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Results for “HEAT SHIELD”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Ablative heat shield design for space shuttle

Ablator heat shield configuration optimization studies were conducted for the orbiter. Ablator and reusable surface insulation (RSI) trajectories for design studies were shaped to take advantage of the low conductance of ceramic RSI and high temperature capability of ablators. Comparative weights were established for the RSI system and for direct bond and mechanically attached ablator systems. Ablator system costs were determined for fabrication, installation and refurbishment. Cost penalties were assigned for payload weight penalties, if any. The direct bond ablator is lowest in weight and cost. A mechanically attached ablator using a magnesium subpanel is highly competitive for both weight and cost.

Seiferth, R. W.↗
NASA NTRS1998

Leading Edge Heat Shield for Wings of Spacecraft

A heat shield for thermally insulating the leading edge of a wing of a spacecraft during ascent and reentry includes a plurality of rigid tiles. Each tile is formed with a pie-shaped element which interlocks with the complementarily-formed element of another tile. The combination of structure afforded by the pie-shaped elements substantially impedes hypersonic flow of any gases that might enter the gaps between tiles.

Stewart, David A.↗
NASA NTRS1972

Silica reinforcement and char reactions in the Apollo heat shield.

Reactions of the silica reinforcement fiber and ablation char of the Apollo heat shield have been investigated by laboratory tests in an arc image furnace (at temperature levels up to 5000 R, pressures up to 0.7 atm, and heat flux similar to reentry) and by an actual reentry test. Microchemical analyses and X-ray diffraction studies have been made to determine the presence of SiC formation in the char. Experimental data and analytical predictions of thermal and density profiles have been compared for the ablation of virgin heat shield and precharred materials to determine the effects of SiC formation on ablation performance. In all analyses, general agreement was found between chemical composition and the thermal predictions for laboratory tests and reentry materials. In all ablated materials, SiC was formed in the front surface of the char. The highest SiC content found was 58% by weight and found in a high-pressure environment. The SiC formed was found to act as a heat sink in the ablation process and can lower the front surface temperature by 300 R.

Cagliostro, D. E.↗
NASA NTRS2011

Regolith-Derived Heat Shield for Planetary Body Entry and Descent System with In Situ Fabrication

In this paper we will discuss a new mass-efficient and innovative way of protecting high-mass spacecraft during planetary Entry, Descent & Landing (EDL). Heat shields fabricated in situ can provide a thermal-protection system (TPS) for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from regolith materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Three regolith processing and manufacturing methods will be discussed: 1) oxygen & metal extraction ISRU processes produce glassy melts enriched in alumina and titania, processed to obtain variable density, high melting point and heat-resistance; 2) compression and sintering of the regolith yield low density materials; 3) in-situ derived high-temperature polymers are created to bind regolith particles together, with a lower energy budget.

Hogue, Michael D.↗
NASA NTRS2012

Regolith-Derived Heat Shield for Planetary Body Entry and Descent System with In Situ Fabrication

In this paper we will discuss a new mass-efficient and innovative way of protecting high-mass spacecraft during planetary Entry, Descent & Landing (EDL). Heat shields fabricated in situ can provide a thermal-protection system (TPS) for spacecraft that routinely enter a planetary atmosphere. By fabricating the heat shield with space resources from regolith materials available on moons and asteroids, it is possible to avoid launching the heat-shield mass from Earth. Two regolith processing and manufacturing methods will be discussed: 1) Compression and sintering of the regolith to yield low density materials; 2) Formulations of a High-temperature silicone RTV (Room Temperature Vulcanizing) compound are used to bind regolith particles together. The overall positive results of torch flame impingement tests and plasma arc jet testing on the resulting samples will also be discussed.

Hogue, Michael D.↗
NASA NTRS1978

Development, fabrication and test of a high purity silica heat shield

A highly reflective hyperpure ( 25 ppm ion impurities) slip cast fused silica heat shield material developed for planetary entry probes was successfully scaled up. Process development activities for slip casting large parts included green strength improvements, casting slip preparation, aggregate casting, strength, reflectance, and subscale fabrication. Successful fabrication of a one-half scale Saturn probe (shape and size) heat shield was accomplished while maintaining the silica high purity and reflectance through the scale-up process. However, stress analysis of this original aggregate slip cast material indicated a small margin of safety (MS. = +4%) using a factor of safety of 1.25. An alternate hyperpure material formulation to increase the strength and toughness for a greater safety margin was evaluated. The alternate material incorporates short hyperpure silica fibers into the casting slip. The best formulation evaluated has a 50% by weight fiber addition resulting in an 80% increase in flexural strength and a 170% increase in toughness over the original aggregate slip cast materials with comparable reflectance.

