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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Study of low-cost fabrication of ablative heat shields

The major objectives were accomplished in three tasks: (1) modification of the ablative material composition for ease of fabrication as well as thermal and mechanical performance; (2) scaled-up, simplified, manufacturing techniques which resulted in cost reductions; and (3) the identification of a significant design problem caused by the differential pressure buildup imposed on mechanically attached ablative heat shield panels during launch.

Norwood, L. B.↗
NASA NTRS1976

Evaluation of dispersion strengthened nickel-base alloy heat shields for space shuttle application

The results obtained in a program to evaluate dispersion-strengthened nickel-base alloys for use in a metallic radiative thermal protection system operating at surface temperatures to 1477 K for the space shuttle were presented. Vehicle environments having critical effects on the thermal protection system are defined; TD Ni-20Cr characteristics of material used in the current study are compared with previous results; cyclic load, temperature, and pressure effects on sheet material residual strength are investigated; the effects of braze reinforcement in improving the efficiency of spotwelded joints are evaluated; parametric studies of metallic radiative thermal protection systems are reported; and the design, instrumentation, and testing of full scale subsize heat shield panels in two configurations are described. Initial tests of full scale subsize panels included simulated meteoroid impact tests, simulated entry flight aerodynamic heating, programmed differential pressure loads and temperatures simulating mission conditions, and acoustic tests simulating sound levels experienced during boost flight.

Johnson, R., Jr.↗
NASA NTRS1995

Ames Research Center Mars/Pathfinder Heat Shield Design Verification ARC-JET Test

Design verification tests were performed on samples representing the aerobrake of the Mars/Pathfinder vehicle. The test specimens consisted of the SLA-561V ablator bonded to the honeycomb structure. The primary objective was to evaluate the ablation materials performance and to measure temperatures within the ablator, at the structural bondline and at the back sheet of the honeycomb structure. Other objectives were to evaluate the effect of ablative repair plug material treatment and voids in the heat shield. A total of 29 models were provided for testing in the Ames 60MW arc-jet facility. Of these, 23 models were flat-faced and six remaining models were curved edge ones, intended to simulate the conditions on the curved rim of the forebody where the maximum shear occurred. Eight sets of test conditions were used. The stagnation point heating rates varied from 47 to 240 W/cm2 and the stagnation pressures from 0.15 to 0.27 atm. (The maximum flight values are 132 W/cm2 and 0.25 atm) The majority of these runs were made at a nominal stagnation pressure of 0.25 atm. Two higher pressure runs were made to check the current (denser) ablation material for spallation, or other forms of thermal stress failure. Over 60% of the flatfaced models yielded good thermocouple data and all produced useful surface recession information. Of the five curved-edge models that were tested, only one gave good data; the remaining ones experienced model-holder failure. The test results can be summarized by noting that no failure of the ablative material was observed on any model. Also, the bondline temperature design limit of 250 C was never reached within an equivalent flight time despite a stagnation point heat load that exceeded the maximum flight value by up to 130%. At heating rates of over 200W/cm2 and stagnation pressures of 0.25 atm, or greater, the average surface recessions exceeded 0.5 cm on some models. The surface roughness increased dramatically at pressures above 0.25 atm and was four times greater at 0.27 atm than at 0.25 atm. Procured repair plug material performed much better than room-temperature cured plugs, as observed in the previous tests. Voids in the ablator did not increase local temperatures and gaps did not grow during testing.

