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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
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 NTRS1978

Aerothermodynamic assessment of corrugated panel thermal protection systems

The results of an analysis of wind-tunnel data for the effect of wave shape on local boundary layer flows and heat transfer are used to assess the feasibility of using corrugated panels on a Single-Stage-To-Orbit (SSTO) vehicle. In the flight assessment, available engineering design procedures are used to compute the inviscid and boundary layer flow for a given flight trajectory. The flight vehicle flow field calculations are then used in conjunction with wind-tunnel corrugated panel heating data correlations to examine the heating penalties associated with using corrugated panels at a number of locations on the vehicle. Severe heating penalties, as a result of the corrugations, were found over large areas of the windward surface. Maximum surface temperature penalties ranged from 100K to 500K and very large structural advantages for the corrugations would be required to override the heating penalties. Computed peak corrugation temperatures were within allowable limits of the equivalent smooth wall material for most leeward and side fuselage locations and corrugated panels will be most likely limited to these locations.

Brandon, H. J.↗
NASA NTRS1982

Aerothermodynamic environment of a Titan aerocapture vehicle

The extent of convective and radiative heating for a Titan aerocapture vehicle is investigated. The flow in the shock layer is assumed to be axisymmetric, steady, viscous, and compressible. It is further assumed that the gas is in chemical and local thermodynamic equilibrium and tangent slab approximation is used for the radiative transport. The effect of the slip boundary conditions on the body surface and at the shock wave are included in the analysis of high-altitude entry conditions. The implicit finite difference techniques is used to solve the viscous shock-layer equations for a 45 degree sphere cone at zero angle of attack. Different compositions for the Titan atmosphere are assumed, and results are obtained for the entry conditions specified by the Jet Propulsion Laboratory.

Tiwari, S. N.↗
NASA NTRS1981

Determination of atmospheric properties for STS-1 aerothermodynamic investigations

A procedure for determining an approximation to the freestream atmospheric properties along the Shuttle entry trajectory is presented. Meteorological data as input is obtained by rawinsondes from surface to 70 km, and meteorological spheres from 60-90 km, launched from Hawaii and California. The Langley Atmospheric Information Retrieval System (LAIRS) developed to approximate the atmospheric freestream properties along the flight path, is outlined, noting temperature and wind data are interpolated in altitude, while gradients and diurnal and semidiurnal coefficients are taken from the COSPAR reference atmosphere. The data are input to a model to project temperature profiles for the Shuttle descent, and the input atmospheric parameters are listed. Efforts are continuing in order to correct discrepancies in the generated profiles for regions below 3 km.

Price, J. M.↗
NASA NTRS1982

Analysis of aerothermodynamic environment of a Titan aerocapture vehicle

The feasibility of an aerocapture vehicle mission has been emphasized recently for inner and outer planetary missions. Aerocapture involves a system concept which utilizes aerodynamic drag to acquire the velocity reduction necessary to obtain a closed planetary orbit from a hyperbolic flyby trajectory. It has been proposed to use the atmosphere of Titan for braking into a Saturn orbit. This approach for a Saturn orbital mission is expected to cut the interplanetary cruise travel time to Saturn from 8 to 3.5 years. In connection with the preparation of such a mission, it will be necessary to provide a complete analysis of the aerodynamic environment of the Titan aerocapture vehicle. The main objective of the present investigation is, therefore, to determine the extent of convective and radiative heating for the aerocapture vehicle under different entry conditions. This can be essentially accomplished by assessing the heating rates in the stagnation and windward regions of an equivalent body.

Tiwari, S. N.↗
NASA NTRS1982

Advancements in aerothermodynamics in support of the Galileo probe

The Galileo probe, scheduled for launch in 1985, will experience the most severe heating environment ever encountered for a planetary mission, and the probe design should be based on a reliable definition of the aerothermal environment during entry. Significant advances in numerical computations make possible the prediction of such complicated flowfields, which include the effects of massive ablation injection, radiative transfer, turbulence, shape change and small angle of attack. Also included is a discussion of modeling and data inputs for numerical computations as deduced from both experimental and analytical studies concerning: radiation, transport and thermodynamic properties; probe afterbody flowfield and heating conditions; and the role of ablator spallation.

Moss, J. N.↗
NASA NTRS1986

Aerothermodynamics of transatmospheric vehicles

A transatmospheric vehicle (TAV) using airbreathing propulsion requires a long acceleration period within the denser part of the atmosphere to reach orbital speed. The long flight time, coupled with the need for a low-drag configuration, results in severe heating of parts of the vehicle. The ascent peak stagnation point and wing leading edge equilibrium wall temperatures are about 3500 K and 2500 K, respectively, likely requiring some form of mass addition cooling. The corresponding temperatures during entry are 1000 K lower. The vehicle windward centerline temperatures are more moderate, with values peaking around 1300 K during both ascent and entry. Therefore, radiative cooling should be effective over large areas of the vehicle. The windward centerline heat loads during entry are comparable to those for low acceleration ascent trajectories. However, ascent heat loads for the stagnation point and the wing leading edge are about three times higher than those during entry. For comparison, the entry heat load for the TAV's stagnation point is about three times higher than the value for Shuttle. Therefore the ascent heat load at the TAV's stagnation point exceeds the Shuttle's entry value by an order of magnitude.

