Space Shuttle Aerothermodynamics Technology Conference. Volume 2: Heating
Conference on aerodynamic heat transfer in space shuttle configurations and procedures and techniques for measuring transfer - Vol. 2
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Aerothermodynamic measurements for space shuttle configuration in hypersonic wind tunnels
The effect of shuttle configuration geometry, angle of attack, and free stream flow conditions on the heat-transfer distribution as influenced by three-dimensional effects, the wing-fuselage shock-interaction, and resultant wing-impingement phenomena are examined. In addition, the data provided information regarding the flow field in the vicinity of the nose and boundary layer transition in the plane of symmetry of the fuselage. The data included measurements of the surface pressure, the heat transfer rate distributions, (using models instrumented with thermocouples and models painted with thermographic phosphor) and schlieren and shadowgraph photographs. Posttest photographs of the painted models supplemented the heat transfer data.
Bertin, J. J.↗
Determination of aerothermodynamic environment uncertainties with application to space shuttle vehicles
In an effort to minimize the cost of the thermal protection system (TPS) for space shuttles, an analysis is made of environmental uncertainties. The basic categories of these uncertainties include aerodynamic heating, trajectory dispersions, atmospheric variations, and vehicle attitude variations. Also discussed are TPS design sensitivity to environmental uncertainties, and uncertainty profiles. Preliminary results indicate moderate TPS cost for a high confidence level environment may be obtained.
Scottoline, C. A.↗
Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development
The analytical effort was directed towards (1) completing the design of the combustor exit instrumentation assembly, (2) analyzing the coolant flow distribution of the cowl leading edge tip section, (3) determining effects of purge gas pressure on AIM performance analysis, and (4) analyzing heat transfer and associated stress problems related to the cowl leading edge tip section and the nozzle shroud assembly for test conditions.
Jilly, L. F.↗
Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development
The analytical effort was directed toward (1) analyzing the latest inputs of possible test conditions and their impact on hardware; (2) review structural capabilities of the AIM unit; (3) analyzing coolant flow paths, heat transfer, and associated problems relative to the cowl leading edge tip section; and (4) redefining requirements of flow-calibrating the air-metering duct.
Jilly, L. F.↗
Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development
The fabrication of the various components of the HRE AIM was completed. The purge system necessary for the cavity bounded by the outer shell assembly and the outer cowl body was studied. Preparations were begun for establishing a format for test data acquisition and reduction.
Jilly, L. F.↗
A study for development of aerothermodynamic test model materials and fabrication technique
A literature survey, materials reformulation and tailoring, fabrication problems, and materials selection and evaluation for fabricating models to be used with the phase-change technique for obtaining quantitative aerodynamic heat transfer data are presented. The study resulted in the selection of two best materials, stycast 2762 FT, and an alumina ceramic. Characteristics of these materials and detailed fabrication methods are presented.
Dean, W. G.↗
The effect of nose geometry on the aerothermodynamic environment of shuttle entry configurations
The effect was studied of nose geometry on the transition criteria for the windward boundary layer, on the extent of separation, on the heat transfer perturbation due to the canopy, and on the surface pressure and the heat transfer in the separated region. The data for each of these problems is analyzed. A literature review that concentrates on separation and the leeward flow-field is presented.
Bertin, J. J.↗
Aerothermodynamic aspects of shock-interference patterns for shuttle configurations during entry.
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Bertin, J. J.↗
Effect of nose geometry on the aerothermodynamic environment of shuttle entry configurations.
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Bertin, J. J.↗
Developing a technology base in planetary entry aerothermodynamics
The long-range objectives of the entry technology program are to insure that an adequate technology base for a great variety of mission options exists. Consideration has been given to the entry of vehicles into the atmospheres of all the planets with the exception of Pluto. The experimental facilities for the studies are discussed, giving attention to shock tubes, the planetary entry radiation facility, ballistic ranges, the expansion tube, and arc jet facilities. Flight experiments are considered along with computational analyses and engineering approximations.
Olstad, W. B.↗
Outer planet probe entry thermal protection. I Aerothermodynamic environment
Solutions are presented for the aerothermal heating environments about blunt probes of various shape entering the candidate atmospheres of Saturn and Uranus. The analysis makes use of a series of interacting codes to predict pressures, cold wall radiative fluxes, cold wall convective fluxes, and corresponding corrections for the radiative and convective fluxes. The solutions show that ablation will either enhance or reduce the radiative flux to the surface depending on the ablation gas composition and ablation rate. Other conclusions suggested by the solutions are discussed.
