Abstract: To investigate the effects of the boundary-layer bleeding on flow fields in Scramjet combustor, wind tunnel test and numerical simulation were used to study the shock train and combustion characteristics. The results are obtained under the inflow condition of Mach number 2.0, stagnation temperature 950K and stagnation pressure 0.82MPa. When the kerosene is co-firing with pilot hydrogen which equivalence ratio is 0.18, the instable combustion induces the shock train oscillating up and down in the isolator. When the equivalence ratio of kerosene equals 0.29, the boundary-layer bleeding has little effect on the flow field as the shock train leading edge is far away from the bleeding zone. As the kerosene equivalence ratio increases, the shock train moves forward to the bleeding zone, and the dynamic evolution process, structure and location of the shock train are changed by the boundary-layer bleeding. Further, the boundary-layer bleeding effectively improves the resistance capability of the isolator and increases the maximum kerosene equivalence ratio from 0.38 to 0.42. Results show that the region of subsonic combustion and supersonic combustion in the field are also influenced by the boundary-layer bleeding.

Key words: Scramjet engine;Boundary-layer bleeding;Shock train;Subsonic combustion/supersonic combustion

摘要： 为了解边界层抽吸对超燃冲压发动机流场的影响，采用风洞试验和数值计算对隔离段激波串特性以及燃烧室燃烧特性进行了研究。结果表明，在发动机入口马赫数2.0，总温950K，总压0.82MPa的来流条件下，当量比为0.18先锋氢气与不同当量比煤油共同燃烧呈不稳定状态，激波串在隔离段内前后振荡传播。当煤油当量比为0.29时，激波串振荡前缘远离抽吸位置，边界层抽吸对发动机流场基本没有影响。随着煤油当量比逐渐增大，激波串前缘位置到达抽吸区附近，边界层抽吸开始产生影响，改变了隔离段内的激波串动态演化过程、形态结构以及位置分布，同时有效提高了隔离段抗反压特性，使得煤油最大当量比可以由0.38增大至0.42。此外，边界层抽吸对发动机内的亚燃/超燃区域分布也会产生影响。

关键词: 超燃冲压发动机;边界层抽吸;激波串;亚燃/超燃