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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

A study of the aerothermal entry environment

Turbulent flow-field solutions are presented for the forebody of a probe entering a nominal Jupiter atmosphere. A two-layer eddy viscosity model is used throughout the study. The solutions that include coupled ablation injection are significant in that they are the first flow-field solutions for a 45-deg spherically blunted cone, the forebody configuration for the Galileo probe. The sphere-cone solutions that include coupled ablation injection show that the radiative heating on much of the conical portion of the probe often exceeds the radiative heating predicted for the corresponding no-injection solution. Turbulence produces a marked decrease in the effectiveness of the ablation products to absorb radiation because of the way in which turbulence alters the structure and composition of the flow. Selected laminar and turbulent solutions are presented for a hyperboloid that demonstrate the effect of turbulence and probe configuration on surface heating and ablative mass loss. Furthermore, the sensitivity of the heating to variations in the turbulence model is presented for the sphere-cone configuration.

Moss, J. N.↗
NASA NTRS1980

Influence of nonequilibrium radiation and shape change on aerothermal environment of Jovian entry body

Radiative transfer equations are derived under nonequilibrium conditions which include multilevel energy transitions. The nonequalibrium results, obtained with and without ablation injection in the shock layer, are found to be greatly influenced by the temperature distribution in the shock layer. In the absence of ablative products, the convective and radiative heating to the entry body are reduced significantly under nonequilibrium conditions. The influence of nonequilibrium is found to be greater at higher entry altitudes. With coupled ablation and carbon phenolic injection, 16 chemical species are used in the ablation layer for radiation absorption. Equilibrium and nonequilibrium results are compared under peak heating conditions. A 45 degree sphere cone, a 35 degree hyperboloid, and a 45 degree ellipsoid were used to study probe shape change. Results indicate that the shock layer flow field and heat transfer to the body are influenced significantly by the probe shape change. The effect of shape change on radiative heating of the afterbodies is found to be considerably larger for the sphere cone and ellipsoid than for the hyperboloid.

Tiwari, S. N.↗
NASA NTRS1981

Influence of nonequilibrium radiation and shape change on aerothermal environment of a Jovian entry body

The influence of nonequilibrium radiative energy transfer and the effect of probe configuration changes on the flow phenomena around a Jovian entry body are investigated. The radiating shock layer flow is assumed to be axisymmetric, viscous, laminar and in chemical equilibrium. The radiative transfer equations are derived under nonequilibrium conditions which include multilevel energy transitions. The equilibrium radiative transfer analysis is performed with an existing nongray radiation model which accounts for molecular band, atomic line, and continuum transitions. The nonequilibrium results are obtained with and without ablation injection in the shock layer. The nonequilibrium results are found to be greatly influenced by the temperature distribution in the shock layer. In the absence of ablative products, the convective and radiative heating to the entry body are reduced under nonequilibrium conditions. The influence of nonequilibrium is found to be greater at higher entry altitudes. With coupled ablation and carbon phenolic injection, 16 chemical species are used in the ablation layer for radiation absorption. Equilibrium and nonequilibrium results are compared under peak heating conditions.

Tiwari, S. N.↗
NASA NTRS1981

Thermal and aerothermal performance of a titanium multiwall thermal protection system

A metallic thermal protection system (TPS) concept the multiwall designed for temperature and pressure at Shuttle body point 3140 where the maximum surface temperature is approximately 811 K was tested to evaluate thermal performance and structural integrity. A two tile model of titanium multiwall and a model consisting of a low temperature reusable surface insulation (LRSI) tiles were exposed to 25 simulated thermal and pressure Shuttle entry missions. The two systems performed the same, and neither system deteriorated during the tests. It is indicated that redesign of the multiwall tiles reduces tile thickness and/or weight. A nine tile model of titanium multiwal was tested for radiant heating and aerothermodynamics. Minor design changes that improve structural integrity without having a significant impact on the thermal protection ability of the titanium multiwall TPS are identified. The capability of a titanium multiwall thermal protection system to protect an aluminum surface during a Shuttle type entry trajectory at locations on the vehicle where the maximum surface temperature is below 811 K is demonstrated.

Avery, D. E.↗
NASA NTRS1982

Wind tunnel tests of Space Shuttle external tank insulation material in the aerothermal tunnel at elevated (1440 deg F) total temperatures

Tests of the space shuttle external tank foam insulation were conducted in the von Karman Gas Dynamics Facility Tunnel C. For these tests, Tunnel C was run at Mach 4 with a total temperature of 1440 F and a total pressure which varied from 30-100 psia. Cold wall heating rates were changed by varying the test article support wedge angle and by adding and removing a shock generator or a cylindrical protuberance. Selected results are presented to illustrate the test techniques and typical data obtained.

