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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。
In order to study the starting performance of a hypersonic phantom inlet at a non-design angle of attack and a low Mach number, the design Mach number Ma = 7 was generated using the streamline tracing, with an 8-7 non-stick basic flow field Of the oblique shock generated by the free flow and the angle of 8 ° oblique compression surface generated; the yaw plane of the oblique shock wave from the free flow and the angle of 7 ° angled compression surface generated) pharyngeal inlet The numerical simulation and hypersonic wind tunnel experiments on two kinds of phantom inlet before and after the boundary layer correction are carried out. The shock system structure of the lip of the inlet model was experimentally observed and recorded. The static pressure distribution along the center of the upper and lower wall of the inlet model was measured along the flow path from the inlet to the outlet. The results show that the increase of the angle of attack and the decrease of the Mach number will have an adverse effect on the starting performance of the intake passage. By modifying the boundary layer of the phantom intake port, the total pressure recovery coefficient , Reduce the internal contraction ratio, thus widening the inlet Mach number and angle of attack range, the intake port design has a positive effect.