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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Structural tests on a tile/strain isolation pad thermal protection system

The aluminum skin of the space shuttle is covered by a thermal protection system (TPS) consisting of a low density ceramic tile bonded to a matted-felt material called strain insulation pad (SIP). The structural characteristics of the TPS were studied experimentally under selected extreme load conditions. Three basic types of loads were imposed: tension, eccentrically applied tension, and combined in-plane force and transverse pressure. For some tests, transverse pressure was applied rapidly to simulate a transient shock wave passing over the tile. The failure mode for all specimens involved separation of the tile from the SIP at the silicone rubber bond interface. An eccentrically applied tension load caused the tile to separate from the SIP at loads lower than experienced at failure for pure tension loading. Moderate in-plane as well as shock loading did not cause a measurable reduction in the TPS ultimate failure strength. A strong coupling, however, was exhibited between in-plane and transverse loads and displacements.

Williams, J. G.↗
NASA NTRS1972

Titanium honeycomb structure

A brazed titanium honeycomb sandwich system for supersonic transport wing cover panels provides the most efficient structure spanwise, chordwise, and loadwise. Flutter testing shows that high wing stiffness is most efficient in a sandwich structure. This structure also provides good thermal insulation if liquid fuel is carried in direct contact with the wing structure in integral fuel tanks.

Davis, R. A.↗
NASA NTRS2006

Structural Continuum Modeling of Space Shuttle External Tank Foam Insulation

The Space Shuttle External Tank is covered with rigid polymeric closed-cell foam insulation to prevent ice formation, protect the metallic tank from aerodynamic heating, and control the breakup of the tank during re-entry. The cryogenic state of the tank, as well as the ascent into a vacuum environment, places this foam under significant stress. Because the loss of the foam during ascent poses a critical risk to the shuttle orbiter, there is much interest in understanding the stress state in the foam insulation and how it may contribute to fracture and debris loss. Several foam applications on the external tank have been analyzed using finite element methods. This presentation describes the approach used to model the foam material behavior and compares analytical results to experiments.

Steeve, Brian↗
NASA NTRS1994

Ferric chloride graphite intercalation compounds prepared from graphite fluoride

The reaction between graphite fluoride and ferric chloride was observed in the temperature range of 300 to 400 C. The graphite fluorides used for this reaction have an sp3 electronic structure and are electrical insulators. They can be made by fluorinating either carbon fibers or powder having various degrees of graphitization. Reaction is fast and spontaneous and can occur in the presence of air. The ferric chloride does not have to be predried. The products have an sp2 electronic structure and are electrical conductors. They contain first stage FeCl3 intercalated graphite. Some of the products contain FeCl2*2H2O, others contain FeF3 in concentrations that depend on the intercalation condition. The graphite intercalated compounds (GIC) deintercalated slowly in air at room temperature, but deintercalated quickly and completely at 370 C. Deintercalation is accompanied by the disappearing of iron halides and the formation of rust (hematite) distributed unevenly on the fiber surface. When heated to 400 C in pure N2 (99.99 vol %), this new GIC deintercalates without losing its molecular structure. However, when the compounds are heated to 800 C in quartz tube, they lost most of its halogen atoms and formed iron oxides (other than hematite), distributed evenly in or on the fiber. This iron-oxide-covered fiber may be useful in making carbon-fiber/ceramic-matrix composites with strong bonding at the fiber-ceramic interface.

Hung, Ching-Cheh↗
NASA NTRS1976

Performance of a mullite reusable surface insulation system in a hypersonic stream

The thermal and structural performance of a large panel of mullite reusable surface insulation (RSI) tiles was determined by a series of aerothermal tests in the Langley 8-foot high-temperature structures tunnel. The test panel was designed to represent a portion of the surface structure on a space shuttle orbiter fuselage along a 1,150 K isotherm with the mullite tile system bonded directly to the primary structure. Aerothermal tests were conducted at a free-stream Mach number of 6.7, a total temperature of 1,880 K, a unit Reynolds number of 4.6 million per meter, and dynamic pressure of 62 kPa. The thermal response of the mullite tile was as predicted, and the bond-line temperature did not exceed the design level of 570 K during a typical entry-heat cycle. Geometric irregularities of the tile gaps affected the tile edge temperatures when exposed to hypersonic flow. The tile coating demonstrated good toughness to particle impacts, but the coating cracked and flaked with thermal cycles. The gap filler of woven silica fibers appeared to hinder flow penetration into the gaps and withstood the flow shear of the present tests.