Rusert, E. L.↗
NASA NTRS1974

Impact of the retained heat shield concept on science instruments

Associated interface problems between the mass spectrometer and the actual probe design are considered along with the problem of producing a clean sample to the gas detection instrument. Of particular interest is the penetration of the heat shield by the mass spectrometer sampling tube, because it must be demonstrated that the sampling tube can penetrate the heat shield and that the mass spectrometer can be supplied with a contaminant-free gas sample, free of contaminants from out-gassing of the heat shield.

Kessler, W. C.↗
NASA NTRS1990

Prototype V-Groove Radiator Heat Shield

Report describes design, fabrication, and testing of heat radiator equipped with multi-V-groove radiator heat shield. Device compact, efficient structure which removes heat from infrared detectors, gamma-ray detectors, and similar instruments aboard Mars Observer spacecraft and radiates heat into outer space. Designed to maintain detector for gamma-ray spectrometer at temperature of 80 K in cold vacuum under heat load of 80 mW. Prototype made of aluminum, though production shields made of aluminized sheets of polyethylene terephthalate.

Petrick, S. Walter↗
NASA NTRS2013

Enabling a Better Aft Heat Shield Solution for Future Mars Science Laboratory Class Vehicles

System studies are described that compare masses and estimated manufacturing costs of options for the as‐flown Mars Science Laboratory (MSL) aft body Thermal Light Weight Ablator (SLA) 561‐V and its thickness was not optimized using the standard TPS Sizer Tool widely used for heat shield design. Use of the TPS sizing tool suggests that optimization of the SLA thickness could reduce the aft heat shield mass by 40 percent. Analysis of the predicted aft‐shell aerothermodynamics suggests that the bulk of MSL class entry vehicle heat shields could incorporate Advanced Flexible Reusable Surface Insulation (AFRSI). AFRSI has a wellestablished record of relatively inexpensive manufacturing and flight certification based on its use on the lee side of the Space Shuttle. Runs with the TPS Sizer show that the AFRSI solution would be 60 percent lighter than the as‐flown SLA. The issue of Reaction Control System (RCS) heating on the aft shell could be addressed by locally impregnating the AFRSI with silicone to enhance its robustness to short bursts ofheating. Stagnation point arcjet testing has shown that silicone impregnated AFRSI performs well at heat rates of 115 W/cm2 and 0.1 atmospheres for a duration of 40 seconds, far beyond conditions that are expected for MSL class vehicles. The paper concludes with a discussion of manufacturing processes for AFRSI, impregnation approaches and relative cost comparisons to the SLA solution.

McGuire, Mary K.↗
NASA NTRS1985

Thermal fatigue tests of a radiative heat shield panel for a hypersonic transport

A pair of corrugation stiffened, beaded skin Rene 41 heat shield panels were exposed to 20,000 thermal cycles between room temperature and 1450 F to evaluate the thermal fatigue response of Rene 41 metallic heat shields for hypersonic cruise aircraft applications. At the conclusion of the tests, the panels retained substantial structural integrity; however, there were cracks and excessive wear in the vicinity of fastener holes and there was an 80-percent loss in ductility of the skin. Shrinkage of the panel which caused the cracks and wear must be considered in design of panels for Thermal Protection Systems (TPS) applications.

Webb, Granville L.↗
NASA NTRS2004

Heat Shielding Characteristics and Thermostructural Performance of a Superalloy Honeycomb Sandwich Thermal Protection System (TPS)

Heat-transfer, thermal bending, and mechanical buckling analyses have been performed on a superalloy "honeycomb" thermal protection system (TPS) for future hypersonic flight vehicles. The studies focus on the effect of honeycomb cell geometry on the TPS heat-shielding performance, honeycomb cell wall buckling characteristics, and the effect of boundary conditions on the TPS thermal bending behavior. The results of the study show that the heat-shielding performance of a TPS panel is very sensitive to change in honeycomb core depth, but insensitive to change in honeycomb cell cross-sectional shape. The thermal deformations and thermal stresses in the TPS panel are found to be very sensitive to the edge support conditions. Slight corrugation of the honeycomb cell walls can greatly increase their buckling strength.

Ko, William L.↗