Characterization and Control of Shock Waves and Compressible Turbulent Boundary Layers through Supersonic Rectangular Ducts

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To Investigate the supersonic flow phenomena related to internal and external flows of high pressure gas pipeline systems, air breathing engines and space planes, an experimental study was carried out on shock wave/boundary layer interaction (pseudo shock wave, or PSW) in a rectangular duct, using a supersonic wind tunnel (pressure vacuum type, Mach4.0 and 2.0). A PSW was visualized by schlieren photography. Detailed distribution of velocity in the PSW was measured by particle imaging velocimetry (PIV). Boundary layer separation after the front shock wave occurred more severely on one of the upper and lower walls than on the other, producing an asymmetric flow field. Deceleration of the core flow around the centerline of the duct was small.Additionally, in order to clarify characteristics of supersonic intakes for ramjet engines, flow structure around and through a small external-compression rectangular intake model with double ramps designed for Mach number 1.9 was investigated experimentally using a vacuum-type wind tunnel and a set of measurement methods composed of colored schlieren photography, laser Doppler velocimetry, and mercury manometry. A critical operation, where shocks induced by ramps attach to the leading edge of the cowl, was attained. A flow plug installed on the downward edge of the intake enabled monotonous compression through a shock train caused by shock-boundary layer interactions.