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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Application of the DSMC method to hypersonic flow about a delta wing

The direct simulation Monte Carlo method was applied to a hypersonic rarefied flow about a delta wing at angle of incidence. Three-dimensional calculations were performed using a single-species gas model that accounts for energy exchange with the internal modes. Computed results were compared with the available experimental data obtained in a nitrogen wind-tunnel experiment. Discrepancies were shown to exist between the experimental and calculated results. These differences may be caused by the inadequately defined freestream conditions.

Celenligil, M. C.↗
NASA NTRS1972

Pitot-pressure distributions of the flow field of a delta-wing orbiter

Pitot pressure distributions of the flow field of a 0.0075-scale model of a typical delta wing shuttle orbiter are presented. Results are given for the windward and leeward sides on centerline in the angle-of-attack plane from wind tunnel tests conducted in air. Distributions are shown for three axial stations X/L = .35, .60, and .98 and for angles of attack from 0 to 60 deg. The tests were made at a Mach number of 7.4 and for Reynolds numbers based on body length from 1,500,000 to 9,000,000. The windward distributions at the two survey stations forward of the body boat tail demonstrate the compressive aspects of the flow from the shock wave to the body. Conversely, the distributions at the aft station display an expansion of the flow that is attributed to body boat tail. On the lee side, results are given at low angles of attack that illustrate the complicating aspects of the canopy on the flow field, while results are given to show the effects of flow separation at high angles of attack.

Cleary, J. W.↗
NASA NTRS1981

A vortex-lattice method for calculating longitudinal dynamic stability derivatives of oscillating delta wings

A nonsteady vortex-lattice method is introduced for predicting the dynamic stability derivatives of a delta wing undergoing an oscillatory motion. The analysis is applied to several types of small oscillations in pitch. The angle of attack varied between + or - 1 deg, with the mean held at 0 deg when the flow was assumed to be attached and between + or - 1 deg and the mean held at 15 deg when both leading-edge separation and wake roll-up were included. The computed results for damping in pitch are compared with several other methods and with experiments, and are found to be consistent and in good agreement.

Levin, D.↗
NASA NTRS1982

Analytical study of vortex flaps on highly swept delta wings

This paper highlights some current results from ongoing analytical studies of vortex flaps on highly swept delta wings. A brief discussion of the vortex flow analysis tools is given along with comparisons of the theories to vortex flap force and pressure data. Theoretical trends in surface pressure distribution for both angle-of-attack variation and flap deflection are correctly predicted by Free Vortex Sheet theory. Also shown are some interesting calculations for attached-flow and vortex-flow flap hinge moments that indicate flaps utilizing vortex flow may generate less hinge moment than attached flow flaps. Finally, trailing-edge flap effects on leading-edge flap thrust potential are investigated and theory-experiment comparisons made.

Frink, N. T.↗
NASA NTRS1984

Delta wings with shock-free cross flow

In order to have a high level of maneuverability, supersonic delta wings should have a cross flow that is free of embedded shock waves. The conical cross flow sonic surface differs from that of plane transonic flow in many aspects. Well-known properties such as the monotone law are not true for conical cross flow sonic surfaces. By using a local analysis of the cross flow sonic line, relevant conditions for smooth cross flow are obtained. A technique to artificially construct a smooth sonic surface and an efficient numerical method to calculate the flow field are used to obtain cones with smooth cross flow.

Sritharan, S. S.↗
NASA NTRS1985

Delta wings with shock-free cross flow

In order to have a high level of maneuverability, supersonic delta wings should have a cross flow that is free of embedded shock waves. The conical cross flow sonic surface differs from that of plane transonic flow in many aspects. Well-known properties such as the monotone law are not true for conical cross flow sonic surfaces. By using a local analysis of the cross flow sonic line, relevant conditions for smooth cross flow are obtained. A technique to artificially construct a smooth sonic surface and an efficient numerical method to calculate the flow field are used to obtain cones with smooth cross flow.

