Pressure distribution in a hydrostatic bearing of multi-wells
Pressure distribution in hydrostatic bearing of multi-wells obtained by use of Navier-Stokes equations
PRESSURE DISTRIBUTION↗
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Force- and Pressure-Distribution Measurements on Eight Fuselages
The present report deals with force- and pressure-distribution measurements on a number of fuselage forms of varying slenderness ratio, varying rearward position of maximum thickness, and varying nose ratio. The effect of these parameters on the force and moment coefficients was determined. The linearity of the difference between the theoretical and experimental fuselage moments with the friction lift made it possible to indicate a neutral point and its travel with the different parameters. The pressure-distribution measurements yielded absolute values for the increase of velocity. A comparison with the theory indicated good agreement at small angles of attack, but considerable differences at greater angles of attack, where potential flow could no longer be assumed.
Lange, G.↗
Measurements of Pressure Distributions and Force Coefficients in a Squeeze Film Damper. Part 1: Fully Open Ended Configuration
Measurements of pressure distributions and force coefficients were carried out in two types of squeeze film dampers, executing a circular centered orbit, an open-ended configuration, and a partially sealed one, in order to investigate the effect of fluid inertia and cavitation on pressure distributions and force coefficients. Dynamic pressure measurements were carried out for two orbit radii, epsilon 0.5 and 0.8. It was found that the partially sealed configuration was less influenced by fluid inertia than the open ended configuration.
Jung, S. Y.↗
Pressure distributions and oil-flow patterns for a swept circulation-control wing
Pressure distributions and photographs of oil flow patterns are presented for a circulation control wing. The model was an aspect ratio four semispan wing mounted on the side wall of the NASA Ames Transonic Wind Tunnel. The airfoil was a 20 percent thick ellipse, modified with circular leading and trailing edges of 4 percent radius, and had a 25.4 cm constant chord. This configuration does not represent a specific wing design, but is generic. A full span, tangetial, rearward blowing, circulation control slot was incorporated ahead of the trailing edge on the upper surface. The wing was tested at Mach numbers from 0.3 to 0.75 at sweep angle of 0 to 45 deg with internal to external pressure ratios of 1.0 to 3.0. Lift and pitching momemt coefficients were obtained from measured pressure distributions at five span stations. When the conventional corrections resulting from sweep angle are applied to the lift and moment of circulation control sections, no additional corrections are necessary to account for changes in blowing efficiency. This is demonstrated for an aft sweep angle of 45 deg. An empirical technique for estimating the downwash distribution of a swept wing was validated.
Keener, Earl R.↗
Pressure distribution for the wing of the YAV-8B airplane; with and without pylons
Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20 and 35 percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60 percent chord) whether or not the pylons are installed for Mach numbers greater than or equal to 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag at transonic Mach number increases.
Saltzman, Edwin J.↗
Pressure distribution over a Clark Y-H airfoil section with a split flap
An investigation was made in the N.A.C.A. 7 by 10 foot wind tunnel of the distribution of air pressure over one chord section of a Clark Y-H airfoil with a split flap. The results obtained are given in the form of diagrams showing the chordwise distribution of pressure on both airfoil and flap as plots of calculated aerodynamic coefficients for both airfoil and flap.
Wenzinger, Carl J↗
Pressure distributions on a 0.02-scale Space Shuttle orbiter nose at Mach 21.5 in helium
Pressure distributions on a 0.02-scale model of the Space Shuttle orbiter forward fuselage were obtained in the 22-inch aerodynamic leg of the Langley Hypersonic Helium Tunnel Facility at a nominal free-stream Mach number of 21.5 and a ratio of specific heats of 1.67 for inclusion in the database of the Shuttle entry air data system (SEADS). The data were measured at model angles of attack of 0 deg to 50 deg in 5 deg increments for zero sideslip angle and at model sideslip angles of -5 deg to 5 deg for angles of attack equal to 5, 20, 35, and 40 deg. These data displayed trends similar to those observed in other wind tunnels at Mach 6 and 10 in air. Specifically noted is a shift in the location of the stagnation point at angles of attack above 15 deg; this effect did not, however, occur in flight. By comparison, the data obtained at Mach 6 in the Langley Hypersonic CF4 Tunnel, corresponding to a lower ratio of specific heats in the postshock region than those in helium and air, showed some reduction of the stagnation point shift at the higher angles of attack. The differences between flight and wind tunnel pressure distributions are believed due primarily to high-temperature gas chemistry effects in flight, which include lower effective specific heat ratios but which were not completely duplicated in the wind tunnels.
