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At least 829 records · Page 46
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Laminar boundary layer on a cone near a plane of symmetry.

The three-dimensional laminar boundary layer near the planes of symmetry of a cone with either outflow or inflow is studied in supersonic and hypersonic flow. Solutions are presented for a variety of values of three parameters, namely those of cross flow, external Mach number, and wall enthalpy ratio. Also asymptotic solutions for ranges of pertinent similarity variables are developed and used in numerical analyses so as to assure exponential behavior as the similarity variables approach infinity. Of special interest in the results is the existence for outflow, i.e., for the windward plane of symmetry, of a new solution branch and for inflow of four solution branches. The physical significance of the new solutions is discussed.

Wu, P.↗
NASA NTRS1974

Aerodynamic heating and loading within large open cavities in cone and cone-cylinder-flare models at Mach 6.7

The aerodynamic heating and loading distributions within large cavities exposed by surface openings to hypersonic flow were determined in the 8-foot high-temperature structures tunnel. Cone and cone-cylinder-flare models with cavities having regular and irregular surface-opening shapes were aerothermally tested at various angles of attack up to 30 deg. Tests were conducted at a Mach number of 6.7, a total temperature of 1800 K, a dynamic pressure of 80 kPa, and a stream unit Reynolds number of 6,000,000 per meter. The results showed that the heating rates at internal reattachment were generally lower but of the same order as the corresponding external heating rates; however, other internal heating rates were an order of magnitude lower. The internal flow showed characteristics of being funneled into jets or being dispersed dependent upon the internal surface contour.

Hunt, R. L.↗
NASA NTRS1974

Results of tests in the NASA/LARC 31-inch CFHT on an 0.010-scale model (32-OT) of the space shuttle configuration 3 to determine the RCS jet flowfield interaction effects on aerodynamic characteristics (IA60/OA105), volume 1

Tests were conducted in the NASA Langley Research Center 31-inch continuous Flow Hypersonic Wind Tunnel to determine RCS jet interaction effect on the hypersonic aerodynamic and stability and control characteristics prior to return to launch site (RTLS) abort separation. The model used was an 0.010-scale replica of the Space Shuttle Vehicle Configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the integrated Orbiter and external tank and the Orbiter alone. RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 7, 20, and 50 psf were investigated. The effects of speedbrake, bodyflap, elevon, and aileron deflections were also investigated.

Thornton, D. E.↗
NASA NTRS1974

Results of tests in the NASA/LaRC 31-inch CFHT on an 0.010-scale model (32-OT) of the space shuttle configuration 3 to determine the RCS jet flowfield interaction effects on aerodynamic characteristics (IA60/0A105), volume 2

Tests were conducted in the NASA Langley Research Center 31-inch continuous flow hypersonic wind tunnel from 14 February to 22 February 1974, to determine RCS jet interaction effect on the hypersonic aerodynamic and stability and control characteristics prior to RTLS abort separation. The model used was an 0.010-scale replica of the space shuttle vehicle configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the intergrated orbiter and external tank and the orbiter alone. RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 7, 20, and 50 psf were investigated. The effects of speedbrake, bodyflap, elevon, and aileron deflections were also investigated.

Thornton, D. E.↗
NASA NTRS1974

Results of tests in the NASA/LaRC 31-inch CFHT on an 0.010-scale model (32-OT) of the space shuttle configuration 3 to determine the RCS jet flowfield interaction effects on aerodynamic characteristics (IA60/0A105), volume 1

Tests were conducted in the 31-inch continuous Flow Hypersonic Wind Tunnel to determine RCS jet interaction effect on the hypersonic aerodynamic and stability and control characteristics prior to RTLS abort separation. The model used was an 0.010-scale replica of the Space Shuttle Vehicle Configuration 3. Hypersonic stability data were obtained from tests at Mach 10.3 and dynamic pressure of 150 psf for the integrated Orbiter and external tank and the Orbiter alone. RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 7, 20, and 50 psf were investigated. The effects of speedbrake, bodyflap, elevon, and airleron deflections were also investigated.

Thornton, D. E.↗
NASA NTRS1974

Aerodynamic heating on 3-D bodies including the effects of entropy-layer swallowing

A relatively simple method was developed previously (authors, 1973) for calculating laminar, transitional, and turbulent heating rates on three-dimensional bodies in hypersonic flows. This method was shown to yield reasonably accurate results for laminar heating on blunted circular and elliptical cones and an earlier version of the space shuttle vehicle. As the boundary layer along the surface grows, more and more of the inviscid-flow mass is entrained into the boundary layer, and the streamlines which passed through the nearly normal portion of the bow shock wave are 'swallowed' by the boundary layer. This phenomenon is often referred to as entropy-layer or streamline swallowing, and it can have a significant effect on the calculated heating rates. An approximate, yet simple, method for including the effects of entropy-layer swallowing in the heating-rate calculations is given.

