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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1960

Effect of Reynolds Number on the Force and Pressure Distribution Characteristics of a Two-Dimensional Lifting Circular Cylinder

A two-dimensional lifting circular cylinder has been tested over a Mach number range from 0.011 to 0.32 and a Reynolds number range from 135,000 to 1,580,000 to determine the force and pressure distribution characteristics. Two flaps having chords of 0.37 and 6 percent of the cylinder diameter, respectively, and attached normal to the surface were used to generate lift. A third configuration which had 6-percent flaps 1800 apart was also investigated. All flaps were tested through a range of angular positions. The investigation also included tests of a plain cylinder without flaps. The lift coefficient showed a wide variation with Reynolds number for the 6-percent flap mounted on the bottom surface at the 50-percent-diameter station, varying from a low of about 0.2 at a Reynolds number of 165,000 to a high of 1.54 at a Reynolds number of 350,000 and then decreasing almost linearly to a value of 1.0 at a Reynolds number of 1,580,000. The pressure distribution showed that the loss of lift with Reynolds number above the critical was the result of the separation point moving forward on the upper surface. Pressure distributions on a plain cylinder also showed similar trends with respect to the separation point. The variation of drag coefficient with Reynolds number was in direct contrast to the lift coefficient with the minimum drag coefficient of 0.6 occurring at a Reynolds number of 360,000. At this point the lift-drag ratios were a maximum at a value of 2.54. Tests of a flap with a chord of 0.0037 diameter gave a lift coefficient of 0.85 at a Reynolds number of 520,000 with the same lift-drag ratio as the larger flap but the position of the flap for maximum lift was considerably farther forward than on the larger flap. Tests of two 6-percent flaps spaced 180 deg apart showed a change in the sign of the lift developed for angular positions of the flap greater than 132 deg at subcriti- cal Reynolds numbers. These results may find use in application to air- craft using forebody strakes. The drag coefficient developed by the flaps when normal to the relative airstream was approximately equal to that developed by a flat plate in a similar attitude.

Lockwood, Vernard E.↗
NASA NTRS1995

Airfoil modification effects on subsonic and transonic pressure distributions and performance for the EA-6B airplane

Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

Allison, Dennis O.↗
NASA NTRS1930

Pressure distribution on the tail surfaces of a PW-9 pursuit airplane in flight

Presented here are pressure distribution data obtained from the tail surfaces of a PW-9 in a number of flight maneuvers. The results given are part of those obtained in an extensive investigation of the pressure distribution over all of the lifting and control surfaces of this airplane. The results are given in tabular and curve form and are discussed briefly with respect to their comparison with existing tail surface design specifications. It is recommended that tail load design loadings should be revised upwards. This is particularly true of leading edge loads, which should be at least doubled for thick sections.

Rhode, Richard V↗
NASA NTRS1949

On the Installation of Jet Engine Nacelles on a Wing Fourth Partial Report: Pressure-Distribution Measurements on a Sweptback Wing with Jet Engine Nacelle

The present report, which deals with pressure-distribution measurements made on a sweptback wing with a jet engine nacelle, is similar to a report on pressure-distribution measurements on a rectangular wing with a jet engine nacelle (second partial report). Here, in investigations preliminary to high-speed measurements, as in the second partial report, useful arrangements and fillet designs have been discovered.

Buschner, R.↗
NASA NTRS1988

Pressure distributions from subsonic tests of an advanced laminar-flow-control wing with leading- and trailing-edge flaps

An unswept, semispan wing model equipped with full-span leading- and trailing-edge flaps was tested in the Langley 14- by 22-Foot Subsonic Tunnel to determine the effect of high-lift components on the aerodynamics of an advanced laminar-flow-control (LFC) airfoil section. Chordwise pressure distributions near the midsemispan were measured for four configurations: cruise, trailing-edge flap only, and trailing-edge flap with a leading-edge Krueger flap of either 0.10 or 0.12 chord. Part 1 of this report (under separate cover) presents a representative sample of the plotted pressure distribution data for each configuration tested. Part 2 presents the entire set of plotted and tabulated pressure distribution data. The data are presented without analysis.

