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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 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Nonequilibrium radiative heating during outer planet atmospheric entry

The contradictory results obtained by investigators assessing the influence of finite-rate ionization on the radiative heating of probes entering the atmospheres of Jupiter and Saturn are discussed. Tiwari and Szema (1979) found that the radiation heating increased, whereas Leibowitz (1973) and Liebowitz and Kuo (1976) found that it decreased in relation to the results obtained when the ionization rate was assumed to be in equilibrium at the local thermodynamic conditions. The study presented here is limited to stagnation shock layers for nonviscous, hydrogen-helium plasmas with cold, nonblowing conditions at the probe heat shield. It is found that the radiative heating comes mainly from the Balmer region of the spectrum, where the shock layer is optically thin.

Nelson, H. F.↗
NASA NTRS1991

A parametric study of manned aerocapture at Mars

The results of a parametric study of corridor width and stagnation point heating are presented for a range of a probable mission designs and vehicle configurations. Entry velocities were varied from 6 to 10 km/s, L/Ds from 0.1 to 1.0, and ballistic coefficients from 100 to 500 kg/sq m. It is found that vehicles with an L/D of 0.4 to 0.5 provide an entry corridor width of at least 1 deg for velocities up to 10 km/s. For entry velocities below approximately 7 km/s, radiative cooling may be possible for the thermal protection system. It is recommended that, at higher entry speeds, ablative heat shields be used. Maximum integrated stagnation point heat loads were equivalent to or less than those experienced by the Space Shuttle on a typical reentry.

Tauber, M. E.↗
NASA NTRS2008

Experimental Investigation of Project Orion Crew Exploration Vehicle Aeroheating in AEDC Tunnel 9

An investigation of the aeroheating environment of the Project Orion Crew Entry Vehicle has been performed in the Arnold Engineering Development Center Tunnel 9. The goals of this test were to measure turbulent heating augmentation levels on the heat shield and to obtain high-fidelity heating data for assessment of computational fluid dynamics methods. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the data for the purpose of helping to define uncertainty margins for the computational method. Data from both the wind tunnel test and the computational study are presented herein.

Hollis, Brian R.↗
NASA NTRS2017

Fiber Optic Temperature Sensors in TPS: Arc Jet Model Design & Testing

Techniques for using fiber optics with Fiber Bragg Gratings (FBGs) have been developed by IFOS Corp. for use in thermal protection systems (TPS) on spacecraft heat shield materials through NASA Phase 1 and 2 SBIR efforts and have been further improved in a recent collaboration between IFOS and NASA that will be described here. Fiber optic temperature sensors offer several potential advantages over traditional thermocouple sensors including a) multiplexing many sensors in a single fiber to increase sensor density in a given array or to provide spatial resolution, b) improved thermal property match between sensor and TPS to reduce heat flow disruption, c) lack of electrical conductivity.

sensors↗
NASA NTRS2019

Modeling Heatshield Erosion due to Dust Particle Impacts for a Martian Entry Vehicle

New surface damage models are presented that predict the heat shield erosion due to dust particle impact when a spacecraft enters the Martian atmosphere. Existing models were based on Apollo-era experimental data and approximate methods for tracking the dust particle trajectories through the shock layer. These legacy methods will be compared against new results based on more sophisticated particle tracking methods and recent experimental data.

Planetary entry↗
NASA NTRS1969

The Ames combined heating facility.

Heating facility for testing ablative heat shield material models in combined convective and radiative heating environment with constant or transient heating conditions

Gowen, F. E.↗
NASA NTRS1971

Evaluation of active cooling systems for a Mach 6 hypersonic transport airframe, part 2

Transpiration and convective cooling concepts are examined for the fuselage and tail surface of a Mach 6 hypersonic transport aircraft. Hydrogen, helium, and water are considered as coolants. Heat shields and radiation barriers are examined to reduce heat flow to the cooled structures. The weight and insulation requirements for the cryogenic fuel tanks are examined so that realistic totals can be estimated for the complete fuselage and tail. Structural temperatures are varied to allow comparison of aluminum alloy, titanium alloy, and superalloy contruction materials. The results of the study are combined with results obtained on the wing structure, obtained in a previous study, to estimate weights for the complete airframe. The concepts are compared among themselves, and with the uncooled concept on the basis of structural weight, cooling system weight, and coolant weight.

Helenbrook, R. G.↗
NASA NTRS1986

Mars vehicle TCS and aerobrake TPS

General design approach, some problems and technology needs for a Mars vehicle manned module Thermal Control System (TCS) and aerobrake Thermal Protection System (TPS) are discussed. The design approach of the TCS will be similar to that of the space station. Mars atmospheric dust storms are identified as an unknown that will impact the design of the Mars landing vehicle and base facility. New technology may be needed for thermal control surfaces to functionally survive the dust storms. The TPS for the Mars aerocapture vehicle will be subject to marginal stagnation heating rates for conjunction class missions and very high heating rates for opposition class missions. New technology TPS materials or an ablative heat shield will be required for the high stagnation heating rate trajectories. No significantly new technology is needed for the manned modules that do not descent to the Mars surface.

