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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.

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At least 883 records · Page 49
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NASA NTRS1975

A three-dimensional solution of flows over wings with leading-edge vortex separation

The application of a new, general, potential flow computational technique to the solution of the subsonic, three-dimensional flow over wings with leading-edge vortex separation is presented. The present method is capable of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings of rather arbitrary planform. The wing geometry is arbitrary in the sense that leading and trailing edges may be curved or kinked and the wing may have arbitrary camber and twist. The method employs an inviscid flow model in which the wing, the rolled-up vortex sheets, and the wake are represented by piecewise continuous quadratic doublet sheet distributions. The Kutta condition is imposed along all wing edges. Strengths of the doublet distributions as well as shape and position of the free fortex sheet spirals are computed in iterative fashion starting with an assumed initial sheet geometry. The method is verified by numerous computed results.

Weber, J. A.↗
NASA NTRS1975

Prediction of span loading of straight-wing/propeller combinations up to stall

A method is presented for calculating the spanwise lift distribution on straight-wing/propeller combinations. The method combines a modified form of the Prandtl wing theory with a realistic representation of the propeller slipstream distribution. The slipstream analysis permits calculations of the nonuniform axial and rotational slipstream velocity field of propeller/nacelle combinations. This nonuniform field was then used to calculate the wing lift distribution by means of the modified Prandtl wing theory. The theory was developed for any number of nonoverlapping propellers, on a wing with partial or full-span flaps, and is applicable throughout an aspect ratio range from 2.0 and higher. A computer program was used to calculate slipstream characteristics and wing span load distributions for a number of configurations for which experimental data are available, and favorable comparisons are demonstrated between the theoretical predictions and the existing data.

Mcveigh, M. A.↗
NASA NTRS1975

Transonic wind-tunnel tests of an F-8 airplane model equipped with 12 and 14-percent thick oblique wings

An experimental investigation was conducted in the Ames 14-foot transonic wind tunnel to study the aerodynamic performance and stability characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing. Two elliptical planform (axis ratio = 8:1) wings, each having a maximum thickness of 12 and 14 percent, were tested. Longitudinal stability data were obtained with no wing and with each of the two wings set at sweep angles of 0, 45, and 60 deg. Lateral directional stability data were obtained for the 12 percent wing only. Test Mach numbers ranged from 0.6 to 1.2 in the unit Reynolds number range from 11.2 to 13.1 million per meter. Angles of attack were between -6 and 22 deg at zero sideslip. Angles of sideslip were between -6 and +6 deg for two angles of attack, depending upon the wing configuration.

Smith, R. C.↗
NASA NTRS1976

A preliminary design study of a laminar flow control wing of composite materials for long range transport aircraft

The results of an aircraft wing design study are reported. The selected study airplane configuration is defined. The suction surface, ducting, and compressor systems are described. Techniques of manufacturing suction surfaces are identified and discussed. A wing box of graphite/epoxy composite is defined. Leading and trailing edge structures of composite construction are described. Control surfaces, engine installation, and landing gear are illustrated and discussed. The preliminary wing design is appraised from the standpoint of manufacturing, weight, operations, and durability. It is concluded that a practical laminar flow control (LFC) wing of composite material can be built, and that such a wing will be lighter than an equivalent metal wing. As a result, a program of suction surface evaluation and other studies of configuration, aerodynamics, structural design and manufacturing, and suction systems are recommended.

Swinford, G. R.↗
NASA NTRS1976

Effect of Krueger nose flaps on the experimental force and moment characteristics of an oblique wing

Experimental force and moment data are presented for an oblique wing mounted on a body of revolution and equipped with Krueger type nose flaps. The effectiveness of these flaps in making the moment curves more linear by controlling the flow separation on the downstream wing panel at high lift coefficients was determined. The investigation of the effects of the Krueger flaps covered two cases: (1) use of the flaps on the downstream wing panel only and (2) use of the flaps on both wing panels. For part of the tests, the Krueger flaps were mounted on nose flaps that were drooped either 5 deg or 10 deg. The wing was elliptical in planform, had an aspect ratio of 6.0 (based on the unswept span) and was tested at sweep angles of 0, 45 deg, and 50 deg. The Mach-number range covered was from 0.25 to 0.95. It was found that the most effective arrangement of the Krueger flaps for making the pitching-, rolling-, and yawing-moment curves more linear at high lift coefficients was having the Krueger flaps mounted on the nose flaps drooped 5 deg and only on the downstream wing panel.

Hopkins, E. J.↗
NASA NTRS1976

Experimental effects of fuselage camber on longitudinal aerodynamic characteristics of a series of wing-fuselage configurations at a Mach number of 1.41

An experimental investigation was conducted to evaluate a method for the integration of a fighter-type fuselage with a theoretical wing to preserve desirable wing aerodynamic characteristics for efficient maneuvering. The investigation was conducted by using semispan wing fuselage models mounted on a splitter plate. The models were tested through an angle of attack range at a Mach number of 1.41. The wing had a leading edge sweep angle of 50 deg and an aspect ratio of 2.76; the wing camber surface was designed for minimum drag due to lift and was to be self trimming at a lift coefficient of 0.2 and at a Mach number of 1.40. A series of five fuselages of various camber was tested on the wing.

