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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch

Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a nonlinear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured results. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.

Ricketts, R. H.↗
NASA NTRS1961

Ordinates and Theoretical Pressure-Distribution Data for NACA 6- and 6A-Series Airfoil Sections with Thicknesses from 2 to 21 and From 2 to 15 Percent Chord, Respectively

Information is presented with which ordinates can be easily obtained for any thickness from 2 to 21 percent chord for NACA 63-, 64-, and 65-series airfoil sections and from 2 to 15 percent chord for NACA 63A-, 64-A, series airfoil sections. In addition, data required for estimation of the theoretical pressure distributions of any of these airfoils are included.

Patterson, Elizabeth W.↗
NASA NTRS1991

Measurements of Pressure Distributions and Force Coefficients in a Squeeze Film Damper. Part 2: Partially Sealed Configuration

Experimental results from a partially sealed squeeze film damper (SFD) test rig, executing a circular centered orbit are presented and discussed. A serrated piston ring is installed at the damper exit. This device involves a new sealing concept which produces high damping values while allowing for oil flow to cool the damper. In the partially sealed damper, large cavitation regions are observed in the pressure fields at orbit radii epsilon equals 0.5 and epsilon equals 0.8. The cavitated pressure distributions and the corresponding force coefficients are compared with a cavitated bearing solution. The experimental results show the significance of fluid inertia and vapor cavitation in the operation of squeeze film dampers. Squeeze film Reynolds numbers tested reach up to Re equals 50, spanning the range of contemporary applications.

Jung, S. Y.↗
NASA NTRS1973

Pressure distribution and velocity profiles in supersonic laminar wakes.

The method of integral relations is applied to calculate the characteristics of laminar wakes in supersonic flow past a wedge with a flat base. An earlier analysis of the flow in the critical recompression region is extended so that the flow near the base can be determined. Results are worked out for flow past a wedge of 10-deg semiangle at Mach number 6 and two values of Reynolds number. The pressure distribution along the axis is calculated, together with velocity profiles at various stations and streamline patterns in the recirculation region. Good agreement with experiment is found.

Holt, M.↗
NASA NTRS1984

A comparison of experimental and theoretical results for leakage, pressure distribution, and rotordynamic coefficients for annular gas seals

The importance of seal behavior in rotordynamics is discussed and current annular seal theory is reviewed. A Nelson's analytical-computational method for determining rotordynamic coefficients for this type of compressible-flow seal is outlined. Various means for the experimental identification of the dynamic coefficients are given, and the method employed at the Texas A and M University (TAMU) test facility is explained. The TAMU test apparatus is described, and the test procedures are discussed. Experimental results, including leakage, entrance-loss coefficients, pressure distributions, and rotordynamic coefficients for a smooth and a honeycomb constant-clearance seal are presented and compared to theoretical results from Nelson's analysis. The results for both seals show little sensitivity to the running speed over the test range. Agreement between test results and theory for leakage through the seal is satisfactory. Test results for direct stiffness show a greater sensitivity to fluid pre-rotation than predicted. Results also indicate that the deliberately roughened surface of the honeycomb seal provides improved stability versus the smooth seal.

Nicks, C. O.↗
NASA NTRS1947

Pressure Distribution Measurements at High Speed and Oblique Incidence of Flow

The present report contains the results of a series of observations obtained for a wing of symmetrical profile for different angles of yaw. The shock tunnel with 0.4m x 0.4m cross section, of the Institute for Gas Dynamics at L.F.A. - Braunchweig was available for this work. The profile used was a 9 percent, thick hyperbolic profile with maximum thickness 40 percent aft, which was calculated by the method of F. Ringleb (2). Since the conformal transformation of this profile is known, the theoretical pressure distribution could be determined exactly for the case of incompressible flow. Then by the use of the Prandtl rule one may extend the comparison of theory and experiment to the case of higher velocity of incident flow.

