Search NASA⌕ Search

SEARCH · Search NASA

Results for “wing aerodynamics”

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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

The aerodynamic design of the oblique flying wing supersonic transport

The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

Vandervelden, Alexander J. M.↗
NASA NTRSDate not supplied

Effect of Geometric Fidelity on the Aerodynamics of a Swept Wing with Scalloped Ice Accretion

The effect of geometric fidelity on the aerodynamics of a swept wing with a “scallop” ice shape was studied. Three geometric fidelity versions of the ice shape were studied. The High Fidelity ice shape maintained all of the highly 3D features of the scallop shape. The 3D Smooth ice shape was smoothly lofted over the High Fidelity shape in order to eliminate all of the locally 3D feature. The third ice shape was the 3D Smooth ice shape with roughness grit attached to the surface. The two 3D Smooth versions of the ice shape exhibited a flowfield characterized by a leading edge separation bubble that rolled into spanwise running vortex. The surface pressure data showed classic leading edge separation bubbles that are observed on 2D airfoils with leading edge horn ice shapes. The High Fidelity ice shape exhibited a flowfield characterized by streamwise vortices that formed downstream of the ice shape. The streamwise vortices observed downstream of the High Fidelity ice shape appeared to reduce the size of the separation suction peak observed downstream of the 3D Smooth ice shapes. This reduced the lift for the High Fidelity ice shape, when compared to the 3D Smooth ice shapes. However, these streamwise vortices may have allowed the flow to remain attached longer and slightly increased the stall angle of attack, even though the maximum lift was lower.

Aircraft icing↗
NASA NTRS1987

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above a Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based on subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outbound control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Seidel, David A.↗
NASA NTRS1987

Investigation of transonic region of high dynamic response encountered on an elastic supercritical wing

Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based upon subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outboard control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.

Seidel, David A.↗
NASA NTRS2013

Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove

Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft s swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First AIAA CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

Bui, Trong↗
NASA NTRS2014

Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove

Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft's swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First American Institute of Aeronautics and Astronautics (AIAA) CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

Bui, Trong T.↗
NASA NTRS1992

Effects of forebody strakes and Mach number on overall aerodynamic characteristics of configuration with 55 deg cropped delta wing

A wind tunnel data base was established for the effects of chine-like forebody strakes and Mach number on the longitudinal and lateral-directional characteristics of a generalized 55 degree cropped delta wing-fuselage-centerline vertical tail configuration. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center at free-stream Mach numbers of 0.40 to 1.10 and Reynolds numbers based on the wing mean aerodynamic chord of 1.60 x 10(exp 6) to 2.59 x 10(exp 6). The best matrix included angles of attack from 0 degree to a maximum of 28 degree, angles of sidesip of 0, +5, and -5 degrees, and wing leading-edge flat deflection angles of 0 and 30 degrees. Key flow phenomena at subsonic and transonic conditions were identified by measuring off-body flow visualization with a laser screen technique. These phenomena included coexisting and interacting vortex flows and shock waves, vortex breakdown, vortex flow interactions with the vertical tail, and vortices induced by flow separation from the hinge line of the deflected wing flap. The flow mechanisms were correlated with the longitudinal and lateral-directional aerodynamic data trends.

Erickson, Gary E.↗
NASA NTRS1982

The equivalent angle-of-attack method for estimating the nonlinear aerodynamic characteristics of missile wings and control surfaces

A method has been developed for estimating the nonlinear aerodynamic characteristics of missile wing and control surfaces. The method is based on the following assumption: if a fin on a body has the same normal-force coefficient as a wing alone composed of two of the same fins joined together at their root chords, then the other force and moment coefficients of the fin and the wing alone are the same including the nonlinearities. The method can be used for deflected fins at arbitrary bank angles and at high angles of attack. In the paper, a full derivation of the method is given, its accuracy demonstrated and its use in extending missile data bases is shown.

