高超声速下表面凸起干扰气动热实验研究

Experimental Investigation on Hypersonic Interference Heating Around Surface Protuberance

对高超声速飞行器表面凸起附近的气流流动和气动加热开展了实验研究和分析。实验在高超声速炮风洞中进行,来流马赫数为8.2、单位雷诺数为9.35×106 m-1。利用薄膜传热测量方法进行了凸起几何形状和边界层状态对干扰流动加热的影响评估。利用流油图谱和纹影摄像法得到了凸起周围的流动特征:若凸起上游边界层未分离,最大峰值热流发生在凸起侧方附近处;若凸起上游边界层完全分离,最大峰值热流通常发生在凸起的上游表面。实验发现最大峰值热流和来流边界层状态关系不大,原因是流动干扰区表现出较强的三维扰动特性,使得来流层流边界层在干扰区内会转变成过渡甚至完全湍流状态。

A study is performed to understand the details of the flow and heating characteristics around surface protuber- ances on hypersonic bodies. Experiments are conducted in a hypersonic gun tunnel at a freestream Mach number of 8.2 and Reynolds number of 9.35 x 108 m-1. The thin-film heat transfer measurements are used to assess the effects of the protuber- ance geometry and boundary layer state on surface heating. Oil-dot visualizations and high-speed schlieren videos are addi- tionally used to qualitatively understand the flowfield around the protuberances. The highest heating is found to the side of the protuberance in interactions in which the incoming boundary layer remains unseparated upstream of the protuberance. In fully separated interactions, the highest heating generally takes place ahead of the protuberance and can become significantly high. The dependence of the maximum heating on the incoming boundary layer state is negligible. This is believed to be caused by the 3-dimensinality of the interactions which causes the incoming laminar boundary layer to become transitional or even fully turbulent.