Tran, Huy K.↗
NASA NTRS2014

Remote Recession Sensing of Ablative Heat Shield Materials

Tests were performed to demonstrate the feasibility of a new method of measuring surface recession of a material sample during arc-jet testing in the NASA Ames mARC subscale developmental facility. The measurement principle was inspired through tracer elements such as Ca and Na which were seen in the spectra taken during the airborne observation campaign of the Stardust re-entry and which could be clearly observed standing out against the emission spectra emitted by postshock layer and glowing surface of the re-entry capsule. The measurement principle involves seeding of the heat shield materials at a defined depth with tracer elements which show strong and characteristic emission lines in the post shock plasma. Once the material recession reaches the seeding depth, these elements get into the hot plasma and show up in the emission spectra. The methodology was successfully demonstrated during arc-jet testing of phenolic impregnated carbonablator (PICA) material which was seeded in depth with a mixture of NaCl and MgCl in powder form. In the emission spectroscopy data, the emission lines of Mg and Na showed up about 1.5 seconds after probe insertion into the arc-jet plasma and vanished after another 2.5 seconds when recessionhad consumed the seeding material. From these data, a recession rate of about 1 mm/s is estimated.The heat flux during the test was measured to be 2575 W/sq cm on a hemispherical heat flux probewhich corresponds to a heat flux of 1036 W/sq cm on the rectangular test articles. An estimate for a lower limit of the surface temperature of 2800K during the test was obtained by fitting Planck radiation to the continuum spectra emitted by the PICA sample. Typical recession rates of PICA during testing in the large arc-jet facilities at similar test conditions are reported to be on the order of 0.05 to 0.1 cm/s which agrees well with the recession rates of 0.05-0.06 cm/s estimated from the emission spectroscopy data. Further tests under better controlled conditions are suggested to quantify this measurement method. Through a different choice of seeding materials with lower melting point, an extension of the measurement principle to monitor char depth seems feasible but was not yet demonstrated. Possible applications besides ground testing are recession and possibly char depth measurements during real re-entry. Detection through emission spectroscopy could be accomplished through ground based or airborne observation as performed during the Stardust and Hayabusa re-entries, or through on-board spectrometers. A suitable mission would be the re-entryof the OSIRIS-REX mission planned for late 2023. The measured data are presented and interpreted, the results and details of future applications are discussed.

instrumentation↗
NASA NTRS1972

Development of fused slurry silicide coatings for tantalum reentry heat shields

A fused slurry silicide coating was developed to provide atmospheric reentry protection for the 90Ta-lOW alloy. Overlaying the silicide with a highly refractory glass greatly improved total lifetime and reliability of the coating system. Low pressure, slow cycle lifetimes in excess of 100 cycles were consistently recorded for 1700 K - 13 and 1300 N/sq m test conditions. A minimum of 25 cycles was obtained for 1810 K - 1300 N/sq m conditions. About 50 simulated reentry cycles (variable temperature, pressure, and stress) were endured by coated 1-inch miniature heat shield panels when exposed to a maximum of 1700 K and either internal or external pressure conditions.

Warnock, R. V.↗
NASA NTRSDate not supplied

Implementing the 3-D woven Mid-Density Carbon Phenolic (3MDCP) Heat Shield for the Mars Sample Return (MSR) Earth Entry Vehicle (EEV)

he Mars Sample Return (MSR)Program will return Martian soil samples to Earth in the early 2030s. Since the biological content of these samples is unknown, and potentially upsetting to Earth’s biosphere, the Earth Entry Vehicle (EEV) is required to be the most reliable entry probe ever devised. Entering Earth’s atmosphere at ~12 km/s, after up to a 6-day in-space free flight phase, the EEV must tolerate the cis-lunar Micro-Meteoroid environment, the near-Earth Orbital debris environment, and then entry heating of over 3,000 W/cm2 and pressures of250 kPa before ballistically impacting Earth’s surface at nearly 45 m/s. To reliably accomplish such a mission, NASA Ames is implementing the 3-D woven Mid-Density Carbon Phenolic (3MDCP) Heat Shield, derived from the prior Heatshield for Extreme Entry Environments(HEEET) material system [1]. 3MDCP development accomplishments to date, along with the future efforts to deliver the flight hardware are presented, along with a discussion of the manufacturing risks and accompanying mitigations achieved.

