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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Dual wing, swept forward swept rearward wing, and single wing design optimization for high performance business airplanes

An investigation was performed to compare closely coupled dual wing and swept forward swept rearward wing aircraft to corresponding single wing 'baseline' designs to judge the advantages offered by aircraft designed with multiple wing systems. The optimum multiple wing geometry used on the multiple wing designs was determined in an analytic study which investigated the two- and three-dimensional aerodynamic behavior of a wide range of multiple wing configurations in order to find the wing geometry that created the minimum cruise drag. This analysis used a multi-element inviscid vortex panel program coupled to a momentum integral boundary layer analysis program to account for the aerodynamic coupling between the wings and to provide the two-dimensional aerodynamic data, which was then used as input for a three-dimensional vortex lattice program, which calculated the three-dimensional aerodynamic data. The low drag of the multiple wing configurations is due to a combination of two dimensional drag reductions, tailoring the three dimensional drag for the swept forward swept rearward design, and the structural advantages of the two wings that because of the structural connections permitted higher aspect ratios.

Rhodes, M. D.↗
NASA NTRS2019

Preliminary Scale-Resolving Simulations of Laminar-To-Turbulent Transition in Swept-Wing Flow

Swept wings and control surfaces are common elements of modern aircraft, and it has been shown both experimentally and theoretically that laminar-to-turbulent transition of the three-dimensional boundary layer that develops over them is highly sensitive to surface roughness. Numerous studies have been conducted on the effect of discrete roughness elements or distributed roughness elements on swept flow transition, however so far limited computational effort has been dedicated to the study of transition over swept wings with randomly distributed micron-sized roughness. In the present work, we set up to reproduce the extensive experimental data base generated by Dagenhart et al for the infinite swept wing NLF(2)-0415. To this purpose, we perform scale-resolving simulations of flow transition over smooth and rough surfaces using a high-order space-time spectral-element Discontinuous-Galerkin solver. Different types of surface roughnesses are implemented by elastically deforming the original mesh. The study shows that the experimental results cannot be accounted for by a perfectly smooth wing and reveals a strong sensitivity of the transition process to the representation of the surface roughness. The crossflow patterns and transition location approach those measured for some of the surface profiles, however a correlation between the wavenumber spectrum of the surface, grid resolution and boundary layer stability is yet to be established.

Denison, Marie↗
NASA NTRS1957

Investigation at Low Speeds of Deflectors and Spoilers as Gust Alleviators on a Model of the Bell X-5 Airplane with 35 Degree Swept Wings and on a High-aspect-ratio 35 Degree Swept-wing-fuselage Model

Results of an investigation at low speeds to determine the gust-alleviation capabilities (reduction in lift-curve slope) of spoilers and deflectors on a 35 degree swept-wing model of high aspect ratio and on a 1/4-scale model of the X-5 airplane with 35 degree swept wings indicate that deflector and spoiler-deflector types of controls can be designed to provide considerable gust alleviation for a swept-wing airplane while still maintaining stability and control.

CONTROLS, SPOILER - COMPLETE WINGS↗
NASA NTRS1977

A comparison of the experimental aerodynamic characteristics of an oblique wing with those of a swept wing

Force and moment characteristics were measured for two trapezoidal oblique wings and a conventional swept wing mounted on a body of revolution at Mach numbers from 0.25 to 2.0. Both oblique wings had the same planform, but differed in profile and flexibility. One of the oblique wings was made of solid steel and had a maximum thickness-to-chord ratio of 4 percent. The other wing was built up by taking an aluminum wing and adding epoxy material to the upper surface to increase the maximum thickness-to-chord ratio to 8.2 percent. The aspect ratio for both oblique wings when swept 45 deg, and for the swept wing with 45 deg of sweep, was 4.1. Data were obtained at unit Reynolds numbers ranging from 3.3 to 8.2 million per meter and were compared with previously obtained data on the aluminum wing before it was built up with epoxy. Wing flexibility designed into the aluminum and built-up aluminum oblique wings increased the range of lift coefficients from 0.30 to 0.70 over which the pitching-moment curves were linear. However, flexibility did not improve the linearity of the rolling-moment curves and produced sizable side forces. At a Mach number of 0.95, the trapezoidal oblique wing had little or no improvement in the lift/drag ratios over those for a conventional swept wing of the same aspect ratio, sweep, and profile.

Hopkins, E. J.↗
NASA NTRS1973

Space shuttle: Pressure investigation of a space shuttle launch configuration consisting of a delta-wing orbiter and a swept-wing booster with canard and tip fans (M equals 0.6 to 1.3). Volume 1, part A: Booster data

Wind tunnel tests to determine the pressure distribution on a space shuttle launch configuration consisting of a delta wing orbiter and a swept wing booster with canard and tip fins were conducted. Pressure data were obtained for the combined orbiter and booster and for the booster alone at Mach numbers from 0.6 to 1.3, angles of attack from minus 8 degrees to plus 10 degrees, and sideslip angles from minus 6 degrees to plus 6 degrees. Pressure data were also obtained for the booster alone without canard at Mach numbers of 0.9 and 1.1. The pressure taps were distributed primarily over the booster upper surface and the orbiter lower surface.

