SUPERSONIC MIXING AND COMBUSTION OF HYDROGEN USING HIGH-ENTHALPY SHOCK TUNNEL

为了研究超燃冲压发动机模型在约7马赫飞行条件下的流动和燃烧，采用了一个喷管出口直径为60 mm的小型高焓激波风洞。试验段设置了双楔板，作为超燃冲压发动机进气和燃烧室模型。利用汽车前缘和导流罩产生的斜激波系统，对压缩过程进行了观察。在燃烧室以1.7马赫的超音速横流正常喷射H^2和/或C_2H_2-空气混合气。用带像增强器的高速摄像机(18000fps)观测到了燃烧气体的自发辐射。对燃烧室壁面静压进行了测量。结果表明，当发生超音速燃烧时，壁面压力增大。

To investigate the flow and combustion in a Scramjet engine model at about Mach 7 flight condition, a small high enthalpy shock tunnel, whose nozzle exit was of 60 mm diameter, was used. Double wedges were set in the test section as a Scramjet engine intake and combustion chamber model. The compression process used the oblique shock systems, which were generated by vehicle forward edge and cowl, was observed. H^2 and/or C_2H_2-air mixture gas was injected normally to supersonic cross flow of Mach 1.7 in the combustion chamber. A self-emission from combustion gas was observed by using a very high-speed video camera (18,000fps) with image intensifier. And also, the wall static pressure in the combustion chamber was measured. It was clarified that the wall pressure increased when the supersonic combustion occurred