Search NASA⌕ Search

SEARCH · Search NASA

Results for “WINGS, RECTANGULAR”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Finite-volume solution of the Euler equations for steady three-dimensional transonic flow

A numerical finite-volume method is proposed for solving the Euler equations of a steady three-dimensional flow. The subsonic regions embedded in steady supersonic flow about aircraft are computed by introducing a time-dependent coordinate system that maps the fuselage and wing into one coordinate surface. The suitability of the method to a delta-winged aircraft is demonstrated, and its applicability to rectangular-winged aircraft and problems in internal transonic flow is pointed out.

Rizzi, A.↗
NASA NTRS1950

Line-vortex theory for calculation of supersonic downwash

The perturbation field induced by a line vortex in a supersonic stream and the downwash behind a supersonic lifting surface are examined to establish approximate methods for determining the downwash behind supersonic wings. Lifting-lines methods are presented for calculating supersonic downwash. A bent lifting-line method is proposed for computing the downwash field behind swept wings. When applied to triangular wings with subsonic leading edges, this method gives results that, in general, are in good agreement with the exact linearized solution. An unbent lifting-line method (horseshoe-vortex system) is proposed for unswept wings. This method is applied to determine downwash behind rectangular wings with aspect ratios of 2 and 4. Excellent agreement with exact linearized theory is obtained for both aspect ratios by placing the lifting line at the 1/2-chord point. The use of lifting-lines therefore appears promising for obtaining estimates of the downwash behind supersonic wings.

Mirels, Harold↗
NASA NTRS1992

Computation of vortex wake flows and control of their effects on trailing wings

The near-vortex-wake flow of a large aspect-ratio rectangular wing is accurately computed by using the thin-layer and full Navier-Stokes (NS) equations. The chordwise section of the wing is a NACA-0012 airfoil and its tip is round. The computations have been carried out on a fine C-O grid using an implicit, upwind, flux-difference splitting, finite-volume scheme. The thin-layer NS results have been obtained with and without flux limiters, and the full NS results have been obtained without flux limiters. Flow transition from laminar to turbulent is mimicked by turning-on the Baldwin-Lomax algebraic model at an experimentally prescribed chord-station location of 0.05. Comparison of computed results and experimental data shows that the full NS results give the best resolution of the near-vortex-wake flow. Next, the strength of the wing-tip vortex has been reduced substantially without reducing the lift coefficient by using flow-injection from a slot along a portion of the wing tip. The flow injection is directed in the wing plane at 45 deg with the wing-tip chord.

Wong, Tin-Chee↗
NASA NTRS1997

Study of the Unsteady Flow Features on a Stalled Wing

The occurrence of large scale structures in the post stall flow over a rectangular wing at high angles of attack was investigated in a small-scale subsonic wind tunnel. Mean and time dependent measurements within the separated flow field suggest the existence of two distinct angle of attack regimes beyond wing stall. The shallow stall regime occurs over a narrow range of incidence angles (2-3 deg.) immediately following the inception of leading edge separation. In this regime, the principal mean flow structures, termed stall cells, are manifested as a distinct spanwise periodicity in the chordwise extent of the separated region on the model surface with possible lateral mobility not previously reported. Within the stall cells and on the wing surface, large amplitude pressure fluctuations occur with a frequency much lower than anticipated for bluff body shedding, and with minimum effect in the far wake. In the deep stall regime, stall cells are not observed and the separated region near the model is relatively free of large amplitude pressure disturbances.

Yon, Steven A.↗
NASA NTRS1934

Effect of retractable-spoiler location on rolling- and yawing-moment coefficients

In this report are presented the results of wind-tunnel tests of retractable spoilers on the upper surface of a Clark Y wing, which have been made as part of an investigation of lateral control devices being conducted by the National Advisory Committee for Aeronautics. Spoilers with chords up to 15.0 percent of the wing chord were tested in several locations on a plain rectangular wing and in two locations on the same wing equipped with a 20.0 percent chord split flap down 60 degrees. Charts are given for four representative angles of attack from which values of rolling- and yawing-moment coefficients may be obtained for spoilers up to 15.0 percent chord located on the upper surface of a Clark Y wing. The tests showed that at low angles of attack practically the same rolling moments can be obtained with a given spoiler at any location back of 30.0 percent of the wing chord, while at high angles of attack there is a definite advantage in locating the spoiler at least as far forward as 30.0 percent of the chord. The yawing moments accompanying a given rolling moment increase positively as the spoiler location is moved forward from the trailing edge of the wing. It is concluded that the 30.0 percent chord location is probably the optimum provided that instantaneous response of the airplane to a control movement can be obtained.

Shortal, J A↗
NASA NTRS1936

Investigations on the amount of downwash behind rectangular and elliptical wings

The downwash behind a wing may be computed from the system of vortices created by the wing and which can be substituted for the latter. If the actual rolling-up process of the free vortex sheet is disregarded, the establishment of boundaries for the downwash factors is contingent upon the premise that no rolling up takes place or else that the developed tip vortices start direct on the effective line. The actual downwash factors for the rectangular wing are rather of the approximate magnitude of the values computable with the developed tip vortices starting direct on the wing.

