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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Effect of Reduced Frequency on Super Maneuver Delta Wing

The unsteady, three-dimensional, Reynolds averaged Navier Stokes equations (NS) are used to simulate and study the flow response around a delta wing undergoing a pitch-up motion up to 90 degrees amplitude. The governing equations are solved time-accurately using the upwind. Roe flux-difference splitting, finite volume scheme. The primary model under consideration consists of a 76 degree swept, sharp-edged delta wing of zero thickness. The Reynolds number and Mach number are 0.45 X 10(exp 6) and .03 respectively. The wing is forced to pitch-up through a ramp function, using a wide range of reduced frequency to provide a good understanding of reduced frequency effect on high amplitude pitch maneuvers. Three reduced frequency values of .0834, 0.134 and 0.04, are used in this study. The computed results are compared with each other as well as the existing experimental data.

Abdelhamid, Yahia A.↗
NASA NTRS1977

Nonlinear, unsteady aerodynamic loads on rectangular and delta wings

Nonlinear unsteady aerodynamic loads on rectangular and delta wings in an incompressible flow are calculated by using an unsteady vortex-lattice model. Examples include flows past fixed wings in unsteady uniform streams and flows past wings undergoing unsteady motions. The unsteadiness may be due to gusty winds or pitching oscillations. The present technique establishes a reliable approach which can be utilized in the analysis of problems associated with the dynamics and aeroelasticity of wings within a wide range of angles of attack.

Atta, E. H.↗
NASA NTRS1986

Steady supersonic Navier-Stokes solutions of a 75 deg delta wing

Steady solutions about a slender sharp edged delta wing in a supersonic freestream for moderate and high angles of attack are obtained numerically by time integration of the unsteady compressible three dimensional laminar Navier-Stokes equations. The main features of the flow, including primary and secondary separation, and vortex position and strength, are adequately simulated in the numerical solutions. Improved resolution of the computational grid in the leading edge region from a previous solution had considerable effect on the accuracy of the solutions. Good agreement between numerical solutions and experimental data was obtained for two cases. A local timestepping procedure is used to speed convergence by approximately a factor of two.

Buter, T. A.↗
NASA NTRS2001

A Study of Grid Resolution, Transition and Turbulence Model Using the Transonic Simple Straked Delta Wing

Three-dimensional transonic flow over a delta wing is investigated using several turbulence models. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition aft of the leading edge or are fully turbulent are performed. These computations show that grid resolution, transition location and turbulence model significantly affect the 3D flowfield.

Bartels, Robert E.↗
NASA NTRS1985

A three-dimensional boundary-layer method for flow over delta wings with leading-edge separation

A three-dimensional, laminar boundary-layer method is applied to the incompressible flow over a slender delta wing at incidence. The predictor-corrector finite-difference scheme of Matsuno is used to difference the governing equations. The method has the advantages that no iterations are required to advance the solution and the cross-flow derivatives are formed independent of the cross-field direction. The difference scheme is demonstrated to yield accurate numerical results when compared to the exact solution of the three-dimensional boundary-layer equations for parabolic flow over a moving flat plate. The method is applied to delta wings of various sweep angles at angles of attack up to 20 deg., with the inviscid solution determined using a higher-order, three-dimensional panel method.

Woodson, S. H.↗
NASA NTRS1995

Progress in Navier-Stokes Simulation of Delta Wing Maneuver Aerodynamics

This paper will summarize recent progress in the numerical simulation of high incidence vortical flow about a generic 65 degree sweep delta wing using the three dimensional, time-dependent, Reynolds averaged, Navier-Stokes (RANS) equations. Computations have been carried out at 15 and 30 degrees angle of attack under subsonic turbulent flow conditions, and compared with experimental data provided by Hanff, Jenkins, and their colleagues. This work has already been published elsewhere and widely disseminated. Computations carried out at 15 degrees angle of attack included static roll angles ranging up to 65 degrees, and a large-amplitude (40 degree), high rate (7 Hz), forced roll motion. There was very good agreement between computed and experimental forces and moments, and static surface pressures. There was a significant hysteresis of the dynamic rolling moment due to the high rate of roll motion. At this angle of attack, no vortex breakdown was observed in the computations or experiment. Computations were also carried out at 30 degrees angle of attack, where vortex breakdown was present in both the computations and experiment. There was overall good agreement in the computed and experimental forces and moments. The static rolling moment varied with roll angle in a highly nonlinear manner, and exhibited three stable trim points and two unstable trim points. This behavior was attributed to the presence of vortex breakdown. Two large-amplitude (30 degrees), high-rate (10 Hz) forced roll motions were computed. The dynamics of the vortex breakdown motion was dramatically visualized by tracking the time-dependent motion of particles released near the delta wing apex. This numerical visualization is analogous to experimental smoke flow techniques. In one of the dynamic cases the breakdown was found to move off the wing, convected downwind of the trailing edge, and later reformed near the trailing edge through an instability of the vortex core. A damped free-to-roll motion was also computed by releasing the wing from rest at 40 degrees of roll. The wing went to the same trim point as in the experiment.

Chaderjian, Neal M.↗
NASA NTRS1972

Lee-side vortices on delta wings at hypersonic speeds.

A fluid-dynamic investigation was carried out to determine the cause of intense heating observed on the lee meridian of hypersonic delta wings and also to derive means for its suppression. Several experimental techniques were combined with analysis of extensive heat-transfer measurements at a freestream Mach number of six in a range of Reynolds number to acquire a general description of the lee-flow structure. With attached leading-edge flow on the delta wings, the dominant feature is a pair of embedded vortices on the lee meridian whose interaction with the boundary-layer is responsible for the observed local heating. On the basis of flow visualization results and heat-transfer correlations, a qualitative vortex flow model is proposed which differs essentially from the conventional inboard separation vortex model.

