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

为了研究马赫数为7的超燃冲压发动机模型内的流动和燃烧，采用了一个小型高焓激波风洞，其喷管出口直径为60 mm。在试验段设置了双楔作为超燃冲压发动机进气道和燃烧室模型。利用飞行器前缘和后缘产生的斜激波系，观察了压缩过程。H^2和/或C_2H_2-空气混合气在燃烧室中垂直喷入马赫数为1.7的超声速横流。利用带像增强器的超高速摄像机（18,000帧/秒）观察到了燃烧气体的自发射。同时，测量了燃烧室壁面静压。结果表明，超声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