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Results for “LOW ASPECT RATIO WING”

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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Flutter analysis of low aspect ratio wings

Several very low aspect ratio flat plate wing configurations are analyzed for their aerodynamic instability (flutter) characteristics. All of the wings investigated are delta planforms with clipped tips, made of aluminum alloy plate and cantilevered from the supporting vehicle body. Results of both subsonic and supersonic NASTRAN aeroelastic analyses as well as those from another version of the program implementing the supersonic linearized aerodynamic theory are presented. Results are selectively compared with the experimental data; however, supersonic predictions of the Mach Box method in NASTRAN are found to be erratic and erroneous, requiring the use of a separate program.

Parnell, L. A.↗
NASA NTRS1984

Computational-experimental pressure distributions on a transonic, low-aspect-ratio wing

A generic, transonic, supercritical, low aspect-ratio wing was tested at design incidence of 5 deg at Mach numbers from 0.25 to 0.96. Oil-flow studies at the design Mach number of 0.85 showed local-flow separation, which, in retrospect, might have been avoided. At Mach 0.82 with unseparated flow, the surface-flow angles were less than 10 deg. Predictions with the FLO22 transonic potential code are good. Lift interference is strong without tunnel-wall suction. Evidence from this study shows that wings that are optimized for mild shock waves and pressure-recovery gradients generally have small, three-dimensional flow at conditions for unseparated flow.

Keener, E. R.↗
NASA NTRS1984

Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

The viscous transonic flow around a low aspect ratio wing was computed by an implicit, three dimensional, thin-layer Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. A low aspect ratio wing with large sweep, twist, taper, and camber is the chosen geometry. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. The flow around the wing was computed for a free stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three dimensional flow separation off the wind surface are observed. Particle path lines indicate that the three dimensional flow separation on the wing surface is part of the roots of the tip vortex formation. The lifting of the tip vortex before the wing trailing edge is observed by following the trajectory of particles release around the wing tip.

Mansour, N. N.↗
NASA NTRS1984

Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

The viscous transonic flow around a low-aspect-ratio wing has been computed using an implicit, three-dimensional, 'thin-layer' Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. The geometry chosen for this study is a low-aspect-ratio wing with large sweep, twist, taper, and camber. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. Using this grid, the flow around the wing was computed for a free-stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three-dimensional flow separation off the wing surface are observed. Particle path lines indicate that the three-dimensional flow separation on the wing surface is part of the roots of the tip-vortex formation. The lifting of the tip vortex before the wing trailing edge is clearly observed by following the trajectory of particles released around the wing tip.

Mansour, N. N.↗
NASA NTRS1986

Numerical simulation of transonic separated flows over low-aspect ratio wings

Transonic flow fields about a low-aspect-ratio advanced technology wing have been computed using a viscous/inviscid zonal approach. The flow field near the wing where viscous effects are important was solved using the 'Reynolds-Averaged Navier-Stokes Equations' in 'thin-layer' form. The Euler equations were used to determine the flow field in regions away from the wing where viscous effects are insignificant. A zonal grid using an H-H topology was generated around the wing by first solving a set of Poisson's equations for the global grid. This grid was then subdivided into separate zones of viscous or inviscid flow as suggested by the flow physics. A series of flow cases were computed and compared with corresponding sets of experimental data. All cases showed good agreement with experiment in terms of the pressure field. Also, a good correlation between computed separated surface flow and experimental oil flow was obtained.

Kaynak, U.↗
NASA NTRS1992

Low aspect ratio wing code validation experiment

An experiment was performed on a low-aspect ratio wing to define the limits of the validity of codes which are used to predict the flow over modern fighter configurations. The focus of the experiment is to provide data at or near the true operating conditions of transonic aircraft. The Reynolds numbers reached in this experiment reproduce those experienced by an aircraft with a 5-m root chord, flying at 14 to 16 km, from Mach 0.6 to 0.8. The angle-of-attack range of the experiment extends into the regime of leading-edge separation.

Olsen, Michael E.↗
NASA NTRS1994

Low aspect ratio wing experiment

This test was initiated to provide validation data on low aspect ratio wings at transonic speeds. The test was conducted so that the data obtained would be useful in the validation of codes, and all boundary condition data required would be measured as part of the test. During the conduct of the test, the measured quantities were checked for repeatability, and when the data would not repeat, the cause was tracked down and either eliminated or included in the measurement uncertainty. The accuracy of the data was in the end limited by wall imperfections of the wind tunnel in which the test was run.