Lippisch, A.↗
NASA NTRS1976

Pressure distributions on a rectangular aspect-ratio-6, slotted supercritical airfoil wing with externally blown flaps

An investigation was made in the 5.18 m (17 ft) test section of the Langley 300 MPH 7 by 10 foot tunnel on a rectangular, aspect ratio 6 wing which had a slotted supercritical airfoil section and externally blown flaps. The 13 percent thick wing was fitted with two high lift flap systems: single slotted and double slotted. The designations single slotted and double slotted do not include the slot which exists near the trailing edge of the basic slotted supercritical airfoil. Tests were made over an angle of attack range of -6 deg to 20 deg and a thrust-coefficient range up to 1.94 for a free-stream dynamic pressure of 526.7 Pa (11.0 lb/sq ft). The results of the investigation are presented as curves and tabulations of the chordwise pressure distributions at the midsemispan station for the wing and each flap element.

Johnson, W. G., Jr.↗
NASA NTRS1925

Pressure distribution on the nose of an airship in circling flight

In recent tests on the pressures occurring on the envelope and control surfaces of the naval airship C-7, it was noted that the pressures on the nose of the airship, while flying in level circling flight, were symmetrically distributed. Such a condition can only occur when the nose of the airship is pointed directly into the wind, and to accomplish this in circling flight, the axis of the airship must then be parallel to the direction of the motion of the nose. The question was raised as to whether the same conditions occur generally on all airships in circling flight. It appears that airships flying in a constant, level, circling flight path will generally head very closely into the wind, and any deviation will be so slight that the distribution of pressure over the nose will be but slightly, if at all, changed from a symmetrical distribution.

Fairbanks, Karl J↗
NASA NTRS2000

Blade Surface Pressure Distributions in a Rocket Engine Turbine: Experimental Work With On-Blade Pressure Transducers

Understanding the unsteady aspects of turbine rotor flowfields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with surface-mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in three respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, two independent unsteady data acquisition systems and fundamental signal processing approaches were used. Finally, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools will contribute to future turbine programs such as those for reusable launch vehicles.

Hudson, Susan T.↗
NASA NTRS1929

Water-Pressure Distribution on Seaplane Float

The investigation presented in this report was conducted for the purpose of determining the distribution and magnitude of water pressures likely to be experienced on seaplane hulls in service. It consisted of the development and construction of apparatus for recording water pressures lasting one one-hundredth second or longer and of flight tests to determine the water pressures on a UO-1 seaplane float under various conditions of taxiing, taking off, and landing. The apparatus developed was found to operate with satisfactory accuracy and is suitable for flight tests on other seaplanes. The tests on the UO-1 showed that maximum pressures of about 6.5 pounds per square inch occur at the step for the full width of the float bottom. Proceeding forward from the step the maximum pressures decrease in magnitude uniformly toward the bow, and the region of highest pressures narrows toward the keel. Immediately abaft the step the maximum pressures are very small, but increase in magnitude toward the stern and there once reached a value of about 5 pounds per square inch. (author)

Thompson, F L↗
NASA NTRS1923

Pressure Distribution over Thick Airfoils - Model Tests

This investigation was undertaken to study the distribution of loading over thick wings of various types. The unloading on the wing was determined by taking the pressure at a number of holes on both the upper and lower surfaces of a model wing in the wind tunnel. The results from these tests show, first, that the distribution of pressure over a thick wing of uniform section is very little different from that over a thin wing; second, that wings tapering either in chord or thickness have the lateral center of pressure, as would be expected, slightly nearer the center of the wings; and, third, that wings tapering in plan form and with a section everywhere proportional to the center section may be considered to have a loading at any point which is proportional to the chord when compared to a wing with a similar constant section. These tests confirm the belief that wings tapering both in thickness and plan form are of considerable structural value because the lateral center of pressure is thereby moved toward the center of the span.

Norton, F H↗
NASA NTRS1937

Calculated and Measured Pressure Distributions over the Midspan Section of the NACA 4412 Airfoil

Pressures were simultaneously measured in the variable-density tunnel at 54 orifices distributed over the midspan section of 5 by 30 inch rectangular model of the NACA 4412 airfoil at 17 angles of attack ranging from -20 degrees to 30 degrees at a Reynolds number of approximately 3,000,000. Accurate data were thus obtained for studying the deviations of the results of potential-flow theory from measured results. The results of the analysis and a discussion of the experimental technique are presented.

Pinkerton, Robert M↗