高超声速进气道设计方法研究

Research on Design of Hypersonic Inlet

针对超燃冲压发动机二维混压式前体/进气道采用斜激波压缩方法设计总压损失较大的问题,采用等熵压缩的方法对进气道下壁进行了流线化改进设计。利用计算流体力学软件Fluent,采用耦合隐式求解器,标准k-ω湍流模型,对进气道流场进行数值模拟,分析比较了采用斜激波压缩方法和等熵压缩方法设计的两种进气道的设计点性能和非设计点性能。研究了燃烧室压力升高对进气道/隔离段流场和起动性能的影响。研究表明,与斜激波压缩进气道相比,采用等熵压缩方法设计的曲线上壁进气道总压恢复系数、流量系数得到明显提高,但其冲压比与前者相比有所降低;进气道进入不起动后,其捕获流量和总压恢复系数急剧下降。

Aiming at the problem of high total pressure loss in a two-dimensional scramjet forebody/inlet designed by the oblique-shock compression method,the streamlined improvement of inlet bottom-wall was designed by isentropic compression method.The flow fields within hypersonic inlet has been simulated using the Fluent CFD code.The current version of this code employs coupled implicit solver and standard k-ω turbulence model.The performance of two inlets（respectively designed by oblique-shock compression method and isentropic compression method） at design point and off-design point are compared and analyzed.The effect of back pressure raise in combustor on inlet/isolator flow field and start performance is simulated.The research indicates that the total pressure recovery coefficient and flow coefficient of the hypersonic inlet designed by isentropic compression method are higher than that of the inlet designed by oblique-shock compression method.But pressure ratio of the first inlet is lower than that of the second.For an unstarted inlet,the captured mass flow and total pressure recovery coefficient will decrease quickly.