Nicolet, W. E.↗
Hypersonic research engine project. Phase 2: Aerothermodynamic integration model, test plan, data item no. 2-2.01
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Jilly, L. F.↗
Results of aerothermodynamic heating tests on a 0.013-scale model solid rocket booster in the NASA/LaRC unitary plan wind tunnel (SH12F)
A 0.013 scale model of the solid rocket booster (SRB) used to launch the space shuttle was tested at a Mach number of 3.7 and Reynolds numbers of 1,500,000 and 3,500,000 per foot. The objective of the test was to obtain aerodynamic heat transfer data on the surface of scaled models of the SRB at simulated full scale reentry flight conditions. Three separate models were utilized to measure film coefficients over an angle of attack range from 0 deg to 180 deg at 0 deg sideslip. All three models were representations of the MCR0200 baseline configuration and varied only by the way they were mounted in the tunnel. Model A, sting mounted thru the model base, was utilized for testing between 0 deg and 40 deg angle of attack. Model B was blade mounted from the top of the model and was tested between 60 deg and 120 deg angle of attack. Model C was sting mounted thru the model nose and utilized for testing between 140 deg and 180 deg angle of attack.
Brewer, E. B.↗
Planetary entry aerothermodynamics. I - Technology requirements and experimental facilities
Consideration has been given to the entry of vehicles into the atmospheres of all planets with the exception of Pluto and to missions for returning samples to Earth from all the planets (excepting Pluto), the satellites of Mars, Jupiter, and Saturn, the Moon, the asteroids, and certain comets. The efficient planning of the space missions including the appropriate design of the spacecraft requires that operational and environmental parameter values for the large range of mission options included be adequately known. The need to obtain the data required as a basis for the computational studies involved makes it necessary to conduct a comprehensive experimental research program. Details concerning the experimental facilities employed in this program are discussed, giving attention to shock tubes, the planetary-entry radiation facility, ballistic ranges, an expansion tube, a hypersonic tetrafluoromethane tunnel, and arcjet facilities.
Olstad, W. B.↗
Planetary entry aerothermodynamics. II - Computational analyses and flight experience
Limitations concerning the possibility to simulate all the significant flow and thermal phenomena occurring during the entry of a space vehicle into a planetary atmosphere make it necessary to rely on computational analyses to obtain the required data for the design of the spacecraft needed for the NASA missions planned for the next two decades. 'Benchmark' computer programs concerned with complete, detailed, and accurate computational solutions of entry problems are considered along with programs representing engineering approximations for cases in which the accuracy provided by the benchmark programs is not needed. The information obtainable by computational analysis has to be supplemented by actual flight experience in order to meet the goals of the NASA entry-technology program. The individual space missions planned for the coming years are examined together with the possibilities for obtaining the data needed to satisfy the entry requirements in each case.
Olstad, W. B.↗
Aerothermodynamic environment for Jovian entry conditions
Solutions are presented for the aerothermal heating environments about blunt probes entering the atmosphere of Jupiter. The analysis employs a novel prediction procedure in which finite differences versions of the thin layer fluid dynamic equations are solved simultaneously with, and coupled to, a fully detailed radiation flux model. Predictions are presented and compared with those of earlier investigators. Radiative blowing corrections are presented for on- and off-stagnation point conditions. The use of a blowing correlation based on the retained mass was found to be useful for off-stagnation point conditions.
Nicolet, W. E.↗
Preliminary analysis of remote infrared imagery of shuttle during entry: An aerothermodynamic flight experiment
The preliminary feasibility of remote high-resolution infrared imagery of the space shuttle orbiter lower surface during entry to obtain accurate measurements of aerodynamic heat transfer to that vehicle was examined. In general, it was determined that such such images can be taken from an existing aircraft/telescope system (the C-141 AIRO) with a minimum modification or addition of systems using available technology. These images will have a spatial resolution of about 0.3 m and a temperature resolution much better than 2.5 percent. The data from these images will be at conditions and at a scale not reproducible in ground based facilities and should aid in the reduction of the prudent factors of safety required to account for phenomenological uncertainties on the thermal protection system design. Principal phenomena to be observed include laminar heating, boundary-layer transition, turbulent heating, surface catalysis, and flow separation and reattachment.
Swenson, B. L.↗