An Experimental Parametric Study of Geometric, Reynolds Number, and Ratio of Specific Heats Effects in Three-Dimensional Sidewall Compression Scramjet Inlets at Mach 6
Since mission profiles for airbreathing hypersonic vehicles such as the National Aero-Space Plane include single-stage-to-orbit requirements, real gas effects may become important with respect to engine performance. The effects of the decrease in the ratio of specific heats have been investigated in generic three-dimensional sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane (where gamma=1.22) and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air (where gamma=1.4). In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.
INVESTIGATION OF HYPERSONIC INLET SHOCK-WAVE BOUNDARY LAYER INTERACTION. PART 2. CONTINUOUS FLOW TEST AND ANALYSES
Abstract : The computational procedure developed in Phase I for estimating laminar boundary layer properties through impinging shock induced laminar boundary layer interactions was modified to account for entropy change of the inviscid streamlines in that isentropic assumptions were used in the initial formulation. Perturbation studies on the effects of wall shear on interaction length have identified the importance of the profile shape factor in such calculations. A Users' Manual for the laminar computational program is contained. Three incident shock models were designed for use in the experimental program conducted in the Ames 3.5 Foot Hypersonic Tunnel. Aided by analysis of the experimental data, a simplified flow model was formulated for the case of a two-dimensional turbulent boundary layer incident shock interaction.
Mach 4 testing of scramjet inlet models
Six scramjet inlet models were tested in a Mach 4 wind tunnel. Wall and pitot pressure were measured and schlieren photographs were taken. Parameters of these models are side-plate sweep angle, contraction ratio, and cowl geometry. The shock pattern inside one of the models, as shown by schlieren photographs, coincides with the calculations. Both mass-capture ratio and total pressure recovery are 50-70%. There seems to be a better sweep angle and a better cowl length for maximum total pressure recovery.