Dollyhigh, S. M.↗
NASA NTRS1977

Aerodynamic loads near cranks, apexes, and tips of thin, lifting wings in incompressible flow

The calculation of the incompressible and irrotational flow in the vicinity of tips and corners of thin, lifting wings is considered. It is shown that the important characteristics of the flow are governed by an eigenvalue problem, which is nonlinear at the trailing edge because of the shed wake (assumed to be in the wing plane). A new solution method was devised because either the existing methods were not valid for the trailing edge case or they would have required excessive amounts of computer time. The new method, which is fundamentally different than the previous ones, was used to calculate solutions for a number of cases, including some for which correct answers had not previously been obtained. Two of these solutions were used to determine the validity of drag and leading-edge-suction distributions near the tips of a delta wing and a swept wing as calculated by using both the vortex lattice method and a kernel function method. The calculations for the swept wing resolved the question of whether or not the induced drag should be zero at the wing tip.

Medan, R. T.↗
NASA NTRS1977

Aerodynamic characteristics at Mach 6 of a hypersonic research airplane concept having a 70 deg swept delta wing

The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.

Clark, L. E.↗
NASA NTRS1977

Numerical calculation of the transonic flow past a swept wing

A numerical method is presented for analyzing the transonic potential flow past a lifting, swept wing. A finite difference approximation to the full potential equation is solved in a coordinate system which is nearly conformally mapped from the physical space in planes parallel to the symmetry plane, and reduces the wing surface to a portion of one boundary of the computational grid. A coordinate invariant, rotated difference scheme is used, and the difference equations are solved by relaxation. The method is capable of treating wings of arbitrary planform and dihedral, although approximations in treating the tips and vortex sheet make its accuracy suspect for wings of small aspect ratio. Comparisons of calculated results with experimental data are shown for examples of both conventional and supercritical transport wings. Agreement is good for both types, but it was found necessary to account for the displacement effect of the boundary layer for the supercritical wing, presumably because of its greater sensitivity to changes in effective geometry.

Jameson, A.↗
NASA NTRS1977

Buffet characteristics of the F-8 supercritical wing airplane

The buffet characteristics of the F-8 supercritical wing airplane were investigated. Wing structural response was used to determine the buffet characteristics of the wing and these characteristics are compared with wind tunnel model data and the wing flow characteristics at transonic speeds. The wingtip accelerometer was used to determine the buffet onset boundary and to measure the buffet intensity characteristics of the airplane. The effects of moderate trailing edge flap deflections on the buffet onset boundary are presented. The supercritical wing flow characteristics were determined from wind tunnel and flight static pressure measurements and from a dynamic pressure sensor mounted on the flight test airplane in the vicinity of the shock wave that formed on the upper surface of the wing at transonic speeds. The comparison of the airplane's structural response data to the supercritical flow characteristics includes the effects of a leading edge vortex generator.

Deangelis, V. M.↗
NASA NTRS1977

Transonic wind-tunnel investigation of the maneuver potential of the NASA supercritical wing concept, phase 1

An investigation was conducted in the NASA Langley 8-foot transonic pressure tunnel at Mach numbers from 0.60 to 0.975 with a variable-wing-sweep airplane model in order to evaluate a series of wings designed to demonstrate the maneuver potential of the supercritical airfoil concept. Both conventional and supercritical wing designs for several planform configurations were investigated with wing sweep angles from 16.0 deg to 72.5 deg, depending on Mach number and wing configuration. The supercritical wing configuration showed significant improvements over the conventional configurations in drag-divergence Mach number and in drag level at transonic maneuver conditions.

Hallissy, J. B.↗
NASA NTRS1977

An elementary analysis of the effect of sweep, Mach number, and lift coefficient on wing-structure weight

Results are presented from an elementary analysis of the effect of sweep angle on the idealized structural weight of swept wings, with cruise Mach number M and lift coefficient C sub L as parameters. The analysis indicates that sweep is unnecessary for cruise Mach numbers below about 0.80, whereas for the higher subsonic speeds, a well defined minimum-weight condition exists at a sweep angle in the neighborhood of 35 deg or 40 deg, depending on M and C sub L. The results further indicate that wing-structure weight increases sharply with Mach number in the high subsonic range, with Mach 0.85 wings weighing half again as much as Mach 0.75 wings. Weight is also shown to increase with cruise lift coefficient, but the effect is not strong for the usual range of design lift coefficients. Minimum wing-structure weight is found to occur at a ratio of thickness to normal chord of about 18 percent, but it is concluded that the thickness ratio for optimum wing design would probably lie in the range of 12 to 15 percent.

