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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 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Multi-shock assembly for protecting a spacecraft surface from hypervelocity impactors

A hypervelocity impact shield assembly for protecting a spacecraft surface from hypervelocity impactors. The shield assembly includes at least one sacrificial impactor disrupting/shocking layer of hypervelocity impactor disrupting/shocking material. A primary spacing element, including space-rated open cell foam material, is positioned between the at least one sacrificial impactor disrupting/shocking layer and a spacecraft surface. A cover member is arranged and disposed relative to the sacrificial impactor disrupting/shocking layer and the primary spacing element to maintain the integrity of the hypervelocity impact shield assembly. In the event of exposure to a hypervelocity impactor, the sacrificial impactor disrupting/shocking layer is perforated while shocking the impactor breaking it into fragments, and/or melting it, and/or vaporizing it, thus providing a dispersion in the form of an expanding debris cloud/plume which spreads the impact energy of the impactor over a volume formed by the primary spacing element between the sacrificial impactor disrupting/shocking layer and the spacecraft surface. This significantly reduces impact lethality at the spacecraft surface. The space-rated open cell foam material provides an extremely lightweight, low-cost, efficient means of spacing and supporting the at least one sacrificial impactor disrupting/shocking layer before, during, and after launch. In a preferred embodiment, the invention is in the form of a multi-shock assembly including a plurality of sacrificial impactor disrupting/shocking layers. In such instance, the hypervelocity impact shield assembly includes a plurality of secondary spacing elements. Each secondary spacing element is positioned adjacent an associated sacrificial impactor disrupting/shocking layer to form a multi-shock subassembly. Thus, a plurality of multi-shock subassemblies are provided which include alternating layers of sacrificial impactor disrupting/shocking layers and secondary spacing elements.

Dvorak, Bruce D.↗
NASA NTRS1997

Future Aerospace Technology in the Service of the Alliance: Sustained Hypersonic Flight - Volume 3

Sustained Hypersonic Flight capability is envisioned for several future military and civilian applications, such as long range immediate reaction reconnaissance, high speed interception of air targets, long range precision strike against hardened or time critical targets, and access to space. The inherent reduction in time-to-target and low vulnerability will permit new operational tactics. The symposium outlined mission needs and operational scenarios for hypersonic vehicles. Addressed were: Technological issues and challenges in external hypersonic aerodynamics and design, propulsion and engine/airframe integration, military ramjet applications, overall system design including structures and materials development needs, and test facilities. Ground based test facilities cannot provide full simulation at Mach numbers above 5. Therefore, the use of Computational Fluid Dynamics design tools is essential. For routine use of the computational tools in analysis, design and optimization, it is necessary to reduce the time for the entire computational process by two orders of magnitude. Hypersonic boundary layer transition remains a critical design issue because of the important impact on engine drag and on heating, which can affect the choice of materials and thermal protection systems. In addition, reliable prediction of scramjet net thrust is an absolute must in resolving hypersonic air breathing vehicle design issues. Due to current facility and computational shortfalls, the development of future hypersonic flight systems requires research flight tests in the technology areas of boundary layer transition and air-breathing propulsion engine performance. For sustained hypersonic flight beyond Mach 6, the supersonic combustion ramjet (scramjet) engine is the only choice for the near future. Only this air-breathing concept offers a significant promise of large reductions in required propellant fractions, increased payload fractions, and reduced size vehicles, together with a foreseeable technological feasibility. Airframe/engine integration, combustor design and thermal management are the predominant engineering tasks. Fuels, hydrogen or hydrocarbon, must be matched to the operational needs of military or civil use. Experience in existing ramjet propelled missiles capable of speeds up to Mach 4 can support the development effort. The potential mission and cost benefits of sustained hypersonic flight to both military and civil applications are tremendous. From the budget point of view, the possibility of sharing development costs between military and civil programs offers a specific advantage.

Source record↗
NASA NTRS1966

Development of Protective Coatings for Chromium-Base Alloys

Chromium alloy sheet was clad with 5 to 10-mil-thick oxidation-resistant nickel-base alloy foils. Specimens also contained 1/2 to 1-mil-thick intermediate layers of platinum, tungsten, and/or W-25Re. Cladding was done by the isostatic hot gas-pressure bonding,.process. The clad chromium-alloy specimens were cyclic oxidation tested at 2100 F and 2300 F for up to 200 hours to determine the effectiveness of these metal claddings in protecting the chromium alloy Cr-5W from oxidation and contamination. Cladding systems consisting of 5-mil-thick Ni-20Cr-20W modified with 3 to 5 weight percent aluminum and containing a 1 /2-mil tungsten diffusion barrier demonstrated potential for long-time service at temperatures as high as 2300 F.