Dejarnette, F. R.↗
NASA NTRS1974

Techniques for aerothermal tests of large, flightweight thermal protection panels in a Mach 7 wind tunnel

Recently developed experimental techniques permit evaluating thermal performance and structural integrity of full-scale panel concepts applicable to reentry and hypersonic vehicles in the Langley 8-foot high-performance structures tunnel. Rate and magnitude of surface heating and differential-pressure loading are independently controlled. Realistic temperature distributions are radiantly preheated into the panel prior to aerodynamic heating, and stream conditions are preselected to sustain the preheat surface heating input during aerodynamic exposure. During tunnel start and shutdown, panels are shielded outside the stream from potentially damaging transient acoustics and buffeting and are then rapidly inserted into the hypersonic flow. Infrared radiometry provides detailed surveys of surface temperatures.

Deveikis, W. D.↗
NASA NTRS1975

Results of test 0A82 in the NASA/LRC 31 inch CFHT on an 0.010-scale model (32-0) of the space shuttle configuration 3 to determine RCS jet flow field interaction and to investigate RT real gas effects

Tests were conducted in the NASA Langley Research Center 31-inch Continuous Flow Hypersonic Wind Tunnel to determine RCS jet interaction effects on hypersonic aerodynamic characteristics and to investigate RT (gas constant times temperature) scaling effects on the RCS similitude. The model was an 0.010-scale replica of the Space Shuttle Orbiter Configuration 3. Hypersonic aerodynamic data were obtained from tests at Mach 10.3 and dynamic pressures of 200, 150, 125, and 100 psf. The RCS modes of pitch, yaw, and roll at free flight dynamic pressure simulation of 20 psf were investigated.

Thornton, D. E.↗
NASA NTRS1975

Space shuttle orbiter trimmed center of gravity extension study. Volume 1: Effects of configuration modifications on the aerodynamic characteristics of the 140 A/B orbiter at Mach 10.3

Longitudinal and lateral-directional characteristics were obtained on several modified versions of the Rockwell International 140 A/B orbiter (0.010 scale). These modifications, designed to extend trim capability to center-of-gravity locations forward of the 65 percent fuselage station, consisted of two forebodies, two canard trimmers, and two body-wing fillets. Tests were performed over an angle-of-attack range of 12 to 36 degrees at a Reynolds number of 1.03 million based on body reference length. Data were obtained with the elevons and body flap deflected at the full-up and full-down design values. This investigation was conducted in the NASA Langley continuous flow hypersonic tunnel.

Bernot, P. T.↗
NASA NTRS1975

Unified nonlinear approach to both weak and strong-interaction problems

The numerical results are obtained for heat transfer, skin-friction, and viscous interaction induced pressure for a step-wise accelerated flat plate in hypersonic flow. In the unified approach here the results are presented for both weak and strong-interaction problems without employing any linearization scheme. With the help of the numerical method used in this work an accurate prediction of wall shear can be made for the problems with plate velocity changes of 1% or larger. The obtained results indicate that the transient contribution to the induced pressure for helium is greater than that for air.

Gupta, R. N.↗
NASA NTRS1976

Results of test MA22 in the NASA/LaRC 31-inch CFHT on an 0.010-scale model (32-0) of the space shuttle configuration 3 to determine RCS jet flow field interaction, volume 1

A wind tunnel test was conducted in the Langley Research Center 31-inch Continuous Flow Hypersonic Wind Tunnel from May 6, 1975 through June 3, 1975. The primary objectives of this test were the following: (1) to study the ability of the wind tunnel to repeat, on a run-to-run basis, data taken for identical configurations to determine if errors in repeatability could have a significant effect on jet interaction data, (2) to determine the effect of aerodynamic heating of the scale model on jet interaction, (3) to investigate the effects of elevon and body flap deflections on jet interaction, (4) to determine if the effects from jets fired separately along different axes can be added to equal the effects of the jets fired simultaneously (super position effects), (5) to study multiple jet effects, and (6) to investigate area ratio effects, i.e., the effect on jet interaction measurements of using wind tunnel nozzles with different area ratios in the same location. The model used in the test was a .010-scale model of the Space Shuttle Orbiter Configuration 3. The test was conducted at Mach 10.3 and a dynamic pressure of 150 psf. RCS chamber pressure was varied to simulate free flight dynamic pressures of 5, 7.5, 10, and 20 psf.