Muttray, H↗
NASA NTRS1973

Wing tip vortex measurements with laser Doppler systems

The vortex velocity field produced by a rectangular wing in a subsonic wind tunnel was measured using two laser Doppler velocimeter systems. One system made three dimensional mean velocity measurements and the other made one dimensional turbulence measurements. The systems and test procedures are described and comparisons of the measurements are made. The data defined a strong spiral motion in the vortex formation process.

Fuller, C. E., III↗
NASA NTRS1934

The effects of full-span and partial-span split flaps on the aerodynamic characteristics of a tapered wing

The investigation was made to determine the effects of full-span and of partial-span split flaps on the aerodynamic characteristics of a tapered wing. Aerodynamic force tests were made in the N.A.C.A. 7 by 10 foot wind tunnel on a highly tapered Clark Y wing equipped with various split flaps. Two sizes of tapered-chord flaps were tested as full-span flaps, and a narrow tapered-chord flap was tested as a partial-span flap by cutting off portions first from the tip and then from the center. The investigation showed that with full-span split flaps the lift and drag characteristics of the tapered wing up to the stall are similar to those of a rectangular wing with flaps of comparable size, but that the stall of the tapered wing with full-span flaps occurs at progressively lower angles of attack with increasing flap deflection up to that for maximum lift. For partial-span tapered split flaps on a tapered wing it was found that the maximum lift is greater, and the lift-drag ratio at maximum lift is less, when the partial-span flap is located at the center of the wing than when it is located at the tip portion.

Wenzinger, Carl J↗
NASA NTRS1992

Unsteady transition measurements on a pitching three-dimensional wing

Boundary layer transition measurements were made during an experimental study of the aerodynamics of a rectangular wing undergoing unsteady pitching motions. The wing was tested at chordwise Mach numbers between 0.2 and 0.6, at sweep angles of 0, 15, and 30 deg, and for steady state, sinusoidal, and constant pitch rate motions. The model was scaled to represent a full size helicopter rotor blade, with chord Reynolds numbers between 2 and 6 x 10(exp 6). Sixteen surface hot-film gages were located along three spanwise stations: 0.08, 0.27, and 0.70 chords from the wing tip. Qualitative heat transfer information was obtained to identify the unsteady motion of the point of transition to turbulence. In combination with simultaneous measurements of the unsteady surface pressure distributions, the results illustrate the effects of compressibility, sweep, pitch rate, and proximity to the wing tip on the transition and relaminarization locations.

Lorber, Peter F.↗
NASA NTRS1978

The practical application of a finite difference method for analyzing transonic flow over oscillating airfoils and wings

Separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady equations for small disturbances was performed. The steady velocity potential was obtained first from the well known nonlinear equation for steady transonic flow. The unsteady velocity potential was then obtained from a linear differential equation in complex form with spatially varying coefficients. Since sinusoidal motion is assumed, the unsteady equation is independent of time. The results of an investigation into the relaxation-solution-instability problem was discussed. Concepts examined include variations in outer boundary conditions, a coordinate transformation so that the boundary condition at infinity may be applied to the outer boundaries of the finite difference region, and overlapping subregions. The general conclusion was that only a full direct solution in which all unknowns are obtained at the same time will avoid the solution instabilities of relaxation. An analysis of the one-dimensional form of the unsteady transonic equation was studied to evaluate errors between exact and finite difference solutions. Pressure distributions were presented for a low-aspect-ratio clipped delta wing at Mach number of 0.9 and for a moderate-aspect-ratio rectangular wing at a Mach number of 0.875.

Weatherill, W. H.↗
NASA NTRS1988

Effect of large amplitude pitching motions on the unsteady aerodynamic characteristics of flat-plate wings

Large-amplitude unsteady motion effects on the aerodynamic force and stability characteristics of flat-plate wings were investigated in a wind tunnel for the cases of two delta wings with respective leading edge sweeps of 70 and 45 deg, and a rectangular wing whose aspect ratio was equal to that of the 70-deg delta wing. Attention was given to the effects of reduced frequency and mean angle of attack. It is found that lags in vortex burst location and separation/reattachment of flow on the upper surface of the wing produced large overshoots and hysteresis loops in normal force and pitching moment coefficients that were a strong function of mean oscillation angle and reduced frequency.

Brandon, Jay M.↗
NASA NTRS2004

Agglomeration Multigrid for an Unstructured-Grid Flow Solver

An agglomeration multigrid scheme has been implemented into the sequential version of the NASA code USM3Dns, tetrahedral cell-centered finite volume Euler/Navier-Stokes flow solver. Efficiency and robustness of the multigrid-enhanced flow solver have been assessed for three configurations assuming an inviscid flow and one configuration assuming a viscous fully turbulent flow. The inviscid studies include a transonic flow over the ONERA M6 wing and a generic business jet with flow-through nacelles and a low subsonic flow over a high-lift trapezoidal wing. The viscous case includes a fully turbulent flow over the RAE 2822 rectangular wing. The multigrid solutions converged with 12%-33% of the Central Processing Unit (CPU) time required by the solutions obtained without multigrid. For all of the inviscid cases, multigrid in conjunction with an explicit time-stepping scheme performed the best with regard to the run time memory and CPU time requirements. However, for the viscous case multigrid had to be used with an implicit backward Euler time-stepping scheme that increased the run time memory requirement by 22% as compared to the run made without multigrid.

