SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Degree of pressure exerted by thermoregulatory mechanisms of body in order to forestall serious disorders caused by overheating or supercooling
Thermostability values of heat resistant poly/p-xylylidene-p-phenylene diamine/ from thermogravimetric analysis
Radiative thermal energy incident effects on pressure-suited man on lunar surface
Thermophysical properties of structural materials, ablators, insulations, seals, and adhesives for manned spacecraft
Thermal radiation effects of lunar environment on pressure-suited man on lunar surface
Explore the source record for details and available documents.
Three thermal protection systems proposed for a hypersonic research airplane were subjected to high heating rates in the Langley 8 foot, high temperature structures tunnel. Metallic heat sink (Lockalloy), reusable surface insulation, and insulator-ablator materials were each tested under similar conditions. The specimens were tested for a 10 second exposure on the windward side of an elevon deflected 30 deg. The metallic heat sink panel exhibited no damage; whereas the reusable surface insulation tiles were debonded from the panel and the insulator-ablator panel eroded through its thickness, thus exposing the aluminum structure to the Mach 7 environment.
The hydrocarbon fuels RP-1, commercial-grade propane, JP-7 and chemically pure propane were subjected to tests in a high pressure fuel coking apparatus in order to evaluate their thermal decomposition limits and carbon deposition rates in heated copper tubes. A fuel thermal stability parametric evaluation was conducted at 136-340 atmospheres, bulk fuel velocities of 6-30 m/sec, and tube wall temperatures of 422-811 K, and the effect of inside wall material on deposit formation was evaluated in tests using nickel-plated tubes. Results show RP-1 deposit formation at wall temperatures between 600 and 800 K, with peak deposit formation near 700 K. Substitution of deoxygenated JP-7 for RP-1 showed no improvement, and the carbon deposition rates for propane fuels were found to be higher than those of either of the kerosene fuels. Nickel plating of the tube walls significantly reduced RP-1 carbon deposition rates.
Attention is given to the design features and comparative performance of four two-phase heat transport systems under consideration by NASA as long term space mission devices for the collection, transportation and rejection of waste heat from spacecraft components over the zero-100 C temperature range. The system types are: a mechanically pumped two-phase loop, a capillary pumped loop, an osmotic pumped loop, and a loop employing a biomorph pump. The principal advantage of a capillary pumped loop is that it is a passive system with few moving parts that inherently possesses high reliability. The working fluid recommended for use in such a thermal utility is NH3, which has the requisite combination of high latent heat of vaporization and high surface tension.
An analysis of various Space Station (MSS) active thermal control systems options under consideration because of their potential low weight, efficiency and reliability is reported. The study addressed ordinary and diode-action body mounted radiators, thermal storage, the area and pumping power requirements for single-phase cooling of cold plates, and single-phase and two-phase active cooling loops. The base configuration considered was a core MSS formed by four habitable modules on which are mounted heat pipe radiators articulated to be always edge-on to the sun. A simulation was performed which accounted for the available heat sinks, several thermal loads and the heat rejection capability. No benefits were found with diode-action radiators if the solar absorption is 0.1 or less, although diode-action heat pipes will maintain a higher level of performance in the presence of coating degradation. Thermal storage becomes important only with radiator coating degradation. Water can be up to three times as efficient as Freon as a heat transfer medium. Finally, single-phase cooling offers a lower system mass than two-phase cooling if varying temperature heat loads can be accommodated.
