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

为了研究高压气体管道系统、吸气式发动机和空间飞机内、外流动的超音速流动现象，利用超音速风洞(压力真空型，Mach4.0和2.0型)对矩形管道内激波/边界层干扰(拟激波)进行了实验研究。用纹影照相技术显示PSW。用粒子成像测速仪(PIV)测量了PSW内的详细速度分布。在前激波作用下，上、下壁面上的边界层分离比另一个壁面上的严重，形成了不对称的流场。此外，为了弄清冲压发动机超音速进气道的特性，采用真空式风洞和彩色纹影摄影、激光多普勒测速、水银测压等测量方法，对马赫数1.9设计的小型外压缩矩形进气道模型的绕流和流过流动结构进行了实验研究。一个关键的操作，其中斜坡引起的冲击附着在罩的前缘，实现了。在进气口的下缘安装了一个流动塞子，通过激波-边界层相互作用引起的激波列实现了单调压缩。

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.