Hunt, L. R.↗
NASA NTRS1995

Ferrix Chloride-Graphite Intercalation Compounds Prepared From Graphite Flouride

The reaction between graphite fluoride and ferric chloride was observed in the temperature range of 300 to 400 C. The graphite fluorides used for this reaction have an sp(sup 3) electronic structure and are electrical insulators. They can be made by fluorinating either carbon fibers or powder having various degrees of graphitization. Reaction is fast and spontaneous and can occur in the presence of air. The ferric chloride does not have to be predried. The products have an sp(sup 2) electronic structure and are electrical conductors. They contain first-stage FeCl3 intercalated graphite. Some of the products contain FeCl2 (center dot) 2H2O, others contain FeF3, in concentrations that depend on the intercalation condition. The graphite intercalated compounds (GIC) deintercalated slowly in air at room temperature, but deintercalated quickly and completely at 370 C. Deintercalation is accompanied by the disappearance of iron halides and the formation of rust (hematite) distributed unevenly on the fiber surface. When heated to 400 C in pure N2 (99.99 vol%), this new GIC deintercalates without losing its molecular structure. However, when the compounds are exposed to 800 C N2, in a quartz tube, they lost most of their halogen atoms and formed iron oxides (other than hematite), distributed evenly in or on the fiber.

Hung, Ching-Cheh↗
NASA NTRS1979

Preliminary design for a space based orbital transfer vehicle

A space-based orbital transfer vehicle has been sized for a 50-metric-ton payload delivery from low-earth-orbit to a geosynchronous orbit. Space basing effected substantial reductions in cryogenic insulation, tank, and body structure. The tank and body structural masses are shown to be lower for space basing because of the larger difference in acceleration loads between the on-orbit case (0.2 g's) and delivery (3.0 g's), the latter applying to ground-based vehicles which are delivered to orbit fully loaded with propellants. Insulation masses are lower because of the absence of an atmosphere and the attendant heat transfer losses. Insulation systems masses are also reduced because of the elimination of the problem of liquefaction and freezing of moisture on the tanks.

Macconochie, I. O.↗
NASA NTRSDate not supplied

Control System Development for A Zero Boil-Off Hydrogen Storage Demonstration With Two-Stage Active Cooling

A NASA team is designing and building a test article to demonstrate the long duration storage of liquid hydrogen via active cooling (cryocoolers) while implementing a two-stage cooling approach. This activity is one of a large portfolio of NASA’s Space Technology Mission Directorate funded activities which focuses on the technology maturation needed for long duration storage of cryogenic liquid in-space. The current state-of-the-art for liquid hydrogen storage on-orbit is on the order of hours while NASA’s planned missions require storage for months, or even years. To enable such a long duration, Zero Boil-Off conditions must be achieved which requires passive technologies to minimize environmental heat loads, but also active cooling to intercept and reject the remaining heat to deep space. The implementation of active cooling results in a significant amount of dry mass added to the vehicle. Utilizing a Two-Stage Cooling approach, analysis and testing to date indicates the Zero Boil-Off of liquid hydrogen can be achieved with less mass and electrical power relative to a single-stage cooling approach where only one cryocooler is used. This activity will demonstrate the fully integrated suite of technologies needed to enable the Two-Stage Cooling approach which includes Multi-Layer Insulation blankets, Low Conductivity Structures, a Tube-On-Tank Heat Exchanger, Tube-On-Shield Heat Exchanger and a Two-Stage Cryogenerator (both 90 Kelvin and 20 Kelvin) with each stage having an independently controlled circulation loop. The bulk of the heat load is intercepted by the 90 Kelvin loop via a thin foil heat exchanger internal to the insulation blankets (also known as a Broad Area Cooling Shield), with the remainder removed at cryofluid temperature (20 Kelvin) by tubes directly welded to the outer tank surface. In this paper, we examine the challenges of controlling the system, and describe the development of control algorithms and software for future testing with liquid hydrogen. The interactions between the cryofans which control circulation loop mass flow rates, electrical heaters which vary the cryocooler lift to simulate the operation of variable-lift flight units, tank heaters which can vary the overall heat load, and the effects of changes in circulation loop pressure as temperatures change, can be complex and must be well understood to maintain steady-state propellant conditions and achieve Zero Boil-Off. Parallel PID loops and watchdog programs implemented on the user interface system will help bring the systems to steady state operation and keep them at selected operating points within acceptable error, while avoiding runaway feedback loops.