Sritharan, S. S.↗
NASA NTRS1986

Steady and unsteady transonic pressure measurements on a clipped delta wing for pitching and control-surface oscillations

Steady and unsteady pressures were measured on a clipped delta wing with a 6-percent circular-arc airfoil section and a leading-edge sweep angle of 50.40 deg. The model was oscillated in pitch and had an oscillating trailing-edge control surface. Measurements were concentrated over a Mach number range from 0.88 to 0.94; less extensive measurements were made at Mach numbers of 0.40, 0.96, and 1.12. The Reynolds number based on mean chord was approximately 10 x 10 to the 6th power. The interaction of wing or control-surface deflection with the formation of shock waves and with a leading-edge vortex generated complex pressure distributions that were sensitive to frequency and to small changes in Mach number at transonic speeds.

Hess, Robert W.↗
NASA NTRS1988

Detailed flowfield measurements over a 75 deg swept delta wing for code validation

Selected results from an experimental investigation documenting the flowfield over 75 deg swept delta wing at an angle-of-attack of 20.5 deg are presented. Results obtained in the investigation include surface flow visualization, off-body flow visualization, and detailed flowfield surveys for various Reynolds numbers. Flowfield surveys at Reynolds numbers of 0.5, 1.0 and 1.5 million were conducted with both a pitot pressure probe and a 5-hole pressure probe; and 3-component laser Doppler velocimeter surveys were conducted at a Reynolds number of 1.0 million. The pitot pressure surveys were obtained at 5 longitudinal stations, the 5-hole probe surveys were obtained at 3 longitudinal stations and the laser Doppler velocimeter surveys were obtained at one station. The accuracy of each instrumentation system is discussed, as well as, discrepancies in the calculation of vorticity using various algorithms.

Kjelgaard, Scott O.↗
NASA NTRS1990

Unsteady surface pressure distributions on a delta wing undergoing large amplitude pitching motions

Wind tunnel experiments were performed on a 70-deg-sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Pressure taps were placed from 35-90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as functions of Reynolds number and pitch rate. The surface pressure distribution was seen to vary at the same frequency as the pitching frequency, though distortion due to the vortex breakdown was observed. Comparing the upstroke (angle of attack increasing) and downstroke (angle of attack decreasing) pressures for a specific angle of attack, a time lag in the pressure distribution was observed. The downstroke pressures were slightly larger at the forward chord locations. Vortex breakdown was seen to have the most significant effect at the 40-45-percent chord location, where an increase in local pressure was apparent, as well as a distortion of the periodic pressure fluctuation.

Thompson, S. A.↗
NASA NTRS1990

Delta wing surface pressures for high angle of attack maneuvers

A wind tunnel experiment was performed on a delta wing with a leading edge sweep of 70 deg. Unsteady pressures were measured on the suction surface of the wing as it was oscillated from 0-30 deg, and 2-60 deg angle-of-attack. Pressure coefficients were measured at different surface locations for two pitch rates. Static pressure measurements were also obtained for comparison with the dynamic results. The Reynolds number was fixed at 420,000, based on the centerline chord length. Pressure measurements were made from 35-90 percent of the chord, along a ray from the apex at 60 percent of the local semispan. Spanwise measurements were also made at a constant chord location, x/c = 75 percent from the apex. The unsteady pressure data over the 0-30 deg angle-of-attack range showed pressures fluctuating in phase with model motion, and little overshoot from the static values. The pressures for the large amplitude motion showed large overshoots from the static values. In addition, during the high angle-of-attack portion of the motion, the upstroke (angle of attack increasing) pressure coefficients were typically much lower than the downstroke values. For the lower pitch rate, there was little difference between upstroke and downstroke pressures at the low angles of attack.