Comer, G.↗
NASA NTRS1975

Material Response of One-Dimensional, Steady-State Transpiration Cooling in Radiative and Convective Environments

A simplified analytical solution for thermal response of a transpiration-cooled porous heat-shield material in an intense radiative-convective heating environment is presented. Essential features of this approach are "two-flux method" for radiative transfer process and "two-temperature" assumption for solid and gas temperatures. Incident radiative-convective heatings are specified as boundary conditions. Sample results are shown using porous silica with CO2 transpiration and some parameters quantitatively show the effect on this transpiration cooling system. Summarized maps for mass injection rate, porosity and blowing correction factor for radiation are obtained in order to realize such a cooling system.

Kubota, Hirotoshi↗
NASA NTRS1987

High Temperature Properties of an Alumina Enhanced Thermal Barrier

The heatshield material requirements for future space vehicles (Aerobraking Orbital Transfer Vehicle & National Aerospace Plane) will depend upon the desired flight capability, configuration and location on the vehicle. These requirements will be more demanding and different from those derived for the materials used in the Shuttle Orbiter thermal protection system. Research was therefore initiated into improving the thermal efficiency of this class of materials by first characterizing their thermal and structural capabilities. Alternate material systems have been developed, tested, and compared with the baseline Shuttle system. This research resulted in the development of several very low density, high porosity (80-90%) materials with enhanced durability and temperature capability. One of the developments was a family of materials referred to as Fibrous Refractory Composite Insulation (FRCI) utilizing a mixture of fibers, each serving a unique purpose. One composition of the FRCI family with two fibers was adopted as a baseline material for use on the third and fourth Orbiters in selected areas due to its strength at a lower density compared to earlier materials. A further improvement in the FRCI family of materials is the Alumina Enhanced Thermal Barrier (AETB), a three-fiber composite. It has a higher temperature capability (composition dependent) than the baseline FRCI as proven by convective heating tests of one composition. AETB was studied to better characterize its performance at high temperature and the mechanisms by which its properties change. In conclusion, the shrinkage of AETB is a factor of six better than baseline FRCI at 1260 C (2300 F) with about a 20% improvement in mechanical properties. This improvement could translate into a 110 C (200 F) higher temperature capability in use as a heat shield material, but further testing in a convective heating environment is required to determine the actual improvement attainable.

Leiser, Daniel B.↗
NASA NTRS1996

Aerothermodynamic Analysis of Commercial Experiment Transporter (COMET) Reentry Capsule

An aerothermodynamic analysis of the Commercial Experiment Transporter (COMET) reentry capsule has been performed using the laminar thin-layer Navier-Stokes solver Langley Aerothermodynamic Upwind Relaxation Algorithm. Flowfield solutions were obtained at Mach numbers 1.5, 2, 5, 10, 15, 20, 25, and 27.5. Axisymmetric and 5, 10, and 20 degree angles of attack were considered across the Mach-number range, with the Mach 25 conditions taken to 90 degrees angle of attack and the Mach 27.5 cases taken to 60 degrees angle of attack. Detailed surface heat-transfer rates were computed at Mach 20 and 25, revealing that heating rates on the heat-shield shoulder ,can exceed the stagnation-point heating by 230 percent. Finite-rate chemistry solutions were performed above Mach 10, otherwise perfect gas computations were made. Drag, lift, and pitching moment coefficients are computed and details of a wake flow are presented. The effect of including the wake in the solution domain was investigated and base pressure corrections to forebody drag coefficients were numerically determined for the lower Mach numbers. Pitching moment comparisons are made with direct simulation Monte Carlo results in the more rarefied flow at the highest Mach numbers, showing agreement within two-percent. Thin-layer Navier-Stokes computations of the axial force are found to be 15 percent higher across the speed range than the empirical/Newtonian based results used during the initial trajectory analyses.