Frink, Neal↗
NASA NTRS1974

Extension of leading-edge-suction analogy to wings with separated flow around the side edges at subsonic speeds

A method for determining the lift, drag, and pitching moment for wings which have separated flow at the leading and side edges with subsequently reattached flow downstream and inboard is presented. Limiting values of the contribution to lift of the side-edge reattached flow are determined for rectangular wings. The general behavior of this contribution is computed for rectangular, cropped-delta, cropped-diamond, and cropped-arrow wings. Comparisons of the results of the method and experiment indicate reasonably good correlation of the lift, drag, and pitching moment for a wide planform range. The agreement of the method with experiment was as good as, or better than, that obtained by other methods. The procedure is computerized and is available from COSMIC as NASA Langley computer program A0313.

Lamar, J. E.↗
NASA NTRS1991

The benchmark aeroelastic models program: Description and highlights of initial results

An experimental effort was implemented in aeroelasticity called the Benchmark Models Program. The primary purpose of this program is to provide the necessary data to evaluate computational fluid dynamic codes for aeroelastic analysis. It also focuses on increasing the understanding of the physics of unsteady flows and providing data for empirical design. An overview is given of this program and some results obtained in the initial tests are highlighted. The tests that were completed include measurement of unsteady pressures during flutter of rigid wing with a NACA 0012 airfoil section and dynamic response measurements of a flexible rectangular wing with a thick circular arc airfoil undergoing shock boundary layer oscillations.

Bennett, Robert M.↗
NASA NTRS1992

The benchmark aeroelastic models program: Description and highlights of initial results

An experimental effort was implemented in aeroelasticity called the Benchmark Models Program. The primary purpose of this program is to provide the necessary data to evaluate computational fluid dynamic codes for aeroelastic analysis. It also focuses on increasing the understanding of the physics of unsteady flows and providing data for empirical design. An overview is given of this program and some results obtained in the initial tests are highlighted. The tests that were completed include measurement of unsteady pressures during flutter of a rigid wing with an NACA 0012 airfoil section and dynamic response measurements of a flexible rectangular wing with a thick circular arc airfoil undergoing shock boundary layer oscillations.

Bennett, Robert M.↗
NASA NTRS1946

Wind-Tunnel Investigation of the Horizontal Motion of a Wing Near the Ground

By the method of images the horizontal steady motion of a wing at small heights above the ground was investigated in the wind tunnel, A rectangular wing with Clark Y-H profile was tested with and without flaps. The distance from the trailing edge of the wing to the ground was varied within the limits 0.75 less than or = s/c less than or = 0.25. Measurements were made of the lift, the drag, the pitching moment, and the pressure distribution at one section. For a wing without flaps and one with flaps a considereble decrease in the lift force and a,drop in the drag was obtained at angles of attack below stalling. The flow separation near the ground occurs at smaller angles of attack than is the case for a great height above the ground. At horizontal steady flight for practical values of the height above the ground the maximum lift coefficient for the wing without flaps changes little, but markedly decreases for the wing with flaps. Analysis of these phenomena involves the investigation of the pressure distribution. The pressure distribution curves showed that the changes occurring near the ground are not equivalent to a change in the angle of attack. At the lower surface of the section a very strong increase in the pressures is observed. The pressure changes on the upper surface at angles of attack below stalling are insignificant and lead mainly to an increase in the unfavorable pressure gradient, resulting in the earlier occurrence of separation. For a wing with flaps at large angles of attack for distances from the trailing edge of the flap to the ground less than 0.5 chord, the flow between the wing end the ground is retarded so greatly that the pressure coefficient at the lower surface of the section is very near its limiting value (P = 1), and any further possibility of increase in the pressure is very small. In the application an approximate computation procedure is given of the change of certain aerodynamic characteristics for horizontal steady flight near the ground.

Serebrisky, Y. M.↗
NASA NTRS1982

Some observations of separated flow on finite wings

Wind tunnel test results for aspects of flow over airfoils exhibiting single and multiple trailing edge stall 'mushroom' cells are reported. Rectangular wings with aspect ratios of 4.0 and 9.0 were tested at Reynolds numbers of 480,000 and 257,000, respectively. Surface flow patterns were visualized by means of a fluorescent oil flow technique, separated flow was observed with a tuft wand and a water probe, spanwise flow was studied with hot-wire anemometry, smoke flow and an Ar laser illuminated the centerplane flow, and photographs were made of the oil flow patterns. Swirl patterns on partially and fully stalled wings suggested vortex flow attachments in those regions, and a saddle point on the fully stalled AR=4.0 wing indicated a secondary vortex flow at the forward region of the separation bubble. The separation wake decayed downstream, while the tip vortex interacted with the separation bubble on the fully stalled wing. Three mushroom cells were observed on the AR=9.0 wing.

Winkelmann, A. E.↗