Unsteady Navier-Stokes computations have been performed for simulating transonic flows over wings with oscillating control surfaces using a locally moving grid and a stationary-mismatched zoning scheme. An F-5 wing and a clipped delta wing are chosen for the present study. The computed unsteady pressures and the response characteristics to the control surface motions are compared with experimental data. The results successfully predict main features of the unsteady pressure profiles, such as the double peaks at the shock wave and at the hinge line.

Navier-Stokes computations for oscillating control surfaces

The Transonic Dynamics Tunnel (TDT) at the National Aeronautics and Space Administration's (NASA) Langley Research Center began research operations in early 1960. Since that time, over 600 tests have been conducted, primarily in the discipline of aeroelasticity. This paper presents a bibliography of the publications that contain data from these tests along with other reports that describe the facility, its capabilities, testing techniques, and associated research equipment. The bibliography is divided by subject matter into a number of categories. An index by author's last name is provided.

A Bibliography of Transonic Dynamics Tunnel (TDT) Publications

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 flowfleld and structural flexibility on the wing responses. For the 10-deg ramp motion, a vortex breakdown is observed. The inviscid interaction with the shock wave plays an essential role in the process of the breakdown observed in the present calculation.

Unsteady Shock-Vortex Interaction on a Flexible Delta Wing

Steady and unsteady aerodynamic data were measured on a rectangular wing with a 12 percent thick supercritical airfoil mounted in the NASA Langley Transonic Dynamics Tunnel. The wing was oscillated in pitch to generate the unsteady aerodynamic data. The purpose of the wind-tunnel test was to measure data for use in the development and assessment of transonic analytical codes. The effects on the wing pressure distributions of Mach number, mean angle of attack, and oscillation frequency and amplitude were measured. Results from the newly-developed XTRAN3S program (a non-linear transonic small disturbance code) and from the RHOIV program (a linear lifting surface kernel function code) were compared to measured data for a Mach number of 0.7 and for oscillation frequencies ranging from 0 to 20 Hz. The XTRAN3S steady and unsteady results agreed fairly well with the measured data. The RHOIV unsteady-result agreement was fair but, of course, did not predict shock effects.

Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch

Transonic steady and unsteady aerodynamic data were measured on a large elastic wing in the NASA Langley Transonic Dynamics Tunnel. The wing had a supercritical airfoil shape and a leading-edge sweepback of 28.8 deg. The wing was heavily instrumented to measure both static and dynamic pressures and deflections. A hydraulically driven outboard control surface was oscillated to generate unsteady airloads on the wing. Representative results from the wind-tunnel tests are presented and discussed. An experimentally predicted dynamic wing instability boundary, sensitive to angle of attack, is reported.

Measured unsteady transonic aerodynamic characteristics of an elastic supercritical wing

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.

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

Pressures were measured with Freon as the test medium. Data taken at M = 0.9 is presented for static and oscillatory deflections of the trailing edge control surface and for the wing in pitch. Comparisons of the static measured data are made with results computed using the Bailey-Ballhaus small disturbance code.

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F. W. Cazier

Papers

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

Static and unsteady pressure measurements on a 50 degree clipped delta wing at M = 0.9

Static and unsteady pressure measurements on a 50 degree clipped delta wing at M = 0.9. [conducted in the Langley Transonic Dynamics Tunnel