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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 NTRS1976

Pressure distributions for a rectangular supersonic inlet at subsonic speeds

Pressure distribution data are provided for a supersonic rectangular inlet at subsonic speeds. Variations in cowl and ramp geometry as well as sideplate sweep were investigated. Tests were made in the Langley 16-foot transonic tunnel and the Langley high speed 7- by 10-foot tunnel for Mach numbers of 0.6, 0.7, and 0.8. Angles of attack investigated were 0 deg, 4 deg, and 8 deg for a range of mass flow ratios.

Fuller, D. E.↗
NASA NTRS1955

Theoretical Prediction of Pressure Distributions on Nonlifting Airfoils at High Subsonic Speeds

Theoretical pressure distributions on nonlifting circular-arc airfoils in two-dimensional flows with high subsonic free-stream velocity are found by determining approximate solutions, through an iteration process, of an integral equation for transonic flow proposed by Oswatitsch. The integral equation stems directly from the small-disturbance theory for transonic flow. This method of analysis possesses the advantage of remaining in the physical, rather than the hodograph, variable and can be applied in airfoils having curved surfaces. After discussion of the derivation of the integral equation and qualitative aspects of the solution, results of calculations carried out for circular-arc airfoils in flows with free-stream Mach numbers up to unity are described. These results indicate most of the principal phenomena observed in experimental studies.

Spreiter, John R↗
NASA NTRS1989

Pressure Distribution and Air Data System for the Aeroassist Flight Experiment

The Aeroassist Flight Experiment (AFE) is designed to provide critical flight data necessary for the design of future Aeroassist Space Transfer Vehicles (ASTV). This flight experiment will provide aerodynamic, aerothermodynamic, and environmental data for verification of experimental and computational flow field techniques. The Pressure Distribution and Air Data System (PD/ADS), one of the measurement systems incorporated into the AFE spacecraft, is designed to provide accurate pressure measurements on the windward surface of the vehicle. These measurements will be used to determine the pressure distribution and air data parameters (angle of attack, angle of sideslip, and free-stream dynamic pressure) encountered by the blunt-bodied vehicle over an altitude range of 76.2 km to 94.5 km. Design and development data are presented and include: measurement requirements, measurement heritage, theoretical studies to define the vehicle environment, flush-mounted orifice configuration, pressure transducer selection and performance evaluation data, and pressure tubing response analysis.

Gibson, Lorelei S.↗
NASA NTRS1961

Low Subsonic Pressure Distributions on Three Rigid Wings Simulating Paragliders with Varied Canopy Curvature and Leading-Edge Sweep

An investigation has been made in the Langley 7- by 10-foot transonic tunnel to determine the subsonic pressure distribution of three paraglider models through an angle-of-attack range from 0 deg to 74 deg. Three rigid metal models simulated a 45 deg basic flat planform paraglider with leading-edge sweep angles of 61.6 deg, 52.5 deg, and 48.6 deg. These configurations resulted in one-half-circle, one-third-circle, and one-quarter-circle semispan trailing-edge curvature when viewed from downstream. The results of the investigation are presented as curves of chordwise pressure distributions at four spanwise locations.

Fournier, Paul G.↗
NASA NTRS1930

Calculation of Pressure Distribution on Airship Hulls

These calculations were based on the shape of the ZR III, with the following simplifications: cars, fins, and rudders removed; all cross sections replaced by equivalent circular cross sections. Under these assumptions the pressure distribution was calculated for the following cases: symmetrical case, or flow parallel to the axis; unsymmetrical case, or flow at an angle to the axis. In both cases the simple potential flow first forms the basis for the determination of the pressure distribution.

Von Karman, Theodor↗
NASA NTRS1961

The Effect of Air Bleed on the Heat Transfer and Pressure Distribution on 30 Conical Flares at a Mach Number of 6.8

The effect of a boundary-layer bleed at the start of a 30 deg half-angle flare upon the shape of the flow boundaries, the pressure distribution on the flare, and the heat transfer to the flare was studied at a Mach number of 6.8. The forebody was an ogive cylinder. Test Reynolds numbers, based on forebody length, ranged from 1 x 10(exp 6) to 7.4 x 10(exp 6). Schlieren photographs showed the effect produced upon the flow boundaries by varying the dimensions of the bleed in both the radial and axial directions and by blunting the lip at the leading edge of the flare. The heat transfer and pressure distribution on the flare were correlated with the shape and nature of the flow boundaries.

Crawford, Davis H.↗
NASA NTRS1976

Pressure distribution on a 1- by 3-meter semispan wing at sweep angles from 0 deg to 40 deg in subsonic flow

A 1- by 3-meter semispan wing of taper ratio 1.0 with NACA 0012 airfoil section contours was tested in the Langley V/STOL tunnel to measure the pressure distribution at five sweep angles, 0 deg, 10 deg, 20 deg, 30 deg, and 40 deg, through an angle-of-attack range from -6 deg to 20 deg. The pressure data are presented as plots of pressure coefficients at each static-pressure tap location on the wing. Flow visualization wing-tuft photographs are also presented for a wing of 40 deg sweep. A comparison between theory and experiment using two inviscid theories and a viscous theory shows good agreement for pressure distributions, normal forces, and pitching moments for the wing at 0 deg sweep.

Yip, L. P.↗
NASA NTRS1976

An investigation to determine the pressure distribution on the 0.0137 scale solid rocket booster forebody (MSFC model 467) at angles of attack at or near 90 deg and high Reynolds numbers in the MSFC High Reynolds Number Wind Tunnel (SA29F)

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.

Ramsey, P. E.↗
NASA NTRS1979

DONBOL: A computer program for predicting axisymmetric nozzle afterbody pressure distributions and drag at subsonic speeds

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.

Putnam, L. E.↗