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Space shuttle abort stage separation wind tunnel tests of orbiter and booster vehicles - booster proximity data, Mach number 3
Space shuttle abort stage separation wind tunnel tests of orbiter and booster vehicles - orbiter proximity data, Mach number 3
Space shuttle abort stage separation wind tunnel tests of orbiter and booster vehicles - booster proximity data, Mach number 2
Space shuttle abort stage separation wind tunnel tests of orbiter and booster vehicles - orbiter proximity data, Mach number 2
The coherent structure in high Reynolds number, Mach 0.6 to 0.8 axisymmetric cold air jets exhausting at atmospheric pressure has been studied. The mean flow and the statistical time averaged turbulence properties have been measured. The coherent structure in the jets was measured using two hot-wire probes and cross correlations. The coherent structure was represented in a wave format. The axial wavelength was measured for several frequencies and was compared to results obtained by other investigators for various Mach and Reynolds numbers. The results indicated that the axial wavelength-frequency relationship was the same for a wide range of Mach and Reynolds numbers.
Lifting surface for reentry vehicles - flutter of sails having delta wing planforms
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Performance of four 16-inch-diameter ram-jet units was determined at free-stream Mach numbers of 0.49 to 1.78 over range of gas total-temperature ratios of 1.0 to 6.1. Time histories of each flight and data on thrust, drag, diffuser efficiency, and combustion are presented. A maximum thrust coefficient of 0.88 and a maximum net acceleration of 5.13 g's were observed for the four units.
Low Mach number supersonic flight data on sonic boom signatures at cutoff Mach number
Results of a free-flight investigation to determine the drag of circular, finite-length cylinders are presented for a Mach number range from about 0.5 to 1.3. Also included are drag results of previous experimental tests of infinite-length cylinders. Drag of circular cylinders at supersonic speeds is largely independent of fineness ratio and Reynolds number; whereas, at subsonic speeds, the drag of finite-length cylinders (fineness rations of about 60 and below) increases as their fineness ratios increase.
Full-potential, Euler, and Navier-Stokes computational fluid dynamics (CFD) codes were evaluated for use in analyzing the flow field about airfoils sections operating at Mach numbers from 0.20 to 0.60 and Reynolds numbers from 500,000 to 2,000,000. The potential code (LBAUER) includes weakly coupled integral boundary layer equations for laminar and turbulent flow with simple transition and separation models. The Navier-Stokes code (ARC2D) uses the thin-layer formulation of the Reynolds-averaged equations with an algebraic turbulence model. The Euler code (ISES) includes strongly coupled integral boundary layer equations and advanced transition and separation calculations with the capability to model laminar separation bubbles and limited zones of turbulent separation. The best experiment/CFD correlation was obtained with the Euler code because its boundary layer equations model the physics of the flow better than the other two codes. An unusual reversal of boundary layer separation with increasing angle of attack, following initial shock formation on the upper surface of the airfoil, was found in the experiment data. This phenomenon was not predicted by the CFD codes evaluated.
An experimental program was conducted to survey the lee side vortex flow field about an ogive-cylinder-frustum-cylinder at angles of attack to 25 degrees for two Reynolds numbers at Mach number 0.8, and one Reynolds number at Mach number 1.96. The data were obtained using miniature 5-port conical pressure probes calibrated for angle of attack and roll angle over a Mach number range of 0.6 to 3.0. The results are presented here as local flow field properties and circulation strengths for various body stations.
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An investigation was made in the Langley 8-foot transonic tunnel and the Langley Unitary Plan wind tunnel, over a Mach number range of 0.6 to 2.16, to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic-cruise fighter. The configuration, which is designed for efficient cruise at Mach number 1.8, is a twin-engine tailless arrow-wing concept with a single rectangular inlet beneath the fuselage and outboard vertical tails and ventral fins. It had untrimmed values of lift-drage ratio ranging from 10 at subsonic speeds to 6.4 at the design Mach number. The configuration was statically stable both longitudinally and laterally.
Wind tunnel tests of a proposed HL-10 lifting body vehicle were conducted to determine the subsonic and transonic aerodynamic characteristics. The conditions under which the tests were conducted are described. The tests indicate that the configuration has slightly positive damping in pitch except at higher angles of attack at Mach numbers of 0.8, 0.9, and 1.0. At supersonic speeds, the configuration has positive damping in pitch for all test conditions. At subsonic and transonic speed, the configuration has positive damping and positive stability in yaw for all test conditions.
Mach number and pressure measurements on afterbody of blunt-nose model in free flight at hypersonic speeds
A diamond wing and body combination was designed to have an area distribution which would result in near optimum zero-lift wave-drag coefficients at a Mach number of 1.00, and decreasing wave-drag coefficient with increasing Mach number up to near sonic leading-edge conditions for the wing. The airfoil section were computed by varying their shape along with the body radii (blending process) to match the selected area distribution and the given plan form. The exposed wing section had an average maximum thickness of about 3 percent of the local chords, and the maximum thickness of the center-line chord was 5.49 percent. The wing had an aspect ratio of 2 and a leading-edge sweep of 45 deg. Test data were obtained throughout the Mach number range from 0.20 to 3.50 at Reynolds numbers based on the mean aerodynamic chord of roughly 6,000,000 to 9,000,000. The zero-lift wave-drag coefficients of the diamond model satisfied the design objectives and were equal to the low values for the Mach number 1.00 equivalent body up to the limit of the transonic tests. From the peak drag coefficient near M = 1.00 there was a gradual decrease in wave-drag coefficient up to M = 1.20. Above sonic leading-edge conditions of the wing there was a rise in the wave-drag coefficient which was attributed in part to the body contouring as well as to the wing geometry. The diamond model had good lift characteristics, in spite of the prediction from low-aspect-ratio theory that the rear half of the diamond wing would carry little lift. The experimental lift-curve slope obtained at supersonic speeds were equal to or greater than the values predicted by linear theory. Similarly the other basic aerodynamic parameters, aerodynamic center position, and maximum lift-drag ratios were satisfactorily predicted at supersonic speeds.