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Pressure distributions were measured over rectangular wings of aspect ratio 2 and triangular wings of aspect ratios 2 and 4 at Mach numbers of 2.46 and 3.36. The investigation includes some comparison of the effects of Mach number, Reynolds number, and thickening the wing root sections on the loading.
Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.
Time dependent pressure distribution and threshold acceleration for bubble formation in longitudinally vibrating flexible liquid filled cylinder
Heat transfer and pressure distribution due to two-dimensional repetitive sine wave distortions on flat plate at Mach 20 in helium
Experimental surface pressure distributions on spherically blunted cone at Mach numbers from 1.50 to 4.63
Base pressure distribution on blunted cones during supersonic wind tunnel tests
Static aerodynamic characteristics, pressure distributions, and ram air inflation of attached inflatable drag chute models at supersonic speed
Pressure distribution of spherically-blunted 60 deg half-angle cone in hypersonic flow
Theoretical solution for pressure distribution over arbitrarily shaped periodic waves using Fourier series
Wing pressure distributions obtained in flight with flush orifice and external tubing orifice installations for Mach numbers from 0.50 to 0.97 are compared. The procedure used to install the external tubing orifice is discussed. The results indicate that external tubing orifice installations can give useful results.
This paper presents a preliminary design method for determining a wing's design pressure distribution and geometry based on airfoil normal Mach numbers and airfoil loading. In this method, the perturbation velocities in supercritical regions are computed from airfoil transonic normal Mach numbers and include the influence of local sweep, taper, and three-dimensional induced velocities, so that the appearance and strength of shocks can be expected to resemble those of the airfoil. The velocities in subcritical wing regions are scaled first with simple sweep theory, and then to achieve the desired load distribution. The method was applied to the design of an oblique flying wing, using a linear potential method. The required wing area could be reduced by 14 percent using this method rather than simple sweep theory.
Pressure-distribution tests were made on the 32-foot whirling arm of the Daniel Guggenheim Airship Institute of a tapered wing to determine the rolling and yawing moments due to an angular velocity in yaw. The model was tested at 0 degree and 5 degrees pitch, -1 degree and 5 degree yaw, and with a full-span flap deflected 60 degrees. The results are given in the form of span load distributions and in calculated moment coefficients. The rolling-moment coefficients are in fairly close agreement with those derived by means of a simple approximate theory even for high deflection of the full-span flap.
Pressure-distribution tests were made on the 32-foot whirling arm of the Daniel Guggenheim Airship Institute of a tapered wing to determine the rolling and the yawing moments due to an angular velocity in yaw. The model was tested at 0 degree and 5 degree pitch; 0 degree, 5 degree, and 10 degree yaw; and with split flaps covering 25, 50, 75, and 100 percent of the wing span and deflected 60 degrees. The results are given in the form of load distributions and as calculated moment coefficients. The experimental values of rolling- and yawing- moment coefficients were in fairly close agreement with theory.
Pressure-distribution tests were made on the 32-foot whirling arm of the Daniel Guggenheim Airship Institute of a rectangular wing of NACA 23012 section to determine the rolling and the yawing moment due to angular velocity in yaw. The model was tested at 0 and 5 degree pitch; 0, +/- 5, and +/- 10 degrees yaw; and with no flap and with split flaps 25, 50 and 75 percent of the wing span and deflected 60 degrees. The results are given in the form of span load distributions and as calculated moment coefficients. The experimental values of rolling- and yawing-moment coefficients were in fairly close agreement with theory.
A flight investigation has been made of the surface pressure distribution and the flow field around a dummy, nonrotating, elliptical spinner over a Mach number range from 0.65 to 0.95, which corresponds to a Reynolds number range from about 1.6 x 10(exp 6) per foot to about 3.9 x 10(exp 6) per foot. The results showed that free-stream conditions were approximated from about 15 to 90 percent of the spinner length, but the local Mach number in the propeller plane varied from about 5 percent less than free stream at a Mach number of 0.65 to about 10 percent less than free stream at a mach number of 0.95.
Pressure-distribution measurements were made on the right wing of a pursuit-type airplane at values of Mach number up to 0.80. The results showed that a considerable portion of the lift was carried by components of the airplane other than the wings, and that the proportion of lift carried by the wings may vary considerably with Mach number, thus changing the bending moment at the wing root whether or not there is a shift in the lateral position of the center of pressure. It was also shown that the center of pressure does not necessarily move outward at high Mach numbers, even though the wing-thickness ratio decreases toward the wing tip. The wing pitching-moment coefficient increased sharply in a negative direction at a Mach lift-curve slope increased with Mach number up to values of above the critical value. Pressures inside the wing were small and negative.
Pressure-distribution tests of an NACA 0009 airfoil with a 30-percent-chord plain flap and three plain tabs, having chords 10, 20, and 30 percent of the flap chord, were made. The purpose of these tests was to continue an investigation to supply structural and aerodynamic section data that may be applied to the design of horizontal and vertical tail surfaces. The results are presented as diagrams of resultant pressures and of resultant-pressure increments for the airfoil with the flap and the 20-percent-chord tab. Increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs are also given. At all unstalled flap and tab deflections, the experimental distributions agree well with those calculated by an analytical method. The agreement is poor, however, then the stalled or the unstalled condition of the flap or tab deflected alone was changed to an unstalled or stalled condition by the simultaneous deflection of both the flap and the tab.
Pressure-distribution tests of an NACA 0009 airfoil with an 80-percent-chord plain flap and three plain tabs, having chord of 10, 20, and 30 percent of the flap chord, were made. Section data suitable for application to the design of horizontal and vertical tail surfaces were obtained. Resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab are presented. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. A comparison of some characteristic slopes for the 30-, the 50-, and the 80-percent-chord flaps, tested in the general investigation of plain flaps for control surfaces, is included. Section aerodynamic and load data have been made available for a wide range of flap and a tab chords to be used on an NACA 0009 airfoil or on other conventional sections.