Hartman, A. S.↗
NASA NTRS1982

Wind tunnel tests of space shuttle solid rocket booster insulation material in the aerothermal tunnel c

Wind tunnel tests of the space shuttle Solid Rocket Booster Insulation were conducted in the von Karman Gas Dynamics Facility Tunnel C. For these tests, Tunnel C was run at Mach 4 with a total temperature of 1100-1440 and a total pressure of 100 psia. Cold wall heating rates were changed by varying the test article support wedge angle. Selected results are presented to illustrate the test techniques and typical data obtained.

Hartman, A. S.↗
NASA NTRS1983

Analysis of aerothermal loads on spherical dome protuberances

Hypersonic flow over spherical dome protuberances was investigated to determine increased pressure and heating loads to the surface. The configuration was mathematically modeled in a time-dependent three-dimensional analysis of the conservation of mass, momentum (Navier-Stokes), and energy equations. A boundary mapping technique was used to obtain a rectangular parallelepiped computational domain, a MacCormack explicit time-split predictor-corrector finite difference algorithm was used to obtain solutions. Results show local pressures and heating rates for domes one-half, one, and two boundary layer thicknesses high were increased by factors on the order of 1.4, 2, and 6, respectively. Flow over the lower dome was everywhere attached while flow over the intermediate dome had small windward and leeside separations. The higher dome had an unsteady windward separation region and a large leeside separation region. Trailing vortices form on all domes with intensity increasing with dome height. Discussion of applying the results to a thermally bowed thermal protection system are presented.

Olsen, G. C.↗
NASA NTRS1983

Aerothermal environment in control surface gaps in hypersonic flow: An overview

Experimental and analytical research to develop an understanding of the fluid and thermal environment in control surface gaps such as the spanwise gap of the wing elevon and chordwise gap of a split elevon configuration typical of the Space Shuttle are summarized. Although the experimental and analytical studies were initiated too late to significantly impact the Shuttle design they do provide a fundamental understanding of the basic fluid/thermal environment in control surface gaps and help to establish a firm data base for vehicle design.

Wieting, A. R.↗
NASA NTRS1983

Aerothermal modeling program, phase 1

The physical modeling embodied in the computational fluid dynamics codes is discussed. The objectives were to identify shortcomings in the models and to provide a program plan to improve the quantitative accuracy. The physical models studied were for: turbulent mass and momentum transport, heat release, liquid fuel spray, and gaseous radiation. The approach adopted was to test the models against appropriate benchmark-quality test cases from experiments in the literature for the constituent flows that together make up the combustor real flow.

Sturgess, G. J.↗
NASA NTRS1983

Aerothermal environment in control surface gaps in hypersonic flow - An overview

Experimental and analytical research by the NASA Langley Research Center to develop an understanding of the fluid and thermal environment in control surface gaps such as the spanwise gap of the wing elevon and chordwise gap of a split elevon configuration typical of the Space Shuttle are summarized. Although the experimental and analytical studies were initiated too late to significantly impact the basic Shuttle design they do provide a fundamental understanding of the basic fluid/thermal environment in control surface gaps and help to establish a firm data base for future vehicle design.

Wieting, A. R.↗
NASA NTRS1983

Analysis of aerothermal loads on spherical dome protuberances

Hypersonic flow over spherical dome protuberances was investigated to determine increased pressure and heating loads to the surface. The configuration was mathematically modeled in a time-dependant three-dimensional analysis of the conservation of mass, momentum (Navier-Stokes), and energy equations. A boundary mapping techique was used to obtain a rectangular parallelepiped computational domain, and a MacCormack explicit time-split predictor-corrector finite difference algorithm was used to obtain solutions. Results show local pressures and heating rates for domes one-half, one, and two boundary layer thicknesses high were increased by factors on the order of 1.4, 2, and 6, respectively. However, because lee-side pressure and thermal loads were reduced the two lower height domes did not experience any net increase in total loads. The total loads on the higher dome were increased by twenty-five percent. Flow over the lower dome was everywhere attached while flow over the intermediate dome had small windward and leeside separations. The higher dome had an unsteady windward separation region and a large leeside separation region. Trailing vortices form on all domes with intensity increasing with dome height. Discussions of applying the results to a thermally bowed thermal protection system are presented.

