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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.
Measurements of the distribution of surface pressures were made on two wing-body combinations employing a triangular wing having an aspect ratio ratio of 2. The first of these combinations had a wing with the NACA 0008-63 section parallel to the air stream while the wing of the second had the NACA 0005-63 section in the stream direction. The measurements were made for Mach numbers from 0.11 to 0.95 at a constant Reynolds number of 3.0 million. Data were also obtained for Reynolds numbers up to 15.0 million at the lower Mach number. A summary of data obtained from previously published force tests of these same models is also included for completeness of the presentation of the effect of wing thickness.
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.
Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane is at the divergence end of the constant area throat section. At high stagnation pressures the fluid stays liquid well into the constant area throat region; at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region, and this anomaly is related to the two-phase process. The fluid is metastably all liquid below the saturation pressure.
Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane appears to be at the divergence end of the constant area throat section. At high stagnation pressures the fluid appears to stay liquid well into the constant area throat region; however, at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region. This anomaly appears to be fundamentally related to the two-phase process and not merely to the present specific nozzle geometry. The fluid appears to be metastably all liquid below the saturation pressure. The data are compared to various flow models. No model adequately describes the whole range of the experiment.
This note presents the results obtained in pressure distribution tests on the right upper wing panel and tip of a Douglas M-3 airplane in flight. These tests are a part of a more extensive investigation of the effect of changes in tip shape on the load distribution, the tip reported herein being the first of a series of tip shapes being tested. The results are given in tables and curves in such form that the load distribution for any conditions may be determined easily. Tests were made at Langley Field by the NACA in the spring of 1930.
This note presents the results obtained in pressure distribution tests on the right upper panel of a Douglas M-3 airplane in flight. These tests are a part of an extensive investigation on the effect of changes in tip shape on the load distribution.
This note presents the results of flight pressure distribution tests on the right upper wing panel of a Douglas M-3 airplane equipped with a semicircular wing tip. The results are given in tables and curves in such form that the load distribution for any force coefficient within the usual range encountered in flight may easily be determined.
An investigation of the pressure distribution on the fuselage nose and the pilot canopy of a supersonic airplane model has been conducted at a Mach number of 1.90 over a wide range of angles of attack and yaw. Boundary layer separation apparently occurred from the upper surface at angles of attack above 24 degrees and from the lower surface at minus 15 degrees. No separation from the sides of the fuselage was evident at yaw angles up to 12 degrees.
A previously developed eigenvalue model is extended to determine polishing pressure distribution by rectangular tools with unequal stiffness in two directions on cylindrical optics. Tool misfit is divided into two simplified one-dimensional problems and one simplified two-dimensional problem. Tools with nonuniform cross-sections are treated with a new one-dimensional eigenvalue algorithm, permitting evaluation of tool designs where the edge is more flexible than the interior. This maintains edge pressure variations within acceptable parameters. Finite element modeling is employed to resolve upper bounds, which handle pressure changes in the two-dimensional misfit element. Paraboloids and hyperboloids from the NASA AXAF system are treated with the AXAFPOD software for this method, and are verified with NASTRAN finite element analyses. The maximum deviation from the one-dimensional azimuthal pressure variation is predicted to be 10 percent and 20 percent for paraboloids and hyperboloids, respectively.