English, J. J.↗
NASA NTRS2009

Mitigation of EMU Glove Cut Hazard by MMOD Impact Craters on Exposed ISS Handrails

Recent cut damages to crewmember extravehicular mobility unit (EMU) gloves during extravehicular activity (EVA) onboard the International Space Station (ISS) has been found to result from contact with sharp edges or pinch points rather than general wear or abrasion. One possible source of cut-hazards are protruding sharp edged crater lips from impact of micrometeoroid and orbital debris (MMOD) particles on external metallic handrails along EVA translation paths. During impact of MMOD particles at hypervelocity an evacuation flow develops behind the shock wave, resulting in the formation of crater lips that can protrude above the target surface. In this study, two methods were evaluated to limit EMU glove cut-hazards due to MMOD impact craters. In the first phase, four flexible overwrap configurations are evaluated: a felt-reusable surface insulation (FRSI), polyurethane polyether foam with beta-cloth cover, double-layer polyurethane polyether foam with beta-cloth cover, and multi-layer beta-cloth with intermediate Dacron netting spacers. These overwraps are suitable for retrofitting ground equipment that has yet to be flown, and are not intended to protect the handrail from impact of MMOD particles, rather to act as a spacer between hazardous impact profiles and crewmember gloves. At the impact conditions considered, all four overwrap configurations evaluated were effective in limiting contact between EMU gloves and impact crater profiles. The multi-layer beta-cloth configuration was the most effective in reducing the height of potentially hazardous profiles in handrail-representative targets. In the second phase of the study, four material alternatives to current aluminum and stainless steel alloys were evaluated: a metal matrix composite, carbon fiber reinforced plastic (CFRP), fiberglass, and a fiber metal laminate. Alternative material handrails are intended to prevent the formation of hazardous damage profiles during MMOD impact and are suitable for flight hardware yet to be constructed. Of the four materials evaluated, only the fiberglass formed a less hazardous damage profile than the baseline metallic target. Although the CFRP laminate did not form any noticeable crater lip, brittle protruding fibers are considered a puncture risk. In parallel with EMU glove redesign efforts, modifications to metallic ISS handrails such as those evaluated in this study provide the means to significantly reduce cut-hazards from MMOD impact craters.

Christiansen, Eric L.↗
NASA NTRS2024

Distributed Sand-Grain Roughness Effects on Blunt Body Hypersonic Transition and Heating

An experimental investigation of distributed sand-grain surface roughness effects on boundary-layer transition and convective heating has been performed. Two representative entry vehicle geometries, a spherical-cap aeroshell and a sphere-cone aeroshell, were considered. Multiple cast ceramic wind tunnel models of each geometry were fabricated with various roughness heights to simulate an ablated thermal protection system. Wind tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Aeroheating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-wall heat transfer distributions from computational flow field simulations.

Brian R. Hollis↗
NASA NTRS1979

Galileo probe thermal protection: Entry heating environments and spallation experiments design

A valid procedure was developed for predicting wall heating and ablation rates about the probe forebody. Entropy layer effects on convective heating rate were analyzed and the computed results are given. A feasibility study to perform an experiment, the selection of a candidate test facility, and the definition of a test matrix are described. The material selection, fabrication, and evaluation of the metal containing carbon-carbon composites for use on the Galileo probe are summarized. The effect of various Jovian atmospheric models on entry heating environment is considered as well as the effect of the nonspherical shape of the planet on entry trajectory.

Balakrishnan, A.↗
NASA NTRS2019

Experimental Investigation of Hexcomb-Pattern Roughness Effects on Transition Onset and Turbulent Heating Augmentation at Mach 6

An experimental investigation of hexcomb-patterned surface roughness effects on boundary-layer transition and turbulent heating has been performed. Two representative entry vehicle geometries, a spherical- cap aeroshell and a sphere-cone aeroshell, were considered. Multiple cast ceramic models of each geometry were fabricated with varying roughness pattern densities and depths that simulated an ablated hexcomb-structure thermal protection system. Wind tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Aeroheating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-w ll heat transfer distributions from computational flow field simulations.

