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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Radiative properties of advanced spacecraft heat shield materials

Experimental results are presented to show the effects of simulated reentry exposure by convective heating and by radiant heating on spectral and total emittance of statically oxidized Inconel 617 and Haynes HS188 superalloys to 1260 K and a silicide coatea (R512E) columbium 752 alloy to 1590 K. Convective heating exposures were conducted in a supersonic arc plasma wind tunnel using a wedge-shaped specimen configuration. Radiant tests were conducted at a pressure of .003 atmospheres of dry air at a flow velocity of several meters per second. Convective heating specimens were subjected to 8, 20, and 38 15-min heating cycles, and radiant heating specimens were tested for 10, 20, 50, and 100 30-min heating cycles. Changes in radiative properties are explained in terms of changes in composition resulting from simulated reentry tests. The methods used to evaluate morphological, compositional and crystallographic changes include: Auger electron spectroscopy; scanning electron microscopy; X-ray diffraction analysis; and electron microprobe analysis.

Cunnington, G. R.↗
NASA NTRS1982

Results of a flight experiment on the catalytic efficiency of the Space Shuttle heat shield

An experiment was performed on the second Space Shuttle flight to verify data from arc-heated wind-tunnel tests indicating that the thermal protection tiles used on the Shuttle's surface are noncatalytic to the recombination of dissociated air. The experiment indicates that the tiles are indeed noncatalytic, and explains the lower than expected temperatures observed in the first two Shuttle flights. The data are reasonably well predicted with computations, using a chemically reactive boundary-layer code. The experimental setup and the computational analysis are described.

Rakich, J. V.↗
DOE OSTI2023

Hypervelocity impact in stellar media: heat shielding, shock fronts and ablation clouds

The Frontier Plasma Science experiments proposed by our team will model the processes occurring during spacecraft atmospheric entries and meteorite planetary collisions by inserting material targets into the Scrape off Layer (SOL) and edge plasma of a tokamak. This project will investigate important plasma-material processes during high-enthalpy atmospheric entries, including heat flux generation and material ablation. Of particular interest is to explore heat and particle flux detachment in front of the sample using concepts and techniques developed in the latest studies of divertor detachment. Exploration missions to the Solar System’s gaseous giants and hyperbolic re-entries into the Earth’s atmosphere require spacecraft that can withstand high velocity (>10 km/s), high heat fluxes (>10 MW/m2), and corresponding enthalpies. Ablative materials have been used as thermal shields to protect the spacecraft from severe heating during entry. However, developing high enthalpy ablating materials is challenging due to the lack of adequate ground testing facilities. Our team stated that the performance of candidate shielding materials could be assessed in a laboratory environment with multiple in-situ diagnostics, which was enabled by the progress in the tokamak research. Modern machines feature relatively long discharges (~10 s) with well-controlled stable plasma conditions at the edge where the heat flux and the flow speed are similar to those experienced during atmospheric entries. The experiments were carried out at the DIII-D National Fusion Facility in San Diego, CA, operated by General Atomics (GA) for the DoE.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1981

Experimental determination of ablation vapor species from carbon phenolic heat-shield materials

The relative concentrations of vapors produced from carbon phenolic composites under thermal loadings approximating those expected at peak heating during vehicle entry into the atmospheres of the outer planets have been determined. The technique of vaporizing the surface of bulk samples by laser irradiation while measuring in situ the vapor species by mass spectrometry is described. Results show that vapor composition varies with irradiance level and with depth of heating (or extent of pyrolysis). Attempts are made to compare these experimental results with the theoretical predictions from computer codes.

Lincoln, K. A.↗
NASA NTRS1983

Stability of silica heat shield materials in the Space Shuttle environment

Sample tiles of Space Shuttle reusable surface insulation (RSI) and fibrous refractory composite insulation (FRCI) have been alternately exposed to simulated mission heating and the Kennedy Space Center launch-pad environment for up to 50 cycles. The two types of tiles were coated with an identical glass composition. The coating on RSI tiles was prone to cracking in the test environment while the same coating on FRCI was not. Absorption of water contaminated with environmental deposits resulted in accelerated devitrification in the RSI tiles. Devitrification in FRCI tiles was significantly less. The rate of environmental degradation of tiles appeared to be dependent on the amount of water absorption.