Thompson, S. A.↗
NASA NTRS1991

Unsteady shock-vortex interaction on a flexible delta wing

Unsteady Navier-Stokes computations have been carried out for simulating transonic flows over a clipped delta wing undergoing oscillatory and ramp motions, including flexibility. The implicit upwind algorithm has been validated by comparing the solutions with experimental data for the oscillatory pitching motion cases. The numerical and experimental results agree well at moderate angles of attack, where a leading-edge vortex develops. The ramp motion cases have demonstrated the effects of unsteadiness of the flow field and structural flexibility on the wing responses. For the 10 deg ramp motion, a vortex breakdown is observed. The interaction with the shock wave plays an essential role in the process of the breakdown observed in the present calculation.

Obayashi, Shigeru↗
NASA NTRS1991

Hypersonic rarefied flow about a delta wing - Direct simulation and comparison with experiment

Three-dimensional simulations of hypersonic rarefied flow about a delta wing are made using the direct simulation Monte Carlo (DSMC) method of Bird, and the results of the computations are compared with recent experimental data obtained in a vacuum wind tunnel at the DLR in Gottingen, Germany. The present study considers Mach 8.89 nitrogen flow for a range of conditions that include Knudsen numbers of 0.016 to 3.505 for an incidence angle of 30 deg, and angles of incidence of 15 to 60 deg for a constant Knudsen number of 0.389. The calculations provide details concerning the flowfield structure and surface quantities. Comparisons between the calculations and the available experimental measurements are made for aerodynamic and overall heat-transfer coefficients and recovery temperature. The agreement between the measured and calculated data are very good, well within the estimated measurement uncertainty. Comparisons are also made with modified Newtonian and free-molecule theories.

Celenligil, M. C.↗
NASA NTRS1993

Measurements of circulation and vorticity in the leading-edge vortex of a delta wing

The present cross-wire measurements of the flowfield above a 75-deg flat-plate delta wing have yielded distributions of velocity, axial vorticity, and circulation which exhibit strongly conical behavior upstream of the breakdown region and away from the apex and trailing edge regions. The conical character of these properties is most pronounced over the central portion of the planform, away from the apex and trailing-edge regions. This behavior is further documented by the roughly linear increase in vortex strength with distance from the apex over the forward area of the wing.

Visser, K. D.↗
NASA NTRS1993

Shock-vortex interaction over a 65-degree delta wing in transonic flow

Transonic flow over a 65-deg swept-back, sharp-edged, cropped delta wing is investigated computationally using the time-accurate solution of the unsteady, compressible, full Navier-Stokes equations with an implicit, upwind, flux-difference splitting, finite-volume scheme. Coarse and fine O-H grids are used to obtain the solution. The grid consists of 125 x 85 x 84 points in the wrap-around, normal, and axial directions, respectively. The results are presented for an angle of attack of 20 deg Mach number of 0.85 and Reynolds number of 3.23 x 10 exp 6. With the fine grid, the results show that a system of shocks has been captured over the upper wing surface and that the leading-edge vortex core experiences an unsteady supersonic vortex breakdown after passing through a spanwise shock near the wing trailing edge. The computed results at a certain time are in good agreement with the experimental data. Topological aspects of the vortex breakdown flowfield are also presented and discussed.

Kandil, Osama A.↗
NASA NTRS2003

Effects of Various Fillet Shapes on a 76/40 Double Delta Wing from Mach 0.18 to 0.7

The effects of linear, diamond, and parabolic fillets on a double delta wing were investigated in the NASA Langley 7 x 10 ft High Speed Tunnel from Mach 0.18 to 0.7 and angles of attack from 4 deg. to 42 deg. Force and moment, pneumatic pressures, pressure sensitive paint, and vapor screen flow visualization measurements were used to characterize the flow field and to determine longitudinal forces and moments. The fillets increased lift coefficient and reduced induced drag without significantly affecting pitching moment. Pressure sensitive paint showed the increase in lift is caused by an increase in suction and broadening of the vortex suction footprint. Vapor screen results showed the mixing and coalescing of the strake fillet and wing vortices causes the footprint to broaden.

Gonzalez, Hugo A.↗