Tauber, M. E.↗
NASA NTRS1987

Computational fluid dynamics and aerothermodynamics

At the present time the efforts on this project are organized into two areas: the development of an approximate stagnation point solution and approximate flowfield studies which can be used to develop and investigate shock jump, electron temperature, radiation, vibration-dissociation coupling, and chemistry models. Progress in each area is discussed.

Carlson, Leland A.↗
NASA NTRS1987

An aerothermodynamic study of slender conical vehicles

A numerical study was performed to assess the applicability of some current techniques which can be used for aerothermal predictions over slender spherically blunted cones. Predictions using a viscous-shock-layer method and several engineering approaches were compared with experimental results from flight and ground-based tests, with each other, and with other detailed results. Good agreement was obtained in comparisons with laminar and turbulent heating data from the Reentry F flight vehicle and with the wind-tunnel data. In particular, the viscous-shock-layer method was shown to yield excellent comparisons and should be useful in providing detailed flowfield and surface values for slender blunted cones. Additional predictions were obtained with these methods for two 5-deg half-angle cones with different nose radii to illustrate the effects of nose bluntness and angle of attack on drag and heat transfer. These results demonstrate the benefit of nose blunting with respect to heating and drag for laminar and transitional flow at zero-lift conditions and the benefit of heating reduction at angle of attack. Detailed comparison of the engineering code predictions with the viscous-shock-layer results for these additional cases generally showed good agreement except for the laminar prodictions at angle of attack on the forward cone surface. Nonequilibrium calculations at 0-deg angle of attack showed that substantial benefits of low surface catalysis existed only in the nose-dominated region.

Thompson, R. A.↗
NASA NTRS1987

Development and application of computational aerothermodynamics flowfield computer codes

Multiple nozzle plume flow field is computed with a 3-D, Navier-Stokes solver. Numerical simulation is performed with a flux-split, two-factor, time asymptotic viscous flow solver of Ying and Steger. The two factor splitting provides a stable 3-D solution procedure under ideal-gas assumptions. An ad-hoc acceleration procedure that shows promise in improving the convergence rate by a factor of three for steady state problems is utilized. Computed solutions to generic problems at various altitude and flight conditions show flow field complexity and three-dimensional effects due to multiple nozzle jet interactions. Viscous, ideal gas solutions for the symmetric nozzle are compared with other numerical solutions.

Venkatapathy, Ethiraj↗
NASA NTRS1987

Experimental aerothermodynamic research of hypersonic aircraft

The 2-D and 3-D advance computer codes being developed for use in the design of such hypersonic aircraft as the National Aero-Space Plane require comparison of the computational results with a broad spectrum of experimental data to fully assess the validity of the codes. This is particularly true for complex flow fields with control surfaces present and for flows with separation, such as leeside flow. Therefore, the objective is to provide a hypersonic experimental data base required for validation of advanced computational fluid dynamics (CFD) computer codes and for development of more thorough understanding of the flow physics necessary for these codes. This is being done by implementing a comprehensive test program for a generic all-body hypersonic aircraft model in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel over a broad range of test conditions to obtain pertinent surface and flowfield data. Results from the flow visualization portion of the investigation are presented.

Cleary, Joseph W.↗
NASA NTRS1988

Computational fluid dynamics and aerothermodynamics

Approximations applicable to the radiating, reacting, and conducting stagnation region of a hypervelocity vehicle were incorporated into a method for rapidly obtaining approximate solutions. This solution utilizes a coordinate system based upon the origin of the radiative losses and includes in a phenomenologically correct manner the effects of chemical and thermal nonequilibrium, and nonequilibrium, nongray radiative transfer. Results were presented which demonstrate the usefulness of the method and indicate which radiation parameters require further study and definition. Excellent comparisons were obtained with published results for the Fire2 data. An axisymmetric nonequilibrium inverse method was modified and extended and used to investigate and compare various vibration dissociation chemistry coupling models and radiative heat transfer approximations. The similarities, differences, and consequences of using these models in the Aero-assist Orbital Transfer Vehicles flight regime will be discussed.

Carlson, Leland A.↗
NASA NTRS1988

An overview of hypersonic aerothermodynamics

This paper briefly reviews some national studies and new programs concerning hypersonic flight. The flight environment that will be encountered by this new class of hypersonic vehicles is described, and the fluid-dynamic and chemical phenomena that occur in hypersonic flight are examined. Ground-based facilities are briefly described, and their use in helping to validate the codes is examined.

Chapman, Gary T.↗