Olsen, G. C.↗
NASA NTRS1983

Executive summary, aerothermal modeling program, phase 1

Submodels used in the combustor analytical models that were successfully used in designing advanced technology combustors were assessed. Specific recommendations for further improvement of model accuracy for combustor design purposes were made. Based upon an exhaustive literature survey, a number of test cases were selected to assess accuracy of submodels of turbulence, turbulence/chemistry interaction, spray combustion, and dilution jet mixing processes within a confined cross-stream. These test cases included simple flows and complex flows with and without swirl. Nonrecirculating and recirculating, and nonreactive and reactive flows were investigated. It was concluded that the current models give qualitative trends for the recirculating secondary flows (as encountered in a gas turbine combustor primary zone), but the predictions are good for the dilution zone.

Srinivasan, R.↗
NASA NTRS1983

Aerothermal modeling program, phase 1

The combustor performance submodels for complex flows are evaluated. The benchmark test cases for complex nonswirling flows are identified and analyzed. The introduction of swirl into the flow creates much faster mixing, caused by radial pressure gradients and increase in turbulence generation. These phenomena are more difficult to predict than the effects due to geometrical streamline curvatures, like the curved duct, and sudden expansion. Flow fields with swirl, both confined and unconfined are studied. The role of the dilution zone to achieve the turbine inlet radial profile plays an important part, therefore temperature field measurements were made in several idealized dilution zone configurations.

Srinivasan, R.↗
NASA NTRS1983

Analytical prediction of aerothermal environment in a wing-elevon cove

The fluid/thermal environment in the cove between the wing and elevon surfaces was a concern throughout the design and initial operational phases of the space shuttle. Experimenal and analytical investigations provided some understanding of the environmental in the wing-elevon cove. An incompressible finite element analysis of flow through straight parallel walls and curved parallel walls to determine the effects of cove geomtry on the fluid/thermal environment is described. Results from this analysis agree qualitatively with experimental data. The centerline gas temperatures and cold wall heating rates are virtually identical for the two cases indicating that the slight curvature has little effect on the overall thermal environment.

Bey, K. S.↗
NASA NTRS1983

Aerothermal modeling, phase 1. Volume 2: Experimental data

The experimental test effort is discussed. The test data are presented. The compilation is divided into sets representing each of the 18 experimental configurations tested. A detailed description of each configuration, and plots of the temperature difference ratio parameter or pattern factor parameter calculated from the test data are also provided.

Kenworthy, M. J.↗
NASA NTRS1984

Aerothermal loads analysis for high speed flow over a quilted surface configuration

Attention is given to hypersonic laminar flow over a quilted surface configuration that simulates an array of Space Shuttle Thermal Protection System panels bowed in a spherical shape as a result of thermal gradient through the panel thickness. Pressure and heating loads to the surface are determined. The flow field over the configuration was mathematically modeled by means of time-dependent, three-dimensional conservation of mass, momentum, and energy equations. A boundary mapping technique was then used to obtain a rectangular, parallel piped computational domain, and an explicit MacCormack (1972) explicit time-split predictor corrector finite difference algorithm was used to obtain steady state solutions. Total integrated heating loads vary linearly with bowed height when this value does not exceed the local boundary layer thickness.

Olsen, G. C.↗
NASA NTRS1984

Aerothermal environment of a blunted three-dimensional nonaxisymmetric body at Mach 6.8

Measurements of surface pressure and heat flux at Mach 6.8 for a nonaxisymmetric body, typical of the forward fuselage of an entry vehicle, have been obtained in the Langley 8-ft high-temperature tunnel. Comparisons of these data with predictions obtained from two numerical methods, which treat the test medium as an ideal gas with a laminar boundary layer, are presented. In general, the numerical results are consistent with each other and follow the data in regions of laminar flow; however, the data for angle of attack 14.9 deg indicate transitional flow on most of the windward surface. Also, the two numerical methods disagree slightly in regions of zero or nearly zero pressure gradients.

Martin, F. W., Jr.↗
NASA NTRS1984

Aerothermal loads analysis for high speed flow over a quilted surface configuration

Attention is given to hypersonic laminar flow over a quilted surface configuration that simulates an array of Space Shuttle Thermal Protection System panels bowed in a spherical shape as a result of thermal gradients through the panel thickness. Pressure and heating loads to the surface are determined. The flow field over the configuration was mathematically modeled by means of time-dependent, three-dimensional conservation of mass, momentum, and energy equations. A boundary mapping technique was then used to obtain a rectangular, parallelepiped computational domain, and an explicit MacCormack (1972) explicit time-split predictor-corrector finite difference algorithm was used to obtain steady state solutions. Total integrated heating loads vary linearly with bowed height when this value does not exceed the local boundary layer thickness.

Olsen, G. C.↗