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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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Time dependent pressure distribution and threshold acceleration for bubble formation in longitudinally vibrating flexible liquid filled cylinder
An aerodynamic investigation was conducted in the MSFC High Reynolds Number Wind Tunnel to determine the pressure distribution over the foresection of the current 146 inch diameter shuttle SRB. The test model consisted of a 0.0137 scale version of the SRB nose cone and a forward portion of the cylindrical body which was approximately 2.7 calibers in length. The pressure distributions are plotted as a function of longitudinal station ratioed to body diameter and circumferential location for each angle of attack and Mach number. A Reynolds number variation study was made for Mach numbers of 0.4 and 0.6 at an angle of attack of 270 deg and roll angle of 180 deg.
A Neumann solution for inviscid external flow was coupled to a modified Reshotko-Tucker integral boundary-layer technique, the control volume method of Presz for calculating flow in the separated region, and an inviscid one-dimensional solution for the jet exhaust flow in order to predict axisymmetric nozzle afterbody pressure distributions and drag. The viscous and inviscid flows are solved iteratively until convergence is obtained. A computer algorithm of this procedure was written and is called DONBOL. A description of the computer program and a guide to its use is given. Comparisons of the predictions of this method with experiments show that the method accurately predicts the pressure distributions of boattail afterbodies which have the jet exhaust flow simulated by solid bodies. For nozzle configurations which have the jet exhaust simulated by high-pressure air, the present method significantly underpredicts the magnitude of nozzle pressure drag. This deficiency results because the method neglects the effects of jet plume entrainment. This method is limited to subsonic free-stream Mach numbers below that for which the flow over the body of revolution becomes sonic.
Tests were conducted in the 2-D test section of the Langley 0.3-meter Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th power. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Tabulated pressure distributions and integrated force and moment coefficients are presented as well as plots of the surface pressure distributions. The data are presented uncorrected for wall interference effects and without analysis.
Heat-transfer and pressure distributions on 60 and 70 deg swept delta wings with turbulent boundary layers
Incompressible two dimensional turbulent boundary layer equations with arbitrary pressure distribution solved by weighted residual method
Incompressible two dimensional turbulent boundary layer equations with arbitrary pressure distribution solved by weighted residual method
The turbulent flow of a subsonic round jet exhausting perpendicularly from a flat plate into a subsonic crosswind of the same temperature was investigated in the Langley V/STOL tunnel. The large test section of this tunnel made it possible to use a jet with the relatively large diameter of 10.16 cm. Pressures were measured on the flat plate at over 400 locations to provide a detailed description of the static pressure distribution. Results are presented in tabular and graphical forms for jet to crossflow velocity ratios ranging from 2 to 10, and comparisons are made with results of other experiments. The results indicate that the dominant flow parameter affecting the pressure distributions is the ratio of jet to crossflow velocities. The distribution of turbulence on the plate also appears to be a function of velocity ratio.
Pressure distributions due to fabrication-type surface roughness on ogive cylinder at transonic speeds
Local pressure distribution on a blunt delta wing
A ceramic guide vane has been designed and tested for operation under high temperature. Previous efforts have suggested that some cooling flow may be required to alleviate the high temperatures observed near the trailing edge region. The present report describes briefly a three-dimensional viscous analysis carried out to calculate the temperature and pressure distribution on the blade surface and in the flow path with a jet of cooling air exiting from the suction surface near the trailing edge region. The data for analysis was obtained from Dr. Craig Robinson. The surface temperature and pressure distribution along with a flowfield distribution is shown in the results. The surface distribution is also given in a tabular form at the end of the document.
Effects of fluid mass flow rate and exit orifice diameter on radial static pressure distribution in compressible vortex flow field
Experimental and predicted wall static-pressure distributions in conical supersonic nozzles
An investigation has been made at Mach numbers of 1.61 and 2.01 and over a range of free-stream Reynolds number per foot from about 1.2 x 10(exp 6) to 8.3 x 10(exp 6) to determine the pressure distributions and wave drags due to two-dimensional fabrication-type surface roughness. Ten types of surface roughness, including step, wave, crease, and swept configurations were investigated. The tests were made on an ogive cylinder of fineness ratio 12.2, the roughness elements covering the cylindrical portion of the model. The results indicate that wave drag is the major component of the drag due to roughness at supersonic speeds. The pressure distributions over the roughness elements were generally found to be in good agreement with linearized two-dimensional theory except for regions of the elements affected by boundary-layer separation and shock detachment. There was little or no effect of Reynolds number except on the pressures within the regions influenced by separation or shock detachment. Inasmuch as most of the roughness configurations were affected by flow separation and shock detachment, there was generally an effect of Reynolds number on the roughness wave drag. This wave drag decreased as the free-stream Reynolds number was decreased.
Local heat-transfer data and static-pressure distributions for Freon-113 condensing inside vertical tube
Pressure distributions over forward portion of project fire space-vehicle configuration at subsonic speed
Aerodynamic characteristics of reentry body - pressure distribution at angles of attack and trajectory analysis
A 2.29 m (7.5 ft.) span high-lift research model equipped with full-span leading-edge slat and part-span double-slotted trailing-edge flap was tested in the Langley 4- by 7-Meter Tunnel to determine the low speed performance characteristics of a representative high aspect ratio suprcritical wing. These tests were performed in support of the Energy Efficient Transport (EET) program which is one element of the Aircraft Energy Efficiency (ACEE) project. Static longitudinal forces and moments and chordwise pressure distributions at three spanwise stations were measured for cruise, climb, two take-off flap, and two landing flap wing configurations. The tabulated and plotted pressure distribution data is presented without analysis or discussion.