Graphite-reinforced resin matrix composites are being considered for spacecraft structural applications because of their light weight, high stiffness, and lower thermal expansion. Thin protective coatings with stable optical properties and the proper ratio of solar absorption (alpha sub s) to thermal emittance (epsilon) minimize orbital thermal extremes and protect these materials against space environment degradation. Sputtered coatings applied directly to graphite/epoxy composite surfaces and anodized coatings applied to thin aluminum foil were studied for use both as an atomic oxygen barrier and as thermal control coatings. Additional effort was made to develop nickel-based coatings which could be applied directly to composites. These coating systems were selected because their inherent tenacity made them potentially more reliable than commercial white paints for long-life space missions. Results indicate that anodized aluminum foil coatings are suitable for tubular and flat composite structures on large platforms in low Earth orbit. Anodized foil provides protection against some elements of the natural space environment (atomic oxygen, ultraviolet, and particulate radiation) and offers a broad range of tailored alpha sub s/epsilon. The foil is readily available and can be produced in large quantities, while the anodizing process is a routine commercial technique.
The Thermal Control Surface Experiment (TCSE) was flown on the NASA Long Duration Exposure Facility (LDEF) with the basic objective of determining the effects of the near-earth orbital environment and the LDEF induced environment on spacecraft thermal control surfaces. The TCSE was a comprehensive experiment that combined in-space measurements with extensive post-flight analyses of thermal control surfaces to determine the effects of exposure to the low earth orbit space environment. The TCSE is the first space experiment to measure the optical properties of thermal control surfaces the way they are routinely measured in the laboratory. The specific mission objectives of the TCSE were to: (1) determine the effects of the natural and induced space environment on thermal control surfaces; (2) provide in-space performance data on thermal control surfaces; (3) provide in-space comparison to ground-based environmental testing of materials; and (4) develop and prove instrumentation to perform in-space optical testing of materials. Various aspects of the experiment are discussed.
The Air Force Space and Missile Systems Center and Wright Laboratory Materials Directorate (WL/ML) have sponsored and effort to effort to reformulate and qualify Illinois Institute of Technology Research Institute (IITRI) spacecraft thermal control coatings. S13G/LO-1, Z93, and YB-71 coatings were reformulated because the potassium silicate binder, Sylvania PS-7, used in the coatings is no longer manufactured. Coatings utilizing the binder's replacement candidate, Kasil 2130, manufactured by The Philadelphia Quartz (PQ) Corporation, Baltimore, Maryland, and undergoing testing at the Materials Directorate's Space Combined Effects Primary Test and Research Equipment (SCEPTRE) Facility operated by the University of Dayton Research Institute (UDRI). The simulated space environment consists of combined ultraviolet (UV) and electron exposure with in site specimen reflectance measurements. A brief description of the effort at IITRI, results and discussion from testing the reformulated YB-71 coating in SCEPTRE, and plans for further testing of reformulated Z93 and S13G/LO-1 are presented.
In this paper we present an experimental and theoretical study of the effect of microcrack damage on ceramic properties. For the experimental investigation, ceramic samples of aluminum oxide and reaction bonded silicon nitride (RBSN) are used. Thermal shock treatment from different temperatures up to 1000 C is applied to produce the microcracks. Both surface and bulk ultrasonic wave methods are used to correlate the change of elastic constants to microstructural degradation and to determine the change in elastic anisotropy induced by microcrack damage. For the theoretical investigation, damage mechanics, which relates microstructural damage to material service life and mechanical failure, is used. The change in elastic properties due to microcrack damage calculated from the theoretical model is compared with the experimental results for determination of the applicability of damage theory. It is shown that two independent experimental methods (bulk wave and surface wave) give the same results for shear moduli of damaged ceramics. The experimental results aagree reasonably well with the moduli predicted from the cracked solid model.