Rampy, J. M.↗
NASA NTRS1951

Error in Airspeed Measurement Due to Static-Pressure Field Ahead of the Wing Tip of a Swept-Wing Airplane Model at Transonic Speeds

As part of a study of means of airspeed measurement at transonic speeds the use of static orifices located ahead of the wing tip has been investigated for possible application to service or research airspeed installations. The local static pressure and local Mach number have been measured at a distance of 1 tip chord ahead of the wing tip of a model of a swept-wing fighter airplane at true Mach numbers between 0.7 and 1.08 by the NACA wing-flow method. All measurements were made at or near zero lift. The local Mach number was found to be essentially equal to the true Mach numbers less than about 0.90. The local Mach number was found to be about 0.97 at a true Mach number of 0.95, and to be about 1.04 at a true Mach number of 1.08. The local Mach number provided a reasonable sensitive measure of true Mach number except for a restricted region near a true Mach number of 1.0 where the local Mach number did not change appreciably with true Mach number. The linear theory was found to predict qualitatively the effect of the fuselage on the static pressure ahead of the wing time but gave a reasonable prediction of the effect of the wing on the static pressure only at Mach numbers below 0.95.

Danforth, Edward C.B.↗
NASA NTRS1953

A Method for Determining Cloud-Droplet Impingement on Swept Wings

The general effect of wing sweep on cloud-droplet trajectories about swept wings of high aspect ratio moving at subsonic speeds is discussed. A method of computing droplet trajectories about yawed cylinders and swept wings is presented, and illustrative droplet trajectories are computed. A method of extending two-dimensional calculations of droplet impingement on nonswept wings to swept wings is presented. It is shown that the extent of impingement of cloud droplets on an airfoil surface, the total rate of collection of water, and the local rate of impingement per unit area of airfoil surface can be found for a swept wing from two-dimensional data for a nonswept wing. The impingement on a swept wing is obtained from impingement data for a nonswept airfoil section which is the same as the section in the normal plane of the swept wing by calculating all dimensionless parameters with respect to flow conditions in the normal plane of the swept wing.

Dorsch, Robert G.↗
NASA NTRSDate not supplied

Ice Shape Classification for Swept Wings

An important consideration for swept-wing ice shapes is the level of geometric fidelity required to accurately capture the aerodynamics. One way to help make the task of addressing this more manageable is to organize the ice shapes into a small number of classifications. These classifications are done by grouping similar ice shape geometries that also have similar iced wing aerodynamics. Ice shape classifications for airfoils are more mature and are reviewed as an example. Swept wing ice shape classifications are based on the well-established airfoil classifications and here five classifications are proposed: leading-edge roughness, streamwise ice, 3D leading-edge horn, highly 3D leading-edge horn and spanwise ridge. Recent data on swept wing icing aerodynamics is used to develop and support this classification. The development, or suppression, of the leading-edge separation vortex on swept wings without a leading-edge device is seen to be key in the classification of the highly 3D leading-edge horn.

Aircraft icing↗
NASA NTRS1990

Predicted and measured in-flight wing deformations of a forward-swept-wing aircraft

An electrooptical flight-deflection measurement system (FDMS) is described in terms of its use in structural testing of the composite forward-swept wing of the X-29 aircraft. The wing deflected shapes measured using the present system are compared to the shapes predicted by NASTRAN and other codes as well as data from ground-test load measurements. The electrooptical FDMS is based on a control unit, two receivers, a target driver, and 12-16 IR LED targets. The FDMS determines the in-flight deflected wing shapes at a variety of altitudes at Mach 0.9, and the results are compared to the analytically predicted wing-twist distributions. The FDMS data describe the predicted increasing streamwise twist with increasing dynamic pressure and suggest that the streamwise twist is more prevalent at the inboard measurement station than at the wing tip. This hook shape is not represented in the predicted data, and suggestions are given for improving the modeling of the X-29 wing.

Lokos, William A.↗
NASA NTRS1986

Forward-swept wing configuration designed for high maneuverability by use of a transonic computational method

A transonic computational analysis method and a transonic design procedure have been used to design the wing and the canard of a forward-swept-wing fighter configuration for good transonic maneuver performance. A model of this configuration was tested in the Langley 16-Foot Transonic Tunnel. Oil-flow photographs were obtained to examine the wind flow patterns at Mach numbers from 0.60 to 0.90. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of the forward-swept-wing configuration with an equivalent aft-swept-wing-configuration showed that, at a Mach number of 0.90 and a lift coefficient of 0.9, the two configurations have the same trimmed drag. The forward-swept wing configuration was also found to have trimmed drag levels at transonic maneuver conditions which are comparable to those of the HiMAT (highly maneuverable aircraft technology) configuration and the X-29 forward-swept-wing research configuration. The configuration of this study was also tested with a forebody strake.

Mann, M. J.↗
NASA NTRS1978

Effects of wing leading-edge flap deflections on subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing

An investigation was conducted to determine the effects of wing leading-edge flap deflections on the subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing. The tests were conducted at Mach numbers from 0.40 to 0.85, corresponding to Reynolds numbers (based on wing mean geometric chord) of 2.37 x 1,000,000 to 4.59 x 1,000,000 and at angles of attack from -3 deg to 22 deg. The configurations under study included a wing-fuselage configuration and a wing-fuselage-strake configuration. Each configuration had multisegmented, constant-chord leading-edge flaps which could be deflected independently or in various combinations.

Henderson, W. P.↗