来流总温对双模态燃烧室模态转换边界的影响

针对煤油燃料双模态超声速燃烧室，开展了来流总温对燃烧室模态转换边界影响的试验研究。试验采用甲烷燃烧加热直连式试验系统，隔离段进口马赫数保持2.0不变，总压为1.05 MPa，来流总温分别为885、1 085、1 285 K。试验中采集了燃烧室沿程壁面压力，并采用一维分析方法得到了燃烧室的工作模态。试验结果表明:来流总温不同时，燃烧室壁面峰值压力位置相同，同时压力峰值与隔离段壁面压力分布和激波串起始位置存在一一对应关系；来流总温上升导致燃烧室超燃-亚燃模态转换时的当量油气比上升；在燃烧室当量油气比不变的条件下，来流总温上升能够导致燃烧室壁面压力下降，隔离段内激波串长度缩短。 

Effects of incoming flow total temperature on mode transition boundary in dual mode scramjet combustor

To estimate the effects of incoming flow total temperatures on mode transition boundary in a kerosene-fueled dual mode combustor, an test study was conducted using a direct-connected methane combustion heating test facility. The values of incoming flow total pressure at the isolator inlet should be kept at 1.05 MPa, the Mach number was about 2.0, the incoming flow total temperatures were 885, 1 085, 1 285 K, respectively. Combustor wall pressures were measured by pressure sensors; the mode of combustion was identified by calculating the Mach number at the isolator exit using a one-dimensional model. Test results indicated that, the locations of the combustor peak wall pressure were the same at different incoming flow total temperatures; there was a one-to-one relationship between the combustor peak pressure value, isolator near wall pressure distribution and shock train leading edge location; with the increase of incoming flow total temperature, the equivalence ratio of mode transition from supersonic to subsonic combustion rose up; under the same kerosene equivalence ratio, as the incoming flow total temperature increased, the combustor wall pressure dropped, the length of shock train got shorter.