Hollis, Brian R.↗
NASA NTRS2023

Experimental Investigation of Distributed Sand-Grain Roughness Effects on Transition Onset and Turbulent Heating at Mach 6

An experimental investigation of distributed sand-grain surface roughness effects on hypersonic boundary-layer transition and turbulent heating has been performed. Two representative entry vehicle geometries, a sphere-cone aeroshell and a spherical-cap aeroshell, were considered. Cast ceramic models of each geometry were fabricated with distributed sand-grain roughness patterns of different heights that simulated an ablated thermal protection system. Wind tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Aeroheating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-wall heat transfer distributions from computational flow field simulations. The boundary-layer transition data were found to correlate with a functional representation developed in prior roughness studies, although the data scatter was greater owing to the height variability of distributed sand-grain roughness.

Brian R Hollis↗
NASA NTRSDate not supplied

Advanced TUFROC Thermal Protection System

TUFROC is a cost-effective, state of the art, high temperature reusable thermal protection system that is flight-proven as the wing leading edge of X-37B. The low-density dual layer system takes advantage of the high temperature capability of a carbon composite at the surface and transitions to a low-conductivity silica-based Shuttle tile material in depth. An updated surface treatment formulation has resulted in a better performing system, called Advanced TUFROC, which demonstrates lower surface temperature due to reduced catalytic atom recombination at the surface, compared to the original system. An overview of TUFROC and recent arc jet testing will be presented.

TUFROC thermal protection system↗
NASA NTRS2020

Hexcomb Pattern Roughness Effects on Blunt Body Transition and Heating

An experimental investigation of hexcomb pattern surface roughness effects on boundary-layer transition and convective heating has been performed. Two representative entry vehicle geometries, a spherical-cap aeroshell and a sphere-cone aeroshell, were considered. Multiple cast ceramic models of each geometry were fabricated with various roughness pattern densities and depths that simulated an ablated hexcomb-structure thermal protection system. Wind tunnel testing was performed at Mach 6 over a range of Reynolds numbers sufficient to produce laminar, transitional, and turbulent flow. Aeroheating and boundary-layer transition onset data were obtained using global phosphor thermography. The experimental heating data are presented herein, as are comparisons to laminar and turbulent smooth-wall heat transfer distributions from computational flow field simulations. The experimental data were used to develop a unified boundary-layer transition correlation for both sandgrain distributed roughness and hexcomb pattern roughness.

Brian R Hollis↗
NASA NTRS1971

Nozzle and cavity wall cooling limitations on specific impulse of a gas-core nuclear rocket.

Experimental and theoretical study of the performance of a gas-core nuclear rocket, showing that in a uranium plasma nuclear rocket with an 8-ft cavity diameter the cavity wall can be cooled up to a power level of 7400 MW when the rocket operates at a pressure of 1000 atm and the propellant mass flow rate is 10 lbm/sec. A maximum cavity specific impulse of 5800 sec was obtained under such operating conditions. The fact that the wall heat flux was much lower for reactor powers below this level is linked to the presence of a relatively cool, opaque insulating layer of seeded propellant between the hot plasma and the solid wall. The additional coolant required for nozzle protection reduced the maximum cavity specific impulse to 5200 sec.

Kascak, A. F.↗
NASA NTRS1973

Single grid accelerator for an ion thrustor

A single grid accelerator system for an ion thrustor is discussed. A layer of dielectric material is interposed between this metal grid and the chamber containing an ionized propellant for protecting the grid against sputtering erosion.

Margosian, P. M.↗
NASA NTRS1984

Liquid-Oxygen-Compatible Cement for Gaskets

Fluorelastomer and metal bonded reliably by new procedure. To cure fluoroelastomer cement, metal plate/gasket assembly placed in vacuum bag evacuated to minimum vacuum of 27 inches (69 cm) of mercury. Vacuum maintained throughout heating process and until assembly returns to ambient room temperature. Used to seal gaskets and O-rings or used to splice layers of elastomer to form non-standard sized O-rings. Another possible use is to apply protective, liquid-oxygen-compatible coating to metal parts.