Kanipe, D. B.↗
NASA NTRS1976

Performance of a mullite reusable surface insulation system in a hypersonic stream

The thermal and structural performance of a large panel of mullite reusable surface insulation (RSI) tiles was determined by a series of aerothermal tests in the Langley 8-foot high-temperature structures tunnel. The test panel was designed to represent a portion of the surface structure on a space shuttle orbiter fuselage along a 1,150 K isotherm with the mullite tile system bonded directly to the primary structure. Aerothermal tests were conducted at a free-stream Mach number of 6.7, a total temperature of 1,880 K, a unit Reynolds number of 4.6 million per meter, and dynamic pressure of 62 kPa. The thermal response of the mullite tile was as predicted, and the bond-line temperature did not exceed the design level of 570 K during a typical entry-heat cycle. Geometric irregularities of the tile gaps affected the tile edge temperatures when exposed to hypersonic flow. The tile coating demonstrated good toughness to particle impacts, but the coating cracked and flaked with thermal cycles. The gap filler of woven silica fibers appeared to hinder flow penetration into the gaps and withstood the flow shear of the present tests.

Hunt, L. R.↗
NASA NTRS1979

Some recent developments in the prediction of shock interaction phenomena at hypersonic speeds

The shock strengths for which either Edney type I or type II shock interference patterns can occur when two oblique shocks of opposite families intersect were determined graphically at Mach 10 by using logarithmic shock polar diagrams. The theoretical region of overlap for the two types of interaction was investigated by observing in the Schlieren system of the Langley 15-inch hypersonic flow apparatus the intersection of oblique shocks generated by two sharp 10 degree wedges as the wedge angles of attack and their relative positions were altered. A range of shock strengths for which either of the two interference patterns can exist was demonstrated.

Crawford, D. H.↗
NASA NTRS1982

Numerical Solution for Navier-Stokes Equations

Carefully selected blend of computational techniques solves complete set of equations for viscous, unsteady, hypersonic flow in general curvilinear coordinates. New algorithm has tested computation of axially directed flow about blunt body having shape similar to that of such practical bodies as wide-body aircraft or artillery shells. Method offers significant computational advantages because of conservation-law form of equations and because it reduces amount of metric data required.

Warsi, Z. U. A.↗
NASA NTRS1981

Ceramic and coating applications in the hostile environment of a high temperature hypersonic wind tunnel

A Mach 7, blowdown wind tunnel was used to investigate aerothermal structural phenomena on large to full scale high speed vehicle components. The high energy test medium, which provided a true temperature simulation of hypersonic flow at 24 to 40 km altitude, was generated by the combustion of methane with air at high pressures. Since the wind tunnel, as well as the models, must be protected from thermally induced damage, ceramics and coatings were used extensively. Coatings were used both to protect various wind tunnel components and to improve the quality of the test stream. Planned modifications for the wind tunnel included more extensive use of ceramics in order to minimize the number of active cooling systems and thus minimize the inherent operational unreliability and cost that accompanies such systems. Use of nonintrusive data acquisition techniques, such as infrared radiometry, allowed more widespread use of ceramics for models to be tested in high energy wind tunnels.

Puster, R. L.↗
NASA NTRS1982

A hot-film static-pressure probe

The procedure described involves calibrating the probe in the gas of interest over the range of temperatures and pressures anticipated in the wind-tunnel tests and then applying the calibration to reduce the data from those tests. Pressure readings from two probes, one having two hot-film sensors and an internal sonic orifice and the other connected to a conventional pressure gage, are recorded for several Reynolds numbers and compared. The conventional static-pressure probe with a pair of hot-film sensors and an interior sonic orifice is shown to provide the means of surveying flowfields in hypersonic flow at a rapid rate and with an accuracy of better than + or - 10%.

Ashby, G. C., Jr.↗
NASA NTRS1982

Development and application of the GIM code for the Cyber 203 computer

The GIM computer code for fluid dynamics research was developed. Enhancement of the computer code, implicit algorithm development, turbulence model implementation, chemistry model development, interactive input module coding and wing/body flowfield computation are described. The GIM quasi-parabolic code development was completed, and the code used to compute a number of example cases. Turbulence models, algebraic and differential equations, were added to the basic viscous code. An equilibrium reacting chemistry model and implicit finite difference scheme were also added. Development was completed on the interactive module for generating the input data for GIM. Solutions for inviscid hypersonic flow over a wing/body configuration are also presented.

Stainaker, J. F.↗
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.↗