基于HIFiRE-2超燃发动机内流道的激波边界层干扰分析

Analysis of Shock Wave Boundary Layer Interactions Based on Internal Flowpath of HIFiRE-2 Scramjet

为了研究高超声速流激波边界层干扰特性,选取HIFi RE-2(The Hypersonic International Flight Research Experimentation2)项目的高超声速流道为研究对象,采用k-ωSST模型在无燃油工况下模拟计算地面试验过程,所得计算结果与试验结果接近。在此基础上,分析激波边界层干扰过程、流动分离现象及入口马赫数对气动热影响。结果表明:随着入口马赫数增大,激波角变小,激波强度提高,在尾喷管中激波反射次数减少;随着入口速度增大,边界层分离区范围变小,回流区的位置逐渐向下游移动;加入气动耗散项后,流场的温度有一定升高,最大温升约为50 K。

In order to investigate the characteristics of shock wave boundary layer interactions of hypersonic flow,the hypersonic flowpath of HIFiRE-2 project is selected as the research object.The k-ωSST model was used to simulate and calculate the ground test process under the condition of no fuel and the calculated results were close to the test results.On this basis,the influence of the process of shock wave boundary layer interactions,flow separation and the influence of inlet Mach number on the aerodynamic heat were analyzed.The results show that with the increase of the inlet Mach number,the shock intensity increases,shock angle and the number of shock reflection decreases in the exhaust nozzle.With the inlet velocity increases,the boundary layer separation zone becomes smaller and the position of the recirculation zone gradually moves downstream.After adding the aerodynamic dissipation term,the maximum temperature of the flow field is increased by 50K.