J C Vander Kam↗
NASA NTRS2018

A Reliability Comparison of Classical and Stochastic Thickness Margin Approaches to Address Material Property Uncertainties for the Orion Heat Shield

The Orion Thermal Protection System (TPS) margin process uses a root-sum-square approach with branches addressing trajectory, aerothermodynamics, and material response uncertainties in ablator thickness design. The material response branch applies a bondline temperature reduction between the Avcoat ablator and EA9394 adhesive by 60 C (108 F) from its peak allowed value of 260 C (500 F). This process is known as the Bond Line Temperature Material Margin (BTMM) and is intended to cover material property and performance uncertainties. The value of 60 C (108 F) is a constant, applied at any spacecraft body location and for any trajectory. By varying only material properties in a random (monte carlo) manner, the perl-based script mcCHAR is used to investigate the confidence interval provided by the BTMM. In particular, this study will look at various locations on the Orion heat shield forebody for a guided and an abort (ballistic) trajectory.

margins↗
NASA NTRS2018

A Reliability Comparison of Classical and Stochastic Thickness Margin Approaches to Address Material Property Uncertainties for the Orion Heat Shield

The Orion Thermal Protection System (TPS) margin process uses a root-sum-square approach with branches addressing trajectory, aerothermodynamics, and material response uncertainties in ablator thickness design. The material response branch applies a bond line temperature reduction between the Avcoat ablator and EA9394 adhesive by 60 C (108 F) from its peak allowed value of 260 C (500 F). This process is known as the Bond Line Temperature Material Margin (BTMM) and is intended to cover material property and performance uncertainties. The value of 60 C (108 F) is a constant, applied at any spacecraft body location and for any trajectory. By varying only material properties in a random (monte carlo) manner, the perl-based script mcCHAR is used to investigate the confidence interval provided by the BTMM. In particular, this study will look at various locations on the Orion heat shield forebody for a guided and an abort (ballistic) trajectory.

FIAT↗
NASA NTRS1975

Reflecting heat shields made of microstructured fused silica

Heat sheidls constructed from selected monodisperse distributions of high-purity fused-silica particles are efficient reflectors of visible and near-UV radiation generated in shock-layer of space probe during atmospheric entry.

Congdon, W. M.↗
NASA NTRS1980

Galileo probe forebody entry thermal protection - Aerothermal environments and heat shielding requirements

Solutions are presented for the aerothermal heating environments and the material thermal response for the forebody heatshield on the candidate 242 kg Galileo probe entering the modeled nominal and cold-dense Jovian atmospheres. In the flowfield analysis, a finite difference procedure was employed to obtain benchmark predictions of pressure, radiation and convective heating rates (both laminar and turbulent) and the corresponding wall blowing obtained under the steady state approximation. The fluxes over the probe flank were found to be in a range where spallation is an important mass loss mechanism. The predicted heating rates were also used as boundary conditions for a charring materials ablation which was used to predict thermochemical based surface recession, mass loss and bondline temperatures. The contingency factor of 30% currently employed by NASA was found to be insufficient for entry into the cold-dense atmosphere.

Nicolet, W. E.↗
NASA NTRS1980

Connector heat shield

Polytetrafluoroethylene tape wrapped around electrical connectors protects them from heat damage during soldering. Tape is easily removed after contacts are joined.

Clarke, S.↗
NASA NTRS1974

Energy transfer in volume-reflecting heat shields

An approximate analytic solution is provided for the unsteady radiative heating of a highly scattering medium under constant heat flux boundary conditions. The accuracy of the method in representing the reflectance, radiative flux, and radiative flux divergence of the materials was also explored, as was the influence of anistropic scattering on these parameters.

Weston, K. C.↗
NASA NTRS1973

Space vehicle engine and heat shield environment review. Volume 1: Engineering analysis

Methods for predicting the base heating characteristics of a multiple rocket engine installation are discussed. The environmental data is applied to the design of adequate protection system for the engine components. The methods for predicting the base region thermal environment are categorized as: (1) scale model testing, (2) extrapolation of previous and related flight test results, and (3) semiempirical analytical techniques.

Mcanelly, W. B.↗
NASA NTRS1974

Castable thermal insulation for use as heat shields

Structural members supporting the afterburners of high thrust rocket engines are subjected to extreme heating, along with severe vibration and high acceleration levels during early lift-off. Chemically-bonded, castable, zircon composite foams were developed and successfully tested to meet specific, laboratory simulated lift-off conditions.

Mountvala, A. J.↗