Spanwise Pressure Distributions Research Articles

Pressure Distribution on the Surface of a Yawed Circular Cylinder (The Effect of End-Plates on the Spanwise Pressure Distribution).

Pressure Distribution on the Surface of a Yawed Circular Cylinder (The Effect of End-Plates on the Spanwise Pressure Distribution).

Full-potential analysis of a supersonic delta wing/body

Experimental spanwise pressure distributions for a 60 deg delta wing/body of approximate fineness ratio 7.6 have been obtained and compared to predictions using full-potential theory. Analysis was performed at Mach 1.6 for angles of attack in the range of 0.8-10 deg, and for Mach numbers ranging from 1.4 to 1.8 at lift coefficient 0.3. The intent of the study was to examine an attached flow approach for maneuver wing design in the presence of a fuselage. For the Mach number and angle-of-attack conditions considered, the full-potential theory accurately modeled the pressure distributions, provided the flow remained attached. By combining the full-potential theory results with empirical shock-induced separation criteria, it was found that the onset of shock-induced separation can be predicted.

A three-dimensional inverse calculation of an axial-flow rotor by the finite-pitch actuator duct method.

A fully three-dimensional inverse solution to the axial-flow rotor of a constant hub to tip ratio by means of the finite-pitch actuator duct method for inviscid and incompressible flow is presented. The blade is expressed by a vortex sheet and a source sheet superposed on the mean surface, and the blade shape is determined to satisfy the specified thickness and loading distributions for free-vortex type flow. Two-dimensional numerical calculations are performed by modifying the present method in order to show the difference between the blade shape obtained by the two-dimensional inverse method and that obtained by the three-dimensional inverse method. Three-dimensional numerical calculations show the influences of the stacking position and the blade spanwise pressure distributions. The direct solutions for the designed two-dimensional and three-dimensional rotors are calculated in order to examine the numerical accuracy of the present inverse calculation.

Spanwise pressure distribution on delta wing with leading-edge vortex flap

The aerodynamic characteristics of a highly swept planar delta wing employing conical leading edge flaps are numerically investigated, using a free vortex sheet method that is based on an advanced, three-dimensional inviscid flow panel method employing quadratic doublet distributions to represent the wing surface and the rolled-up vortex sheet and wake. Upward flap deflection shifts the negative pressure peak inboard of the basic wing and develops a significant suction pressure on the flap that then produces thrust component in the direction of flight; overall drag is thereby reduced.

Effects of Tip Endwall Contouring on the Three-Dimensional Flow Field in an Annular Turbine Nozzle Guide Vane: Part 2—Numerical Investigation

This paper describes the numerical investigation of the three dimensional flow through a low speed, low aspect ratio, high turning annular turbine nozzle guide vane with meridional tip endwall contouring. This rotational flow field has been simulated using a finite volume discretization and a time marching technique to solve the three dimensional, time dependent Euler equations expressed in a cylindrical coordinates system. The results are presented under the form of contour plots, spanwise pitch-averaged distributions and blade static pressure distributions. Detailed comparisons with the measurements described in part I of the paper are also provided.Copyright © 1985 by ASME

Induced drag and lift of wing by the piecewise continuous kernel function method

T piecewise continuous kernel function method is extended to compute induced drag on three dimensional wings in subsonic flow The method was used to compute induced drag and lift on a wide number of wings, and the results compared very well with other existing methods Trie method was applied to compare induced drag and lift of nontwisted planar swept forward and back wings with the equivalent M or W type configuration It was found that the spanwise load distribution and the induced drag are in fluenced by the sweep angle of the lifting surface. Fur therrriore, it was found that the spanwise pressure distribution of the M wing type offers some advantages that can be used to improve the configuration performance characteristics as one wishes

Improved numerical method for unsteady lifting surfaces in incompressible flow

y0 6 An improved numerical method has been developed for unsteady thin rectangular wings in incompressible flow. To satisfy the wing boundary conditions, the spanwise pressure distribution is assumed to be stepwise constant, while the chordwise integral is reduced to a finite sum using several kinds of quadratures; Cauchy and logarithmic singularities are treated appropriately. The present method gives excellent results over a wide range of reduced frequency k (k = 0.0-5.0) within reasonable computer time. Extension to compressible flow and to swept tapered wings applications is discussed briefly.

Warping of delta wings for minimum drag

A study was conducted of the possibility to use the results obtained by Bera (1974) in the design of wings of low drag. The case in which the wings are subjected to a given lift is investigated. The results of the study are presented in a number of graphs showing conical twist distributions, basic wing camber surfaces, optimum wing camber surfaces, and sample spanwise pressure distribution of basic and minimum drag shapes

Results of a strong interaction, wake-like model of supersonic separated and reattaching turbulent flows

Results of an interaction theory for supersonic separated andreattaching turbulent boundary layers are presented and compared with recent experiments for flow past a compression ramp. Effects of angle, Mach number, Reynolds number, and upstream pressure gradient are considered for situations where the critical point is located upstream of the trailing edge. When the critical point falls downstream of the trailing edge the whole region of separated flow is influenced by length. In these short ramp flows the peak pressure attains a maximum at a critical angle and then decreases with increasing angle. It is shown that this effect is responsible for the spanwise pressure distributions measured by Whitehead and Keyes for flow over a delta wing with a trailing edge flap. Results are also presented for a turbulent boundary layer-shock wave interaction.

The Effects of Blunt Leading Edges on Delta Wings at Mach 5.8

Pressure distributions were measured on a series of four delta wings with subsonic and supersonic leading edges, both sharp and blunt. The blunt leading edge radius was about 0.5 per cent of root chord. Schlieren studies were also made to determine top and side view shock locations. The tests were conducted at a nominal Mach number of 5.8, and at Reynolds numbers between 0.335 x 10^6 and 0.901 x 10^6 based on root chord. Angular settings covered a range -0.2 ≤ w/V ≤ 0.5 in pitch at zero yaw {about -11.5° ≤ a ≤ + 30°), and a range of v/V = ± 0.125 (about ± 7.2°) at a fixed angle of pitch of 11.5°. The effects of bluntness were found to be small. Also, the pressures produced by shock wave interactions with the boundary layer, and the inviscid pressures generated by the blunt leading edges, were found to be small compared with the inviscid pressures producing lift on the basic wing. Spanwise pressure distributions show no similarity to those obtained by linearized theory. Centerline lower surface pressure in pitch at zero yaw is bracketed between the Newtonian value ΔP/q = 2(w/V)^2 and the two-dimensional exact value.