Applin, Zachary T.↗
NASA NTRS1988

Pressure distributions from subsonic tests of an advanced laminar-flow-control wing with leading- and trailing-edge flaps

An unswept, semispan wing model equipped with full-span leading- and trailing-edge flaps was tested in the Langley 14- by 22-Foot Subsonic Tunnel to determine the effect of high-lift components on the aerodynamics of an advanced laminar-flow-control (LFC) airfoil section. Chordwise pressure distributions near the midsemispan were measured for four configurations: cruise, trailing-edge flap only, and trailing-edge flap with a leading edge Kreuger flap of either 0.10 or 0.12 chord. Part 1 of this report presents a representative sample of the plotted pressure distribution data for each configuration tested. Part 2 (under separate cover) presents the entire set of plotted and tabulated pressure distribution data. The data are presented without analysis.

Applin, Zachary T.↗
NASA NTRS1926

Pressure Distribution on the C-7 Airship

This investigation was made for the purpose of determining the aerodynamic pressure distribution encountered on a "C" class airship in flight. It was conducted in two parts: (a) tests on the tail surfaces in which the pressures at 201 points were measured and (b) tests on the envelope in which 190 points were used, both tests being made under as nearly identical flight conditions as possible, so that the results could be combined and the pressure distribution over the entire airship obtained.

Crowley, J W , Jr↗
NASA NTRS1977

Pressure distributions on a 1- by 3-meter semispan wing with a nonstreamwise tip in subsonic flow

Pressure distributions on a 1- by 3-meter semispan wing with a tip that is streamwise at 0 deg sweep are presented. The tip becomes nonstreamwise as the wing is swept. At 0 deg sweep angle, the semispan wing has a taper ratio of 1.0 and NACA 0012 airfoil section contours. The test was conducted in the Langley V/STOL tunnel at a freestream dynamic pressure of 2.46 kPa. Pressure distributions are presented for sweep angles of 10, 20, 30, and 40 with an angle of attack range from -6 deg to 20 deg. The data are presented without analysis or discussion.

Yip, L. P.↗
NASA NTRS1975

Results of an investigation of the 0.003-scale space shuttle external tank MSFC model 460 in the NASA/MSFC 14 x 14 inch Trisonic Wind Tunnel to determine static pressure distributions during reentry (TA2F), volume 3

Static pressure distributions for the external tank (ET) at reentry conditions are presented. Basic configuration of the model was the MCR 0200 ET modified to include a rectangular crossbar at the aft ET/orbiter attach point. Mach numbers were 1.96, 3.48, and 4.96. Reynolds number per foot at these Mach numbers were 6.95 million, 6.42 million, and 4.95 million, respectively. Angle of attack range was -8 to 100 degrees and roll angle was 0 to 315 degrees.

Ramsey, P. E.↗
NASA NTRS1980

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. I - Pressure distribution

Asymptotic solutions are derived for the pressure distribution in the interaction of a weak normal shock wave with a turbulent boundary layer. The undisturbed boundary layer is characterized by the law of the wall and the law of the wake for compressible flow. In the limiting case considered, for 'high' transonic speeds, the sonic line is very close to the wall. Comparisons with experiment are shown, with corrections included for the effect of longitudinal wall curvature and for the boundary-layer displacement effect in a circular pipe.

Messiter, A. F.↗
NASA NTRS1980

Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. Part 1: Pressure distribution. Part 2: Wall shear stress. Part 3: Simplified formulas for the prediction of surface pressures and skin friction

An asymptotic description is derived for the interaction between a shock wave and a turbulent boundary layer in transonic flow, for a particular limiting case. The dimensionless difference between the external flow velocity and critical sound speed is taken to be much smaller than one, but large in comparison with the dimensionless friction velocity. The basic results are derived for a flat plate, and corrections for longitudinal wall curvature and for flow in a circular pipe are also shown. Solutions are given for the wall pressure distribution and the shape of the shock wave. Solutions for the wall shear stress are obtained, and a criterion for incipient separation is derived. Simplified solutions for both the wall pressure and skin friction distributions in the interaction region are given. These results are presented in a form suitable for use in computer programs.

