高超声速双模块内转式进气道的流动特性研究-Part I： 设计状态

Flowfield of Hypersonic Bimodule Inward-Turning Inlet-Part I: Design Point

关键词:  超燃冲压发动机  进气道  激波  边界层干扰  流向涡  基准流场

Key words:  Scramjet  Inlet  Shock wave  Boundary layer interaction  Streamwise vortices  Basic flowfield

摘要:

为了进一步研究几何非对称的高超声速内转式进气道流动特性，通过特征线法设计了基于多道激波加等熵压缩波的基准流场，在此基础上通过流线追踪法设计了双模块内转式进气道。通过数值仿真和风洞试验相结合的方法，获得了内转式进气道的内外流特性。研究结果表明：在内转进气道最大半径对应的角区位置存在大量的边界层堆积，受第二道激波/边界层干扰，在激波根部卷起锥形旋涡。在内转式进气道内部，唇罩激波和管道边界层干扰显著，管道内存在自唇罩指向压缩面的强周向压力梯度，从而诱导管道内边界层均往一处汇聚，卷起大尺度流向涡。仿真和试验结果表明，在来流马赫数5.74，攻角0°状态下，进气道气动性能优良，出口总压恢复系数系数达到0.58，最大抗反压为112倍。

Abstract:

In order to further study the flow characteristics of the geometrically asymmetric hypersonic internal-turning inlet， a basic flowfield contained multiple shocks and isentropic compression waves are designed by the method of characteristics， bi-module inward-turning inlets are then designed. Through the combination of numerical simulation and wind tunnel test， the internal and external flow characteristics of the internal turning inlet are obtained. The research results show that there is a large amount of boundary layer accumulated at the corner area corresponding to the maximum radius of the inner turning inlet， which is driven by the second shock wave/boundary layer， rolling up a conical-shaped vortex at the foot of the shock wave. In the internal duct， the lip shock / boundary layer interaction is fairly strong. A circumferential pressure gradient pointed from the cowl lip to the compression surface is formed， which induces the internal boundary layer to converge and roll up large-scale vortices. Simulation and test results show that under the condition of incoming flow Mach number of 5.74 and angle of attack of 0°， the inlet aerodynamic performance is excellent， among which the outlet total pressure recovery coefficient reaches 0.58， and the maximum sustainable back-pressure is 112 times.