cryogenics↗
NASA NTRS1990

Effects of regolith/megaregolith insulation on the cooling histories of differentiated asteroids

The cooling histories of differentiated asteroids are calculated employing a variety of thermal-conductivity structures to simulate the potential insulating effects of regolith and megaregolith layers on the cooling rate. It was found that a combination of relatively thick megaregolith and regolith can potentially reduce the core cooling rate by more than a factor of 10 below the rate predicted by models with conventional thermal conductivity structure. Thus, differences in cratering (regolith production) histories may have resulted in radically different cooling rates for asteroids of similar radius, or in similar cooling rates from asteroids of different radius.

Haack, Henning↗
NASA NTRS1984

Alternate nozzle ablative materials program

Four subscale solid rocket motor tests were conducted successfully to evaluate alternate nozzle liner, insulation, and exit cone structural overwrap components for possible application to the Space Shuttle Solid Rocket Motor (SRM) nozzle asasembly. The 10,000 lb propellant motor tests were simulated, as close as practical, the configuration and operational environment of the full scale SRM. Fifteen PAN based and three pitch based materials had no filler in the phenolic resin, four PAN based materials had carbon microballoons in the resin, and the rest of the materials had carbon powder in the resin. Three nozzle insulation materials were evaluated; an aluminum oxide silicon oxide ceramic fiber mat phenolic material with no resin filler and two E-glass fiber mat phenolic materials with no resin filler. It was concluded by MTI/WD (the fabricator and evaluator of the test nozzles) and NASA-MSFC that it was possible to design an alternate material full scale SRM nozzle assembly, which could provide an estimated 360 lb increased payload capability for Space Shuttle launches over that obtainable with the current qualified SRM design.

Kimmel, N. A.↗
NASA NTRS1977

Experimental performance of an ablative material as an external insulator for a hypersonic research aircraft

An ablative material composed of silica-filled elastomeric silicone was tested to evaluate its thermal and structural performance as an external insulator, or heat shield, for a hypersonic research aircraft. The material was also tested to determine whether it would form a durable char layer when initially heated and thereafter function primarily as an insulator with little further pyrolysis or char removal. Aerothermal tests were representative of nominal Mach 6 cruise conditions of the aircraft, and additional tests were representative of Mach 8 cruise and interference heating conditions. Radiant heating tests were used to simulate the complete nominal Mach 6 surface-temperature history. The silica char that formed during aerothermal tests was not durable. The char experienced a general and preferential surface recession, with the primary mechanism for char removal being erosion. Tests revealed that radiant heating is not a valid technique for simulating aerodynamic heating of the material.

Puster, R. L.↗
NASA NTRS1989

High temperature insulation barrier composite

A composite material suitable for providing insulation for the nozzle structure of the Space Shuttle and other similar surfaces is disclosed. The composite layer is comprised of an outer skin layer of nickel chromium and an interleaved inner region comprising a top layer of nickel chromium foil which acts as a primary convective shield. There are at least two layers of alumina batting adjacent to the layers of silicon carbide fabric. An additional layer of nickel chromium foil is used as a secondary convective shield. The composite is particularly advantageous for use as nozzle insulation because of its ability to withstand high reentry temperatures, its flexibility, oxidation resistance, low conductivity, and light weight.