Elmore, N. L.↗
NASA NTRS1984

Effect of alloy chemistry and exposure conditions on the oxidation of titanium

Multiwall is a new thermal protection system concept for advanced space transportation vehicles. The system consists of discrete panels made up of multiple layers of foil gage metal. Titanium is the proposed candidate metal for multiwall panels in the reentry temperature range up to 675 C. Oxidation and embrittlement are the principal concerns related to the use of Ti in heat shield applications. The results of a broad study on the oxidation kinetics of several titanium alloys subjected to different exposure conditions are described. The alloys include commercially pure titanium, Ti-6Al-4V, and Ti-6Al-2Sn-4Zr-2Mo. Oxidation studies were performed on these alloys exposed at 704 C in 5-760 torr air pressure and 0 to 50% relative humidity. The resulting weight gains were correlated with oxide thickness and substrate contamination. The contamination depth and weight gains due to solid solutioning were obtained from microhardness depth profiles and hardness versus weight percent oxygen calibration data.

Unnam, J.↗
NASA NTRS1986

Heating distribution comparison between asymmetric and symmetric blunt cones

An experiment was performed to compare the heating distribution between symmetric and asymmetric large-angle blunt cones, with cone angles of 100, 120, and 140 deg. These hot-wall data were obtained from models made from typical thermal protection insulation for proposed aeroassisted orbital transfer vehicles. Experimental data are compared with predictions using a boundary-layer integral matrix procedure with kinetics to determine how well the heating distribution over an asymmetric cone could be approximated using axisymmetric solutions for a cone and spherical segment. In addition, a relationship between the stagnation-point heat-transfer rate and the bow-shock standoff distance for these cones is discussed. The heat-distribution data from the symmetric and asymmetric cones were very similar. Numerical results compared well with the measured wall temperatures at the stagnation point but slightly underpredicted them over the conical portion of the models.

Stewart, D. A.↗
NASA NTRS1991

An XPS study of the stability of Fomblin Z25 on the native oxide of aluminum

Thin films of Fomblin Z25, a perfluoropolyalkylether lubricant, were vapor deposited onto clean, oxidized aluminum and sapphire surfaces, and their behavior at different temperatures was studied using x ray photoelectron spectroscopy (XPS) and thermal desorption spectroscopy (TDS). It was found that the interfacial fluid molecules decompose on the native oxide at room temperature, and continue to decompose at elevated temperatures, as previous studies had shown to occur on clean metal. TDS indicated that different degradation mechanisms were operative for clean and oxidized aluminum. On sapphire substrates, no reaction was observed at room temperature. Our conclusion is that the native oxide of aluminum is neither passive nor protective towards Fomblin Z25. At high temperatures (150 C) degradation of the polymer on sapphire produced a debris layer at the interface with a chemical composition similar to the one formed on aluminum oxide. Rubbing a Fomblin film on a single crystal sapphire also induced the decomposition of the lubricant in contact with the interface and the formulation of a debris layer.

Herrera-Fierro, Pilar↗
NASA NTRS1992

X-ray photoelectron spectroscopy study of the stability of Fomblin Z25 on the native oxide of aluminum

Thin films of Fomblin Z25, a perfluoropolyalkylether lubricant, were vapor deposited onto clean, oxidized aluminum, and onto sapphire surfaces, and their behavior at different temperatures was studied using X-ray photoelectron spectroscopy and temperature desorption spectroscopy (TDS). The interfacial fluid molecules decompose on the native oxide at room temperature, and continue to decompose at elevated temperatures, as previous studies had shown to occur on the clean metal. TDS indicated that different degradation mechanisms were operative for clean and oxidized aluminum. On sapphire substrates, no reaction was observed at room temperature. The native oxide of aluminum is neither passive nor protective towards Fomblin Z25. At higher temperatures (150 C), degradation of the polymer on sapphire produced a debris layer at the interface with a chemical composition similar to the one formed on aluminum oxide. Rubbing a Fomblin film on a single crystal sapphire also induced the decomposition of the lubricant in contact with the interface and the formation of a debris layer.

Herrera-Fierro, Pilar↗
NASA NTRS1995

Screen-Cage Ion Plating Of Silver On Polycrystalline Alumina

Screen-cage ion plating (SCIP) cost-effective technique offering high throwing power for deposition of adherent metal films on ceramic substrates. Applies silver films to complexly shaped substrates of polycrystalline alumina. Silver adheres tenaciously and reduces friction. SCIP holds promise for applying lubricating soft metallic films to high-temperature ceramic components of advanced combustion engines. Other potential uses include coating substrates with metal for protection against corrosion, depositing electrical conductors on dielectric substrates, making optically reflective or electrically or thermally conductive surface layers, and applying decorative metal coats to ceramic trophies or sculptures.

Spalvins, Talivaldis↗