Kyser, A. C.↗
NASA NTRS1977

Theoretical prediction of over-wing-blowing aerodynamics

A theoretical method is developed for predicting the aerodynamic characteristics of wings with over-wing-blowing jet. The method allows the jet to stay above the wing surface as well as to wash the surface. It accounts for the wing-jet interaction due to differences between the jet and freestream dynamic pressures and Mach numbers, in addition to the jet entrainment. For the former effect, the quasi-vortex-lattice method is used to satisfy the jet and wing boundary conditions. For the latter, a new theory was developed to calculate the jet entrained flow for given jet properties. Comparison of predicted results with available data of various configurations shows reasonably good agreement. Further theoretical analysis indicates that it is aerodynamically advantageous to locate the jet exit near and ahead of the wing leading edge, and that the camber shape has significant effect on the induced drag.

Lan, C. E.↗
NASA NTRS1977

Load distribution on a close-coupled wing canard at transonic speeds

This paper reports on a wind-tunnel test where load distributions were obtained at transonic speeds on both the canard and wing surfaces of a closely-coupled wing-canard configuration. The investigation included detailed component and configuration arrangement studies to provide insight into the various aerodynamic interference effects for the leading-edge vortex flow conditions encountered. Data indicate that increasing the Mach number from 0.70 to 0.95 caused the wing leading-edge vortex to burst over the wing when the wing was in the presence of the high canard. For some of the outboard span locations, the leading-edge vortex reattachment streamline intersects the wing trailing edge inboard of these span locations, thus, the Kutta condition was not satisfied. In general, the effect of adding a canard was to reduce the lift inboard and somewhat increase the lift outboard similar to the trends that would have been expected had the flow been attached.

Gloss, B. B.↗
NASA NTRS1977

Wing design by numerical optimization

A study was conducted to assess the feasibility of performing computerized wing design by numerical optimization. The design program combined a full potential, inviscid aerodynamics code with a conjugate gradient optimization algorithm. Three design problems were selected to demonstrate the design technique. The first involved modifying the upper surface of the inboard 50% of a swept wing to reduce the shock drag subject to a constraint on wing volume. The second involved modifying the entire upper surface of the same swept wing (except the tip section) to increase the lift-drag ratio subject to constraints on wing volume and lift coefficient. The final problem involved modifying the inboard 50% of a low-speed wing to achieve good stall progression. Results from the three cases indicate that the technique is sufficiently accurate to permit substantial improvement in the design objectives.

Hicks, R. M.↗
NASA NTRS1973

Preliminary lift and drag characteristics of the F-8 supercritical wing airplane

A flight research program is being conducted using a supercritical wing attached to a modified F-8 fuselage to demonstrate the potential benefits of the supercritical wing concept in a realistic cruise flight environment. One of the primary goals of the flight program is to determine the drag-divergence Mach number, that is, the Mach number at which the total vehicle drag increases significantly. Another objective of the program is to show how well wind-tunnel data can be used to predict the performance of a supercritical wing configuration in the transonic Mach number region from 0. 9 to 1. 0. Comparisons of this nature are important in view of the uncertainties in the present methods for extrapolating wind-tunnel-model data to flight Reynolds numbers, determining the effect of wall interference on the model results, and simulating the flow separation characteristics on the model wing to correspond to the large-scale wing. However, in comparing these wind-tunnel and flight results, it must be taken into account that this supercritical wing test configuration is of only intermediate scale und that a full-scale transport would have Reynolds numbers three to five times larger.

Jon S. Pyle↗
NASA NTRS1977

Aerodynamic characteristics of an F-8 aircraft configuration with a variable camber wing at Mach numbers from 1.5 to 2.0

A 0.1-scale model of an F-8 aircraft was tested over a range of Mach numbers from 1.5 to 2.0. Reynolds number of 4.12 million was based on wing mean-aerodynamic chord for angles of attack varying from -2 deg to +12 deg. The model was equipped with an advanced-technology-conformal-variable-camber wing (ATCVCW) having simple hinge flaps. Data were also obtained for the model equipped with the basic F-8 wing and conventional flaps. Model variables included aileron and wing trailing edge deflections and horizontal tail incidence. The ATCVCW configuration produced slight improvements in lift-curve slope, drag, and static longitudinal stability over that of the basic F-8 wing configuration. Flap effectiveness was essentially the same for both wings.

Boltz, F. W.↗
NASA NTRS1976

Comparison of aerodynamic theory and experiment for jet-flap wings

Aerodynamic theory predictions made for a jet flapped wing were compared with experimental data obtained in a fairly extensive series of tests in the Langley V/STOL tunnel. The tests were made on a straight, rectangular wing and investigated two types of jet flap concepts: a pure jet flap with high jet deflection and a wing with blowing at the knee of a plain trailing edge flap. The tests investigated full and partial span blowing for wing aspect ratios of 8.0 and 5.5 and momentum coefficients from 0 to about 4. The total lift, drag, and pitching moment coefficients predicted by the theory were in excellent agreement with experimental values for the pure jet flap, even with the high jet deflection. The pressure coefficients on the wing, and hence the circulation lift coefficients, were underpredicted, however, because of the linearizing assumptions of the planar theory. The lift, drag, and pitching moment coefficients, as well as pressure coefficients, were underpredicted for the wing with blowing over the flap because of the failure of the theory to account for the interaction effect of the high velocity jet passing over the flap.

Gainer, T. G.↗