Adamson, T. C., Jr.↗
NASA NTRS1935

Full-scale Force and Pressure-distribution Tests on a Tapered U.S.A. 45 Airfoil

This report presents the results of force and pressure-distribution tests on a 2:1 tapered USA 45 airfoil as determined in the full-scale wind tunnel. The airfoil has a constant-chord center section and rounded tips and is tapered in thickness from 18 percent at the root to 9 percent at the tip. Force tests were made throughout a Reynolds Number range of approximately 2,000,000 to 8,000,000 providing data on the scale effect in addition to the conventional characteristics. Pressure-distribution data were obtained from tests at a Reynolds Number of approximately 4,000,000. The aerodynamic characteristics given by the usual dimensionless coefficients are presented graphically.

Parsons, John F↗
NASA NTRS1978

Pressure Distributions for the GA(W)-2 Airfoil with 20% Aileron, 25% Slotted Flap and 30% Fowler Flap

Surface pressure distributions were measured for the 13% thick GA(W)-2 airfoil section fitted with 20% aileron, 25% slotted flap and 30% Fowler flap. All tests were conducted at a Reynolds number of 2.2 x 10 to the 6th power and a Mach number of 0.13. Pressure distribution and force and moment coefficient measurements are compared with theoretical results for a number of cases. Agreement between theory and experiment is generally good for low angles of attack and small flap deflections. For high angles and large flap deflections where regions of separation are present, the theory is inadequate. Theoretical drag predictions are poor for all flap-extended cases.

Wentz, W. H., Jr.↗
NASA NTRS1977

Pressure distributions induced by elevon deflections on swept wings and adjacent end-plate surfaces at Mach 6

Surface pressure distributions are presented for regions where three-dimensional separated flow effects are prominent on swept-wing-elevon-end-plate models of 0 degree, 50 degree, and 70 degree sweepback, and with 0 degree, 10 degree, 20 degree, and 30 degree elevon deflections. Surface-oil-flow photographs and pressure distributions on the flat-plate wing, elevon, and end-plate surfaces are presented for numerous geometric variations, including various spacings between the elevon and the end plate, with and without a tip fin. The data, for a free-stream Mach number of 6 and a wing-root-chord Reynolds number of 20 x 10 to the sixth power, reveal considerably larger regions of elevon induced loads on the adjacent end-plate surface than would be anticipated by using inviscid flow analyses.

Kaufman, L. G., II↗
NASA NTRS1973

Heat transfer and pressure distributions on hemisphere-cylinders in methane-air combustion products at Mach 7

Heat-transfer and pressure distributions were measured over the surfaces of three hemisphere-cylinder models tested at a nominal Mach number of 7 in the Langley 8-foot high-temperature structures tunnel which uses methane-air products of combustion as a test medium. The results showed that the heat-transfer and pressure distributions over the surface of the models were in good agreement with experimental data obtained in air and also with theoretical predictions.

Weinstein, I.↗
NASA NTRS1944

Effect of Compressibility on Pressure Distribution over an Airfoil with a Slotted Frise Aileron

Pressure distribution measurements were made over an airfoil with slotted Frise aileron up to 0.76 Mach at various angles of attack and aileron defections. Section characteristics were determined from these pressure data. Results indicated loss of aileron rolling power for deflections ranging from -12 Degrees to -19 Degrees. High stick forces for non-differential deflections incurred at high speed, which were due to overbalancing tendency of up-moving aileron, may precipitate serious control difficulties. Detailed results are presented graphically.

WINGS-REPUBLIC P-47B-3 WIND TUNNEL TESTS(NACA-HIGH↗