Olsen, Mike↗
NASA NTRS1952

Effect of Aspect Ratio on the Low-Speed Lateral Control Characteristics of Untapered Low-Aspect-Ratio Wings Equipped with Flap and with Retractable Ailerons

A low-speed wind-tunnel investigation was made to determine the lateral control characteristics of a series of untapered low-aspect-ratio wings. Sealed flap ailerons of various spans and spanwise locations were investigated on unswept wings of aspect ratios 1.13, 1.13, 4.13, and 6.13; and various projections of 0.60-semispan retractable ailerons were investigated on the unsweptback wings of aspect ratios 1.13, 2.13, and 4.13 and on a 45 degree sweptback wing. The retractable ailerons investigated on the unswept wings spanned the outboard stations of each wing; whereas the plain and stepped retractable ailerons investigated on the sweptback wing were located at various spanwise stations. Design charts based on experimental results are presented for estimating the flap aileron effectiveness for low-aspect-ratio, untapered, unswept.

Fischel, Jack↗
NASA NTRS1977

Flight measurements of lifting pressures for a thin low-aspect-ratio wing at subsonic, transonic, and low supersonic speeds

Pressure distributions in the form of differential pressure coefficients are presented for several wing chordwise and spanwise stations. Also presented are the results of limited analysis which show aircraft configuration effects, Mach number effects on the local wing loadings, comparisons of selected measured wing pressures with predicted pressures, and comparisons of wing loadings during right-turn and left-turn maneuvers.

Byrdsong, T. A.↗
NASA NTRS1953

Aerodynamic characteristics of low-aspect-ratio wings at high supersonic Mach numbers

This paper presents some recently obtained data on the aerodynamic characteristics of low-aspect-ratio wings at supersonic Mach numbers of 4.04 and 6.9 and discusses some new methods of predicting the lift and drag of such wings. Data on lifting wings in the Mach number range above 2.5 are not plentiful and most of the available data may be found in NACA RM L51D17, NACA RM L51D30, NACA RM L52D15a, NACA RM L52K19, NACA RM L53D03, and NACA RM L53D30a

Ulmann, Edward F↗
NASA NTRS1985

Lift due to thickness for low aspect ratio wings in incompressible flow

The problem under consideration is a numerical study of the effects of thickness on lift for low aspect ratio wings in steady incompressible inviscid flow at moderate angles of attack. At these angles of attack the flow separates along the leading edge giving rise to a lift substantially higher than that computed by classical attached flow potential theory. The problem is treated as a perturbation expansion in a small thickness parameter. The lifting elements of the flow are modeled using a nonlinear vortex lattice method which replaces the leading and trailing edge vortex sheets by segmented straight vortex filaments. The thickness elements of the flow are modeled with a mean plane source distribution and a modification to the wing boundary conditions. Results are obtained for wings with biconvex and NACA 0012 sections which compare well with available experimental data. The important observation that the effect of thickness is to decrease the lift is made.

Dodbele, S. S.↗
NASA NTRS1987

Loss of lift due to thickness for low-aspect-ratio wings in incompressible flow

The problem under consideration is a numerical study of the effects of thickness on lift for low-aspect-ratio wings in steady incompressible inviscid flow at moderate angles of attack. At these angles of attack the flow separates along the leading edge giving rise to a lift substantially higher than that computed by classical attached-flow potential theory. The problem is treated as a perturbation expansion in a small thickness parameter. The lifting elements of the flow are modeled using a nonlinear vortex-lattice method which replaces the leading and trailing-edge vortex sheets by segmented straight vortex filaments. The thickness elements of the flow are modeled with a mean-plane source distribution and a modification to the wing boundary conditions. Results are obtained for wings with biconvex and NACA 0012 sections which compare well with available experimental data. The important observation that the effect of thickness is to decrease the lift is made.

Dodbele, S. S.↗
NASA NTRS1957

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

A full- scale rocket-powered model of a cruciform canard missile configuration with a low- aspect - ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed- control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift - curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift- .curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta= -0.3 deg . The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic- center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number.The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Brown, Clarence A , Jr↗
NASA NTRS1991

Evaluation of a three-dimensional Navier-Stokes solver for transonic flow over a low-aspect-ratio wing mounted in a wind tunnel

A three dimensional Navier-Stokes solver is evaluated for transonic flow over a thin, swept, low-aspect ratio wing. The computational study was undertaken in support of a wind tunnel experimental program. The computational results are compared to experimental surface pressure data obtained in a cryogenic wind tunnel with an adaptive wall test section. The results show favorable agreement over a wide range of conditions, further the numerical results provide additional data of the complex three-dimensional flow field. Differences in the predictions and experiment suggest a need to conduct further experiments to evaluate the adaptive wall testing technique, and to model the tunnel sidewall boundary layer in the computations.

Milholen, William E., II↗