A method for integrating Aeroheating analysis into conceptual reusable launch vehicle (RLV) design is presented in this thesis. This process allows for faster turn-around time to converge a RLV design through the advent of designing an optimized thermal protection system (TPS). It consists of the coupling and automation of four computer software packages: MINIVER, TPSX, TCAT, and ADS. MINIVER is an Aeroheating code that produces centerline radiation equilibrium temperatures, convective heating rates, and heat loads over simplified vehicle geometries. These include flat plates and swept cylinders that model wings and leading edges, respectively. TPSX is a NASA Ames material properties database that is available on the World Wide Web. The newly developed Thermal Calculation Analysis Tool (TCAT) uses finite difference methods to carry out a transient in-depth 1-D conduction analysis over the center mold line of the vehicle. This is used along with the Automated Design Synthesis (ADS) code to correctly size the vehicle's thermal protection system (TPS). The numerical optimizer ADS uses algorithms that solve constrained and unconstrained design problems. The resulting outputs for this process are TPS material types, unit thicknesses, and acreage percentages. TCAT was developed for several purposes. First, it provides a means to calculate the transient in-depth conduction seen by the surface of the TPS material that protects a vehicle during ascent and reentry. Along with the in-depth conduction, radiation from the surface of the material is calculated along with the temperatures at the backface and interior parts of the TPS material. Secondly, TCAT contributes added speed and automation to the overall design process. Another motivation in the development of TCAT is optimization. In some vehicles, the TPS accounts for a high percentage of the overall vehicle dry weight. Optimizing the weight of the TPS will thereby lower the percentage of the dry weight accounted for by the TPS. Also, this will lower the cost of the TPS and the overall cost of the vehicle.
A method for integrating Aeroheating analysis into conceptual reusable launch vehicle RLV design is presented in this thesis. This process allows for faster turn-around time to converge a RLV design through the advent of designing an optimized thermal protection system (TPS). It consists of the coupling and automation of four computer software packages: MINIVER, TPSX, TCAT and ADS. MINIVER is an Aeroheating code that produces centerline radiation equilibrium temperatures, convective heating rates, and heat loads over simplified vehicle geometries. These include flat plates and swept cylinders that model wings and leading edges, respectively. TPSX is a NASA Ames material properties database that is available on the World Wide Web. The newly developed Thermal Calculation Analysis Tool (TCAT) uses finite difference methods to carry out a transient in-depth I-D conduction analysis over the center mold line of the vehicle. This is used along with the Automated Design Synthesis (ADS) code to correctly size the vehicle's thermal protection system JPS). The numerical optimizer ADS uses algorithms that solve constrained and unconstrained design problems. The resulting outputs for this process are TPS material types, unit thicknesses, and acreage percentages. TCAT was developed for several purposes. First, it provides a means to calculate the transient in-depth conduction seen by the surface of the TPS material that protects a vehicle during ascent and reentry. Along with the in-depth conduction, radiation from the surface of the material is calculated along with the temperatures at the backface and interior parts of the TPS material. Secondly, TCAT contributes added speed and automation to the overall design process. Another motivation in the development of TCAT is optimization.
No abstract available
The thermal energy storage (TES) system has the capability to efficiently preserve thermal energy directly derived from the energy source, minimizing any conversion losses. Especially latent heat storage offers distinct advantages, including a substantial increase in energy storage density and minimization of temperature fluctuations within the plants. However, the phase change material (PCM) employed in latent heat storage has low thermal conductivity. Consequently, various studies are being conducted to enhance heat transfer. One approach to enhance heat transfer involves utilizing metal foam to maximize the heat transfer area. However, modeling metal foam with its intricate structure is a challenging task in numerical analysis. For this reason, ongoing research focuses on simplifying the modeling of metal foam. Nevertheless, fully encompassing all the characteristics of actual metal foam proves to be a challenging task for the simplified analytical model. The objective of this paper is to interpret the simple lattice metal foam analysis model from the perspective of behavior induced by buoyancy. When comparing the analysis results of solid PCM and liquid PCM with the same thermal conductivity under changes in porosity and gravity direction, we conducted an analysis to discern the trends in effective thermal conductivity that are overestimated due to convection. In the analysis, a constant heat flux of 10 kW and a constant surface boundary condition of 350 K were applied, and a sensitivity study regarding the mesh was conducted. The results indicate that, from the perspective of gravity in the simple lattice model, the solid analysis yields an effective thermal conductivity 29-47% higher compared to the liquid analysis. Additionally, as porosity increases, there is an observed increase of 24-33% in effective thermal conductivity.