Ashby, George C., Jr.↗
Pressure distributions and shock shapes for a bent-nose biconic at incidence
Initial results of a NASA study of the lift-drag characteristics of a 12.84/7 deg biconic model intended for airbraking during atmospheric entry of probes to Mars, Venus, Saturn, and Titan are reported. Pressure distributions and shock shapes were measured in the Langley 20 in. Mach 6 tunnel with the spherically blunted bent-nose model set at angles from 0-25 deg. Pressure distributions and shock shapes where computed using the STEIN flowfield code, which features a MacCormack scheme to integrate the three-dimensional Euler equations, the Rankine-Hugoniot jump conditions to model shock waves as discontinuities, and requires a supersonic condition at every step. A comparison was made between measured and predicted values. The leeward shock angle was found to be predictable to within 3% for all angles of attack, while parabolized Navier-Stokes equations are regarded as offering more accurate results than the STEIN code for surface pressure distributions.
Miller, C. G., III↗
Variable Sweep Transition Flight Experiment (VSTFE)-Parametric Pressure Distribution Boundary Layer Stability Study and Wing Glove Design Task
The Variable Sweep Transition Flight Experiment (VSTFE) was initiated to establish a boundary-layer transition data base for laminar flow wing design. For this experiment, full-span upper-surface gloves will be fitted to a variable sweep F-14 aircraft. The results of two initial tasks are documented: a parametric pressure distribution/boundary-layer stability study and the design of an upper-surface glove for Mach 0.8. The first task was conducted to provide a data base from which wing-glove pressure distributions could be selected for glove designs. Boundary-layer stability analyses were conducted on a set of pressure distributions for various wing sweep angles, Mach numbers, and Reynolds number in the range of those anticipated for the flight-test program. The design procedure for the Mach 0.8 glove is described, and boundary-layer stability calculations and pressure distributions are presented both at design and off-design conditions. Also included is the analysis of the clean-up glove (smoothed basic wing) that will be flight-tested initially and the analysis of a Mach 0.7 glove designed at the NASA Langley Research Center.
Rozendaal, Rodger A.↗
Subsonic pressure distributions near a wing-fin juncture on a supersonic arrow-wing configuration with wing-mounted vertical fins
A generic supersonic arrow-wing configuration has been extensively studied to develop a consistent experimental data base to aid in verifying analytical prediction methods. One of the wing configurations studied was a twisted and cambered wing with wing-mounted vertical fins and trailing-edge flaps. The pressure data obtained on that configuration was limited in the wing-fin juncture region, and no pressures were obtained on the fin itself. However, detailed data in the juncture region and on the fin are essential to check out the analytical prediction methods. Therefore, the model was modified to provide details of the juncture pressure field as well as pressures on the fin. The results of this experimental investigation showed that the addition of a fin helped improve the wing pressure distribution outboard of the fin. Deflection of the trailing-edge flap did not adversely affect this trend. The pressure distributions on the vertical fin were generally affected more by change in angle of attack than by flap deflection. The spanwise pressure distribution on the fin itself indicated a load reversal with angle of attack.
Huffman, J. K.↗
Ring baffle pressure distribution and slosh damping in large cylindrical tanks
An investigation was conducted to determine the pressure loads and damping associated with rigid ring baffles in relatively large cylindrical tanks. The radial and circumferential pressure distribution, as well as the damping, was measured on a ring baffle subjected to fundamental antisymmetric slosh in a 284-cm-diameter rigid tank. Experimental and analytical data are presented as a function of slosh velocity or amplitude, baffle spacing, and baffle locations both above and below the liquid surface. Results suggest that pressure distributions and damping values can be determined from available theories for the design of single and multiple baffle configurations.
Scholl, H. F.↗
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.↗
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.↗
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↗
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.↗
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.↗
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.↗
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↗