中文摘要：

为了研究高超声速咽式进气道在非设计迎角以及低马赫数下的起动性能,利用流线追踪生成了设计马赫数Ma=7,具有8-7无粘基本流场（即俯仰平面内的斜激波由和自由来流呈8°夹角的斜压缩面产生;偏航平面内的斜激波由和自由来流呈7°夹角的斜压缩面产生）的咽式进气道,并对边界层修正前后的两种咽式进气道进行了数值模拟和高超声速风洞实验。实验观测和记录了各个来流条件下进气道模型唇口的激波系结构,测量了沿进气道模型上下壁面中心线从气流进口到出口的沿程静压分布。结果表明：迎角的增大和来流马赫数的减小都会对进气道的起动性能造成不利的影响,通过对咽式进气道进行边界层修正,可以提高进气道的总压恢复系数,减小内收缩比,从而扩宽进气道起动的马赫数以及迎角范围,对进气道设计有着积极的作用。

英文摘要：

The starting performance of hypersonic jaws inlet in both design and off-design conditions are studied in this paper.Using streamline tracing method,ajaws inlet is designed with 8-7 planar shock wave（the ramp in pitch plane is inclined at 8° to the free stream and the one is 7° in yaw plane,yielding planar shocks）in the design condition of Mach 7.And then,the inlet with and without the boundary layer correction is investigated in Mach 5,6 and 7 respectively by numerical simulations and wind tunnel experiments.With static pressure distribution measured in experiments,the impacts of free stream Mach number and angle of attack on flow distribution and starting performance of jaws inlet are analyzed and compared.The results indicate that both increasing angle of attack and decreasing free stream Mach number will show negative effects on starting performance of the hypersonic jaws inlet.In addition,boundary layer correction improves the total pressure recovery coefficient of jaws inlet and reduces internal contraction ratio,and the working range of the inlet is enlarged.