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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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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
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.
Pressure distributions on four nacelle cowl models of the same length and highlight area but different geometries external to the highlight are compared. The diameter ratio (ratio of highlight diameter to maximum diameter) of the four cowls was 0.854 and the length ratio (ratio of cowl length to maximum diameter) was 0.439. The cowls had the same internal geometry from the highlight to the throat with a contraction ratio (ratio of highlight area to throat area) of 1.250. Data for two other cowls which had a diameter ratio of 0.880, a length ratio of 0.400 and a contraction ratio 1.250 are also included. All the cowls had rows of static pressure orifices on the top and bottom surfaces. Mass-flow ratio was varied between 0.27 and 0.93. Some data were obtained between angles of attack from -2.1deg and 4.1deg. The test was conducted in the Langley 16-Foot Transonic Tunnel.
Pressure distributions and aerodynamic characteristics of spoiler-type controls on trapezoidal wing at supersonic speed
Flight pressure distributions on vertical stabilizers and speed brakes of X-15 aircraft at Mach 1 to 6
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Propeller slipstream effects as determined from wing pressure distribution of a large-scale six-propeller vtol model at static thrust
Iterative method for designing supersonic body of revolution from surface pressure distribution
Heat transfer and pressure distribution over reentry configuration for five-stage scout vehicle