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Pressures were simultaneously measured in the variable-density tunnel at 54 orifices distributed over the midspan section of 5 by 30 inch rectangular model of the NACA 4412 airfoil at 17 angles of attack ranging from -20 degrees to 30 degrees at a Reynolds number of approximately 3,000,000. Accurate data were thus obtained for studying the deviations of the results of potential-flow theory from measured results. The results of the analysis and a discussion of the experimental technique are presented.
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An investigation was conducted in the NASA/LaRC Low-Turbulence Pressure Tunnel on a 0.03614-scale orbiter model of a 089B configuration with a 139B configuration nose forward of F.S. 500. The tailcone was the TC sub 4 design and was instrumented with eighty-nine pressure orifices. Control surfaces were deflected and three wind tunnel mounting techniques were investigated over an angle-of-attack range from -2 deg to a maximum of 18 deg. In order to determine the sensitivity of the tailcone to changes in Reynolds number, most of the test was made at a Mach number of 0.20 over a Reynolds number range of 2.0 to 10 million per foot. A few runs were made at a Mach number of 0.30 at Reynolds numbers of 4.0, 6.0, and 8 million per foot.
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Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at a Mach number of 6 and a free-stream ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of selected tunnel runs for flow visualization of the entrance plane and cowl region. Although these movies could not show the internal flow, the effect of the internal flow on the external flow was evident by way of spillage. The purpose is to provide a preliminary data release for the investigation. The models, facility, and testing methods are described, and the test matrix and a tabulation of tunnel runs are provided. Line plots highlighting the stated parametric effects and a representative set of schlieren photographs are presented without analysis.
Gas kinetic pressure profile and mass density of propagating current sheet in argon pinch discharge, using piezoelectric transducer
Effects of elevon-induced three-dimensional shock-wave turbulent boundary-layer interactions on hypersonic aircraft surfaces are analyzed. Detailed surface pressure and heating rate distributions obtained on wing-elevon-fuselage models representative of aft portions of hypersonic aircraft are compared with analytical and experimental results from other sources. Examples are presented that may be used to evaluate the adequacy of current theoretical methods for estimating the effects of three-dimensional shock-wave turbulent boundary-layer interactions on hypersonic aircraft surfaces.
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