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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Leeward flow over delta wings at supersonic speeds

A survey was made of the parameters affecting the development of the leeward symmetric separated flow over slender delta wings immersed in a supersonic stream. The parameters included Mach number, Reynolds number, angle of attack, leading-edge sweep angle, and body cross-sectional shape, such that subsonic and supersonic leading-edge flows are encountered. It was seen that the boundaries between the various flow regimes existing about the leeward surface may conveniently be represented on a diagram with the components of angle of attack and Mach number normal to the leading edge as governing parameters.

Szodruch, J. G.↗
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 NTRS1960

Some Divergence Characteristics of Low-Aspect-Ratio Wings at Transonic and Supersonic Speeds

The problem of chordwise, or camber, divergence at transonic and supersonic speeds is treated with primary emphasis on slender delta wings having a cantilever support at the trailing edge. Experimental and analytical results are presented for four wing models having apex half-angles of 5 deg, 10 deg, 15 deg, and 20 deg. A Mach number range from 0.8 to 7.3 is covered. The analytical results include calculations based on small-aspect-ratio theory, lifting-surface theory, and strip theory. A closed-form solution of the equilibrium equation is given, which is based on low-aspect-ratio theory but which applies only to certain stiffness distributions. Also presented is an iterative procedure for use with other aerodynamic theories and with arbitrary stiffness distribution.

Woolston, Donald S.↗
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 NTRS1996

Natural Rolling Responses of a Delta Wing in Transonic and Subsonic Flows

The unsteady, three-dimensional, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the natural rolling response of slender delta wings of zero thickness at moderate to high angles of attack, to transonic and subsonic flows. The governing equations of fluid flow and dynamics of the present multi-disciplinary problem are solved using the time-accurate solution of the NS equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The main focus is to analyze the effect of Mach number and angle of attack on the leading edge vortices and their breakdown, the resultant rolling motion, and overall aerodynamic response of the wing. Three cases demonstrate the natural response of a 65 deg swept, cropped delta wing in a transonic flow with breakdown of the leading edge vortices and an 80 deg swept delta wing in a subsonic flow undergoing either damped or self-excited limit-cycle rolling oscillations as a function of angle of attack. Comparisons with an experimental investigation completes this study, validating the analysis and illustrating the complex details afforded by computational investigations.

Menzies, Margaret A.↗
NASA NTRS1982

Lateral aerodynamics of delta wings with leading edge separation

An unsteady vortex lattice method is presented for the calculation of the aerodynamic forces acting on lifting surfaces undergoing complex three dimensional motion. For the present case the nonsymmetric motion of a slender delta wing was considered and the resulting lateral characteristics were calculated. The flow separation line was specified along the wing leading edge and the emanating vortex sheet shape and rollup was then calculated. Numerical results are presented for the combined high angle of attack and side slip condition and for the wing constant roll and coning motions.

Katz, J.↗
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 NTRS1992

Navier-Stokes dynamics and aeroelastic computations for vortical flows, buffet and aeroelastic applications

The accomplishments achieved during the period include conference and proceedings publications, journal papers, and abstracts which are either published, accepted for publication or under review. Conference presentations and NASA highlight publications are also included. Two of the conference proceedings publications are attached along with a Ph.D. dissertation abstract and table of contents. In the first publication, computational simulation of three-dimensional flows around a delta wing undergoing rock and roll-divergence motions is presented. In the second publication, the unsteady Euler equations and the Euler equations of rigid body motion, both written in the moving frame of reference, are sequetially solved to simulate the limit-cycle rock motion of slender delta wings. In the dissertation abstract, unsteady flows around rigid or flexible delta wings with and without oscillating leading-edge flaps are considered.

Kandil, Osama A.↗
NASA NTRS1986

A direct and inverse boundary layer method for subsonic flow over delta wings

A new inverse boundary layer method is developed and applied to incompressible flows with laminar separation and reattachment. Test cases for two dimensional flows are computed and the results are compared with those of other inverse methods. One advantage of the present method is that the calculation of the inviscid velocities may be determined at each marching step without having to iterate. The inverse method was incorporated with the direct method to calculate the incompressible, conical flow over a slender delta wing at incidence. The location of the secondary separation line on the leeward surface of the wing is determined and compared with experiment for a unit aspect ratio wing at 20.5 deg incidence. The viscous flow in the separated region was calculated using prescribed skin friction coefficients.

Woodson, S. H.↗
NASA NTRS1980

Aerodynamic and inlet flow characteristics of several hypersonic airbreathing missile concepts

Four conceptual hypersonic missile configurations were examined experimentally and theoretically. Two of the concepts employed twin module bottom-mounted engines and two were designed for upper surface inlets or engines with the intent of reducing the vehicle observables. The tests were conducted at Mach 6 and Reynolds numbers of 6 to 7.5 x 10 to the 6th per foot. Flow field surveys in the vicinity of the engine inlet were made on all configurations and force and moment tests were conducted on three of the vehicles. Stability and control characteristics of the bottom-mounted engine configurations which incorporated slender, low wings were dominated by strong vortices that promoted severe pitchup tendencies. The shock layer and flow quality in the vicinity of the bottom-mounted engine inlets were dependent on nose shape. The spatula-like upper surface engine concept demonstrated good performance and had uniform flow entering the engine inlet, while the upper surface inlet concept with a highly swept forebody incurred large gradients due to interactions with leading edge shocks.

Dillon, J. L.↗