Onstott, Joseph W.↗
NASA NTRS1983

Structural concepts for future space transportation orbiters

Two structural concepts for a fully reusable STS orbiter are compared with respect to weight and the ability to satisfy structural goals. The cryogenic propellant tanks for both concepts use welded, blade-stiffened aluminum skins for leak-free containment and compatibility with LOX and LH2. The tanks are sized primarily by pressure, although some areas are sized by bending and fracture mechanics considerations. The tank concept is regarded as state-of-the-art, even though it requires a closed cell foam insulation with a dry air purge to prevent air liquefaction and ice formation. It is noted that the foam will require testing to verify its design life for use in the STS. One concept has nonintegral tanks suspended inside an insulated aluminum airframe and thrust structure. Here, a durable thermal protection system external insulation is mechanically attached to the airframe, allowing a buckled skin design. The other concept uses a novel structural arrangement consisting of a separate tank/thrust structure that supports a hot advanced-carbon-carbon aeroshell structure.

Taylor, A. H.↗
NASA NTRS1984

Shell tile thermal protection system

A reusable, externally applied thermal protection system for use on aerospace vehicles subject to high thermal and mechanical stresses utilizes a shell tile structure which effectively separates its primary functions as an insulator and load absorber. The tile consists of structurally strong upper and lower metallic shells manufactured from materials meeting the thermal and structural requirements incident to tile placement on the spacecraft. A lightweight, high temperature package of insulation is utilized in the upper shell while a lightweight, low temperature insulation is utilized in the lower shell. Assembly of the tile which is facilitated by a self-locking mechanism, may occur subsequent to installation of the lower shell on the spacecraft structural skin.

Macconochie, I. O.↗
NASA NTRS1974

Performance of LI-1542 reusable surface insulation system in a hypersonic stream

The thermal and structural performance of a large panel of LI-1542 reusable surface insulation tiles was determined by a series of cyclic heating tests using radiant lamps and aerothemal tests in the Langley 8-foot high-temperature structures tunnel. Aerothermal tests were conducted at a free-stream Mach number of 6.6, a total temperature of 1830 K, Reynolds numbers of 2.0 and 4,900,000 per meter, and dynamic pressures of 29 and 65 kPa. The results suggest that pressure gradients in gaps and flow impingement on the header walls at the end of longitudinal gaps are sources for increased gap heating. Temperatures higher than surface radiation equilibrium temperature were measured deep in gaps and at the header walls. Also, the damage tolerance of the LI-1542 tiles appears to be very high. Tile edge erosion rate was slow; could not be tolerated in a shuttle application. Tiles soaked with water and subjected to rapid depressurization and aerodynamic heating showed no visible evidence of damage.

Hunt, L. R.↗
NASA NTRS1993

Positron lifetime spectroscopy for investigation of thin polymer coatings

In the aerospace industry, applications for polymer coatings are increasing. They are now used for thermal control on aerospace structures and for protective insulating layers on optical and microelectronic components. However, the effectiveness of polymer coatings depends strongly on their microstructure and adhesion to the substrates. Currently, no technique exists to adequately monitor the quality of these coatings. We have adapted positron lifetime spectroscopy to investigate the quality of thin coatings. Results of measurements on thin (25-micron) polyurethane coatings on aluminum and steel substrates have been compared with measurements on thicker (0.2-cm) self-standing polyurethane discs. In all cases, we find positron lifetime groups centered around 560 psec, which corresponds to the presence of 0.9-A(exp 3) free-volume cells. However, the number of these free-volume cells in thin coatings is larger than in thick discs. This suggests that some of these cells may be located in the interfacial regions between the coatings and the substrates. These results and their structural implications are discussed in this report.

Singh, Jag 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.↗