超声速进气道压力估算方法及验证

Estimation Method and Verification of Pressure in Supersonic Inlet

进气道载荷的预示和限制是超声速飞行器设计中的关键问题。以典型颌下进气超声速飞行器为研究对象,对其进气道流场进行数值仿真,研究不同马赫数、攻角、侧滑角及余气系数条件下的进气道压力特性;针对进气道压力工程估算及设计需求,使用无量纲和解耦的方法,对进气道压力经验公式进行拟合;反算飞行试验中的进气道压力,并与测量结果进行对比。结果表明:进气道压力随马赫数增大而增大,随余气系数增大而减小;正常工作包线内,较小的攻角、侧滑角对进气道压力影响不明显;进气道压力经验公式计算值与飞行试验测量值符合较好,具有较高的精度。

Prediction and limitation of inlet pressure are important for supersonic vehicles.Numerical simulation on high speed inlet flow is conducted,and pressure characteristics with different Mach numbers,attack angles,yaw angles and excess air coefficients is analyzed.Based on dimensionless and decoupled method,empirical formula for high speed inlet pressure is fitted and verified by flight results.The results show that,with the Mach number increasing and excess air coefficient decreasing,the pressure in inlet increases.Attack angle and yaw angle have little influence on inlet pressure in normal work envelope.The inlet pressure estimated by empirical formula is in good agreement with measured data in flight tests,which means that the formula established in this paper has a high precision.