Ransone, P. O.↗
NASA NTRS2012

Heat Shield Employing Cured Thermal Protection Material Blocks Bonded in a Large-Cell Honeycomb Matrix

A document describes a new way to integrate thermal protection materials on external surfaces of vehicles that experience the severe heating environments of atmospheric entry from space. Cured blocks of thermal protection materials are bonded into a compatible, large-cell honeycomb matrix that can be applied on the external surfaces of the vehicles. The honeycomb matrix cell size, and corresponding thermal protection material block size, is envisioned to be between 1 and 4 in. (.2.5 and 10 cm) on a side, with a depth required to protect the vehicle. The cell wall thickness is thin, between 0.01 and 0.10 in. (.0.025 and 0.25 cm). A key feature is that the honeycomb matrix is attached to the vehicle fs unprotected external surface prior to insertion of the thermal protection material blocks. The attachment integrity of the honeycomb can then be confirmed over the full range of temperature and loads that the vehicle will experience. Another key feature of the innovation is the use of uniform-sized thermal protection material blocks. This feature allows for the mass production of these blocks at a size that is convenient for quality control inspection. The honeycomb that receives the blocks must have cells with a compatible set of internal dimensions. The innovation involves the use of a faceted subsurface under the honeycomb. This provides a predictable surface with perpendicular cell walls for the majority of the blocks. Some cells will have positive tapers to accommodate mitered joints between honeycomb panels on each facet of the subsurface. These tapered cells have dimensions that may fall within the boundaries of the uniform-sized blocks.

Zell, Peter↗
NASA NTRS2014

A New Ablative Heat Shield Sensor Suite Project

A new sensor suite is developed to measure performance of ablative thermal protection systems used in planetary entry vehicles for robotic and human exploration. The new sensor suite measures ablation of the thermal protection system under extreme heating encountered during planetary entry. The sensor technology is compatible with a variety of thermal protection materials, and is applicable over a wide range of entry conditions.

Thermal Protection↗
NASA NTRS1978

Aerothermodynamic environment for Jovian entry with silica heat shield

Solutions are presented for the stagnation-region shock-layer equations, including radiative transfer with spectral lines and silica ablation during Jovian entry. Results for variations of entry angle, sphere-cone configuration, and atmospheric model are given. The effect of silica ablation on the radiative and convective surface heating is correlated with the ratio of the wall to free-stream mass flux. Correlations are also given for spectral distributions. The effect of newly obtained SiO radiation properties on the surface heating is examined.

Green, M. J.↗
NASA NTRS2003

Finite-Rate Ablation Boundary Conditions for Carbon-Phenolic Heat-Shield

A formulation of finite-rate ablation surface boundary conditions, including oxidation, nitridation, and sublimation of carbonaceous material with pyrolysis gas injection, has been developed based on surface species mass conservation. These surface boundary conditions are discretized and integrated with a Navier-Stokes solver. This numerical procedure can predict aerothermal heating, chemical species concentration, and carbonaceous material ablation rate over the heatshield surface of re-entry space vehicles. In this study, the gas-gas and gas-surface interactions are established for air flow over a carbon-phenolic heatshield. Two finite-rate gas-surface interaction models are considered in the present study. The first model is based on the work of Park, and the second model includes the kinetics suggested by Zhluktov and Abe. Nineteen gas phase chemical reactions and four gas-surface interactions are considered in the present model. There is a total of fourteen gas phase chemical species, including five species for air and nine species for ablation products. Three test cases are studied in this paper. The first case is a graphite test model in the arc-jet stream; the second is a light weight Phenolic Impregnated Carbon Ablator at the Stardust re-entry peak heating conditions, and the third is a fully dense carbon-phenolic heatshield at the peak heating point of a proposed Mars Sample Return Earth Entry Vehicle. Predictions based on both finite-rate gas- surface interaction models are compared with those obtained using B' tables, which were created based on the chemical equilibrium assumption. Stagnation point convective heat fluxes predicted using Park's finite-rate model are far below those obtained from chemical equilibrium B' tables and Zhluktov's model. Recession predictions from Zhluktov's model are generally lower than those obtained from Park's model and chemical equilibrium B' tables. The effect of species mass diffusion on predicted ablation rate is also examined.