Hemsch, M. J.↗
NASA NTRS1950

Use of source distributions for evaluating theoretical aerodynamics of thin finite wings at supersonic speeds

A series of publications on the source-distribution methods for evaluating the aerodynamics of thin wings at supersonic speeds is summarized, extended, and unified. Included in the first part are the deviations of: (a) the linearized partial-differential equation for unsteady flow at a substantially constant Mach number. b) The source-distribution solution for the perturbation-velocity potential that satisfies the boundary conditions of tangential flow at the surface and in the plane of the wing; and (c) the integral equation for determining the strength and the location of sources to describe the interaction effects (as represented by upwash) of the bottom and top wing surfaces through the region between the finite wing boundary and the foremost Mach wave. The second part deals with steady-state thin-wing problems. The third part of the report approximates the integral equation for unsteady upwash and includes a solution of approximate equation. Expressions are then derived to evaluate the load distributions for time-dependent finite-wing motions.

Evvard, John C↗
NASA NTRS2018

Comparison of Iced Aerodynamic Measurements on a Swept Wing from Two Wind Tunnels

Artificial ice shapes of various geometric fidelity were tested on a wing model based on the Common Research Model. Low Reynolds number tests were conducted at Wichita State University's Beech Memorial Wind Tunnel, and high Reynolds number tests were conducted at ONERA's F1 wind tunnel. The aerodynamic performance data from the two facilities were compared at matched or similar Reynolds and Mach number to ensure that the results and trends observed at low Reynolds number could be applied and continued to high Reynolds number. For both clean and iced configurations, the data from Wichita State University and F1 agreed well at matched or similar Reynolds and Mach numbers. The lift and pitching moment curves agreed very well for most configurations. There appeared to be 0.2-0.3deg offset in the angle of attack between the Wichita State University and F1 data, possibly due to different flow angularities in the test sections of the two facilities. There was also an offset in the drag values between the two facilities from an unknown cause. Overall, the data compared very well between the low Reynolds number test at Wichita State University tunnel and the high Reynolds number test at F1. This indicated that data from the low Reynolds number tests could be used to understand iced-swept-wing aerodynamics at high Reynolds number.

Federal Aviation Administration↗
NASA NTRS1981

Effect of leading-edge vortex flaps on aerodynamic performance of delta wings

The effect of leading-edge vortex flaps on the aerodynamic characteristics of highly swept-back wings is analytically investigated, using the free vortex sheet method. The method, based on a three-dimensional inviscid flow model, is an advanced panel type employing quadratic doublet distributions to represent the wing surface, rolled-up vortex sheet and wake and is capable of computing forces, moments and surface pressures.

Reddy, C. S.↗
NASA NTRS1985

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Wood, R. M.↗
NASA NTRS1999

Parallel Computation of Sensitivity Derivatives with Application to Aerodynamic Optimization of a Wing

This paper focuses on the parallel computation of aerodynamic derivatives via automatic differentiation of the Euler/Navier-Stokes solver CFL3D. The comparison with derivatives obtained by finite differences is presented and the scaling of the time required to obtain the derivatives relative to the number of processors employed for the computation is shown. Finally, the derivative computations are coupled with an optimizer and surface/volume grid deformation tools to perform an optimization to reduce the drag of a three-dimensional wing.

Biedron, Robert T.↗
NASA NTRS1995

Subsonic-to-Hypersonic Aerodynamic Characteristics for a Winged, Circular-Body, Single-Stage-to-Orbit Spacecraft Configuration

Experimental aerodynamic characteristics were obtained for a generic, winged, circular-body, single-stage-to-orbit spacecraft configuration. The baseline configuration was longitudinally stable and trimmable at almost all Mach numbers from 0.15 to 10.0--with the exception occurring at low supersonic speeds. Landing speed and subsonic-to-hypersonic longitudinal stability and control appear to be within design guidelines. Lateral-directional instabilities found over the entire speed range, however, create a problem area for this configuration. Longitudinal aerodynamic predictions made utilizing the Aerodynamic Preliminary Analysis System (APAS) were in qualitative, often quantitative agreement with experimental values.

Phillips, W. P.↗