进口斜激波、膨胀波干扰下等直隔离段内的激波串特性

Characteristics of Shock-train in a Straight Isolator with Interference of Incident Shock Waves and Corner Expansion Waves

超燃冲压发动机的隔离段在实际工作中会受到进气道唇罩激波及肩部膨胀扇的显著干扰,本文针对这一特定问题进行了专门研究。提出了唇罩入射激波及肩部膨胀扇的模拟方法,并利用德国Achen的风洞试验对其进行了检验,而后以此研究了入射激波及肩部膨胀扇干扰下隔离段内激波串的基本形态,并分析了出口反压和激波入射位置的影响。仿真结果表明:当激波串在隔离段内不断前移时,受唇罩入射激波及其反射激波的干扰,其高速核心区交替地偏向上下壁面;与无激波入射的情况相比,此时激波串的耐反压能力显著降低,且入射点位置越高,降低幅度越大,管道内的沿程静压分布规律与Waltrup经验公式偏离程度也越来越大。该文结果可为进气道/隔离段的一体化设计提供依据。

The operation of the isolator of a scramjet is substantially affected by cowl induced shocks and corner expansion waves. The current study attempts to reveal the influence of these incident waves on the flow patterns and pressure recovery characteristics of the shock-train in the isolator. First, a simulation method for the incident shock and the corner expansion waves is advanced and validated with Achen wind tunnel experimental results in Germany. Then, the basic flow patterns of the shock-train with the influence of the incident shock waves are analyzed. The effect of the back pressure ratio and incident position of the cowl induced shock is also obtained. The results indicate that the supersonic core flow region of the shock-train deflects alternatively to the top surface and the bottom surface of the isolator with its upstream propagation because of the influence of the incident shock and reflected shocks. As compared with the case without incident shocks, the pressure gain across the shock-train is substantially decreased due to the incident shock waves. The pressure distribution deviates from the Waltrup empirical formula continuously with the rise of the incident point of the cowl induced shocks. The results obtained may provide a design guide for the integration of the inlet and the isolator.