Chen, Y.-K.↗
NASA NTRS2010

Versatile honeycomb matrix heat shield

A thermal protection system for atmospheric entry of a vehicle, the system including a honeycomb structure with selected cross sectional shapes that receives and holds thermally cured thermal protection (TP) blocks that have corresponding cross sectional shapes. Material composition for TP blocks in different locations can be varied to account for different atmospheric heating characteristics at the different locations. TP block side walls may be attached to all, or to less than all, the corresponding honeycomb structure side walls.

Zell, Peter T.↗
NASA NTRS2001

Vapor Grown Carbon Fiber/Phenolic Matrix Composites for Rocket Nozzles and Heat Shields

The ablation and mechanical and thermal properties of vapor grown carbon fiber (VGCF)/phenolic resin composites were evaluated to determine the potential of using this material in solid rocket motor nozzles. Composite specimens with varying VGCF loading (30%-50% wt) including one sample with ex-rayon carbon fiber plies were prepared and exposed to a plasma torch for 20 s with a heat flux of 16.5 MW/sq m at approximately 1650 C. Low erosion rates and little char formation were observed, confirming that these materials were promising for rocket motor nozzle materials. When fiber loadings increased, mechanical properties and ablative properties improved. The VGCF composites had low thermal conductivities (approximately 0.56 W/m-C) indicating they were good insulating materials. If a 65% fiber loading in VGCF composite can be achieved, then ablative properties are projected to be comparable to or better than the composite material currently used on the Space Shuttle Reusable Solid Rocket Motor (RSRM).

Patton, R. D.↗
NASA NTRS1978

The response of heat-shield materials to intense laser radiation

Experimental results for the response of ATJ graphite, Carbitex 100, and carbon phenolic to intense continuous-wave laser radiation are presented. Both penetration and mass-loss test techniques are used and compared. The results are also compared with a simple ablation theory applicable to laser irradiation. Reasons for the disparity between experiment and theory, and applicability of the results to other heating situations, such as planetary entry, are discussed.

Lundell, J. H.↗
NASA NTRS1971

Heat shields for aircraft - A new concept to save lives in crash fires.

A passenger compartment surrounded by a fire-retardant shell, to protect the occupants long enough for the fire to burn out or for fire-fighting equipment to reach the aircraft and extinguish it, is proposed as a new concept for saving lives in crash fires. This concept is made possible by the recent development of two new fire-retardant materials: a very lightweight foam plastic, called polyisocyanurate foam, and an intumescent paint. Exposed to heat, the intumescent paint expands to many times its original thickness and insulates the surface underneath it. Demonstration tests are illustrated, described and discussed. However, some problems, such as preventing fuselage rupture and protecting windows, must be solved before such a system can be used.

Neel, C. B.↗
NASA NTRS1997

Analysis of Mars Pathfinder Entry Data, Aerothermal Heating, and Heat Shield Material Response

The Mars Pathfinder heatshield contained several thermocouples and resistance thermometers. A description of the experiment, the entry data, and analysis of the entry environment and material response is presented. In particular, the analysis addresses uncertainties of the data and the fluid dynamics and material response models. The calculations use the latest trajectory and atmosphere reconstructions for the Pathfinder entry. A modified version of the GIANTS code is used for CFD (computational fluid dynamics) analyses, and FIAT is used for material response. The material response and flowfield are coupled appropriately. Three different material response models are considered. The analysis of Pathfinder entry data for validation of aerothermal heating and material response models is complicated by model uncertainties and unanticipated data-acquisition and processing problems. We will discuss these issues as well as ramifications of the data and analysis for future Mars missions.

Milos, Frank↗
NASA NTRS1974

Energy transfer in volume-reflecting heat shields

An approximate analysis of radiative transfer in highly scattering materials was developed based on the Kubelka-Munk differential equations--a set of two differential equations representing the spatial rate of change of radiative half-fluxes within the scattering media. These approximate solutions of the Kubelka-Munk equations together with analytic solutions for the steady state temperature distribution for two types of boundary conditions are given. These solutions show the influence of back surface reflectance, scattering power, incident radiative flux parameter and boundary conductive flux parameter on overall reflectance and temperature distributions. This radiation field analysis, adapted to spherical geometry, was applied to the evaluation of the thermal performance of teflon and fritted quartz as heat protection materials for entry into the atmosphere of Jupiter.

Weston, K. C.↗