Wood, William A.↗
NASA NTRSDate not supplied

Prediction and Enhancement of Thermal Protection Systems from Meteoroid Damage using a Smooth Particle Hydrodynamic Code

Interplanetary spacecraft are exposed to meteoroid fluxes that range in speeds from 10 to 72 km/sec, far above the capability of today’s test facilities to provide predictions for the likelihood of spacecraft critical penetration. Of special interest are sample return missions, which (though protected by shielding) must often survive years of exposure to the meteoroid environment in order to re-enter Earth’s atmosphere with their scientific cargo. This paper describes the simulation of meteoric material damage to thermal protection systems (TPS) housed beneath protective “garage” (shielding) enclosures using the Smooth Particle Hydrodynamics Code (SPHC) operated by the Institute for Defense Analyses and Stellingwerf Consulting in support of ongoing NASA tasks. The study outlined in this paper considered the impact effect of both meteoric materials such as iron, ice, and chondrites (dunnite), and non-meteoric materials such as aluminum and nylon against both external shielding materials (single and dual aluminum bumpers) and Heat shield for Extreme Entry Environments Technology (HEEET) TPS materials, used alone and in conjunction with shielding. A general predictive damage equation to HEEET TPS is developed from these SPHC simulations for velocities up to 70 km/sec.

Hypervelocity Impact, hydrocode, meteoroid and orb↗
NASA NTRS2017

3D Material Response Analysis of PICA Pyrolysis Experiments

Primarily interested in improving ablation modeling for use in inverse reconstruction of flight environments on ablative heat shields. Ablation model is essentially a component of the heat flux sensor, so model uncertainties lead to measurement uncertainties. Non-equilibrium processes have been known to be significant in low density ablators for a long time, but increased accuracy requirements of the reconstruction process necessitates incorporating this physical effect. Attempting to develop a pyrolysis model for implementation in material response based on the PICA data produced by Bessire and Minton. Pyrolysis gas species molar yields as a function of temperature and heating rate. Several problems encountered while trying to fit Arrhenius models to the data led to further investigation of the experimental setup.

Oliver, Brandon A.↗
NASA NTRS1971

Design and evaluation of active cooling systems for Mach 6 cruise vehicle wings

Active cooling systems, which included transpiration, film, and convective cooling concepts, are examined. Coolants included hydrogen, helium, air, and water. Heat shields, radiation barriers, and thermal insulation are considered to reduce heat flow to the cooling systems. Wing sweep angles are varied from 0 deg to 75 deg and wing leading edge radii of 0.05 inch and 2.0 inches are examined. Structural temperatures are varied to allow comparison of aluminum alloy, titanium alloy, and superalloy structural materials. Cooled wing concepts are compared among themselves, and with the uncooled concept on the basis of structural weight, cooling system weight, and coolant weight.

Mcconarty, W. A.↗
NASA NTRS1987

Smart Houses

GWS takes plans for a new home and subjects them to intensive computerized analysis that does 10,000 calculations relative to expected heat loss and heat gain, then provides specifications designed specifically for each structure as to heating, cooling, ventilation and insulation. As construction progresses, GWS inspects the work of the electrical, plumbing and insulation contractors and installs its own Smart House Radiant Barrier. On completion of the home, GWS technicians use a machine that creates a vacuum in the house and enables computer calculation of the air exchanged, a measure of energy efficiency. Key factor is the radiant barrier, borrowed from the Apollo program. This is an adaptation of a highly effective aluminized heat shield as a radiation barrier holding in or keeping out heat, cold air and water vapor.

Source record↗
NASA NTRS1994

Protective coating for ceramic materials

A protective coating for ceramic materials such as those made of silicon carbide, aluminum oxide, zirconium oxide, aluminoborosilicate and silicon dioxide, and a thermal control structure comprising a ceramic material having coated thereon the protective coating. The protective coating contains, in admixture, silicon dioxide powder, colloidal silicon dioxide, water, and one or more emittance agents selected from silicon tetraboride, silicon hexaboride, silicon carbide, molybdenum disilicide, tungsten disilicide and zirconium diboride. In another aspect, the protective coating is coated on a flexible ceramic fabric which is the outer cover of a composite insulation. In yet another aspect, a metallic foil is bonded to the outer surface of a ceramic fabric outer cover of a composite insulation via the protective coating. A primary application of this invention is as a protective coating for ceramic materials used in a heat shield for space vehicles subjected to very high aero-convective heating environments.

Kourtides, Demetrius A.↗
NASA NTRS2014

Radiative Heating in MSL Entry: Comparison of Flight Heating Discrepancy to Ground Test and Predictive Models

During the recent entry of the Mars Science Laboratory (MSL), the heat shield was equipped with thermocouple stacks to measure in-depth heating of the thermal protection system (TPS). When only convective heating was considered, the derived heat flux from gauges in the stagnation region was found to be underpredicted by as much as 17 W/sq cm, which is significant compared to the peak heating of 32 W/sq cm. In order to quantify the contribution of radiative heating phenomena to the discrepancy, ground tests and predictive simulations that replicated the MSL entry trajectory were performed. An analysis is carried through to assess the quality of the radiation model and the impact to stagnation line heating. The impact is shown to be significant, but does not fully explain the heating discrepancy.

MSL↗