Rao, D. M.↗
NASA NTRS1984

Numerical and experimental determination of secondary separation on delta wings in subsonic flow

A quasi-three-dimensional boundary layer method for conical inviscid flows is presented. The model was developed to characterize the flow fields over slender delta wings, particularly the shed vortex sheet. A steady, incompressible, laminar boundary layer is assumed and a solution is obtained with the Smith (1966) finite difference code. Predictions are compared with oil flow patterns and surface pressure measurements for 74 and 80 deg delta wings at 0-20 deg angles of attack. The model is capable of accurately predicting the secondary separation lines. Inviscid velocities derived from panel methods lead, however, to inaccurate pressure distributions.

Dejarnette, F. R.↗
NASA NTRS1959

Evaluation of the Levy Method as Applied to Vibrations of a 45 deg Delta Wing

The Levy method which deals with an idealized structure was used to obtain the natural modes and frequencies of a large-scale built-up 45 deg. delta wing. The results from this approach, both with and without the effects of transverse shear, were compared with the results obtained experimentally and also with those calculated by the Stein-Sanders method. From these comparisons it was concluded that the method as proposed by Levy gives excellent results for thin-skin delta wings, provided that corrections are made for the effect of transverse shear.

Kruszewski, Edwin T.↗
NASA NTRS1961

Heat-Transfer Measurements on the Apexes of Two 60 deg Sweptback Delta Wings (Panel Semiapex Angle of 30 deg) Having 0 deg and 45 deg Dihedral at a Mach Number of 4.95

An experimental investigation was conducted to evaluate the heat-transfer rates at the apex of two 60 degree sweptback delta wings (panel semi-apex angle of 30 degrees) having cylindrical leading edges and 0 degrees and 45 degree positive dihedral. The models tested might correspond to the first several feet of a hypersonic reentry vehicle. The tests were conducted at a Mach number of 4.95 and a stagnation temperature of 400 F. nominal test-section unit Reynolds numbers varied from 2 x 10(exp 6) to 12 x 10(exp 6) per foot. The results of the investigation indicated that the laminar heat-transfer distributions (ratio of local to stagnation-line heating rate) about the models normal to the leading edges were in close agreement with two-dimensional blunt-body theory. The three-dimensional stagnation point heat-transfer rate on the 0 degree dihedral model was in excellent agreement with theory and the stagnation-line heat transfer on the straight portion of the leading edge of both models approached a constant level 12 percent above the theoretical stagnation-line level on an isolated swept infinite cylinder. When the heating rates on the 45 degree dihedral model (planform sweep of 69.3 degree) were compared with those on the 0 degree dihedral model (planform sweep of 60 degrees) at equal angles of attack and equal lifts greater than zero, the stagnation-line heating rates on the 45 degrees dihedral model were, in general, considerably lower as a result of the difference in effective sweeps of the leading edges. On the wing panels inboard from the stagnation lines, the differences in heating were very small. The stagnation-line heat-transfer variation with angle of attack, the shift in stagnation-line location, and the reduction in stagnation-line heat transfer resulting from the increase in effective sweep when positive dihedral is incorporated into a constant-panel 0 degree dihedral wing, all agreed with the results of a theoretical study made of highly swept delta wings with large positive dihedral.

Gunn, Charles R.↗
NASA NTRS1963

A Critical Study of Delta-Wing Configurations for the Supersonic Transport Application

The supersonic transport study currently being made by industry and government anticipates a larger performance advance than that attained by any previous generation of commercial aircraft. Such a jump to relatively high supersonic flight speeds with the accompanying economic and safety requirements of transport aircraft operation accentuates problems of off-design performance and operational flexibility that are difficult to resolve. Ames Research Center, as part of a coordinated NASA program, is attempting, by low-speed, high-speed, flight-simulation, and analytic research to increase the understanding of the critical problems and to provide information needed for their solution, Although these problems are being considered for a number of airplane configurations, special attention at Ames has been directed toward the delta-wing airplane for which serious problems, particularly at low speed, have been foreseen. Because of this emphasis in research effort, consideration will be given here to the delta-wing airplane. It should be understood, however, that many of the problems discussed herein are of equal concern for other configurations.

TRANSPORT AIRCRAFT↗
NASA NTRS1986

Visualization and flow surveys of the leading edge vortex structure on delta wing planforms

In the present experimental investigation of thin delta wing vortex breakdown, for the cases of sweep angles of 70, 75, 80, and 85 deg, and smoke flow visualization/laser light sheet technique is used to obtain cross sectional views of the leading edge vortices as they break down. A combination of lateral and longitudinal cross sectional views furnishes data on the three-dimensional character of the vortex before, during, and after breakdown. Velocity measurements conducted with a laser Doppler anemometer on the 70 deg sweep delta, at 30 deg angle-of-attack, indicate that when breakdown occurs the core flow is transformed from a jet-like to a wake-like flow.

Payne, F. M.↗
NASA NTRS1990

An embedded grid formulation applied to a delta wing

Applications using a three-dimensional embedded grid scheme are made to high angle of attack viscous flow over two bodies: a slender cone using the conical approximation and a 75 deg swept delta wing. The embedded grids are used principally to improve the numerical resolution of the separated vortical flow above the body. Detailed comparisons are made with experimental measurements of the velocity field over the delta wing. The prediction of the maximum steamwise velocity is improved using two levels of embedded grid refinement but is still less than the experimental measurements available from a laser velocimeter.

Krist, Sherrie L.↗