Cavity ignition was investigated in a rocket-based combined cycle combustor. The Mach number, total pressure, and total temperature of the inflow were 2.92, 2.60 MPa, and 1650 K, respectively. A backward-facing step was employed to replace the rocket engine that was shut down. The expansion fan and the reattachment shock wave resulting from the backward-facing step distorted the mainstream. The combustor was 40 mm in height (H) at the entrance. Three typical cases with $d/H=2.0$, 4.5, and 7.5 were reported in this study, where d was the distance between the backward-facing step and the cavity. Schlieren images and flame chemiluminescence images were captured to reveal the ignition processes. The ignition process in the case with $d/H=2.0$ was similar to the conventional cavity ignition because only the rear of the cavity was subjected to the expansion fan. Two combustion modes, namely, the local combustion mode and the cavity stabilized scramjet mode, were witnessed in this case. The first combustion mode was characterized by a small reaction zone in the fore part of the cavity. The primary reaction zone of the second mode was in the region near the cavity ramp. The rapid flame spreading in the cavity shear layer led to the transition between these two combustion modes. In the case with $d/H=4.5$, the reattachment shock wave enlarged the cavity recirculation flow. The expansion fan decreased the static pressure upstream of the cavity, which made the combustor more sensitive to backpressure. Although the first two combustion modes of this case were similar to those of the case with $d/H=2.0$, a new combustion mode (unsteady jet-wake stabilized scramjet mode) was observed in this case. The mode transition between the cavity stabilized scramjet mode and the unsteady jet-wake stabilized scramjet mode was attributed to the backpressure resulting from the heat release. The backpressure gradually pushed the reaction zone upstream. In the case with $d/H=7.5$, the cavity was not directly exposed to the expansion fan or the reattachment shock wave. However, the ignition process in this case was more complex. The combustor successively experienced the aforementioned three combustion modes and eventually settled into the ramjet mode. The mode transition between the last two modes was the result of thermal choking. The flame propagated upstream rapidly when the combustor was chocked.

在基于火箭的联合循环燃烧器中研究了空腔点火。流入的马赫数、总压力和总温度分别为 2.92、2.60 MPa 和 1650 K。采用后向步骤来更换已关闭的火箭发动机。膨胀扇和后向台阶产生的重附冲击波扭曲了主流。入口处的燃烧器高度 ( H )为 40 mm 。三个典型案例$d/H=2.0$本研究报告了 、4.5 和 7.5，其中d是后向台阶和空腔之间的距离。捕获纹影图像和火焰化学发光图像以揭示点火过程。点火过程$d/H=2.0$类似于传统的腔体点火，因为只有腔体的后部受到膨胀风扇的影响。在这种情况下，见证了两种燃烧模式，即局部燃烧模式和腔稳定超燃冲压发动机模式。第一种燃烧模式的特点是空腔前部有一个小的反应区。第二种模式的主要反应区位于靠近腔斜坡的区域。空腔剪切层中的快速火焰蔓延导致这两种燃烧模式之间的转变。在这种情况下$d/H=4.5$，重附激波扩大了腔体再循环流。膨胀风扇降低了腔体上游的静压，使燃烧室对背压更加敏感。虽然本案例的前两种燃烧模式与$d/H=2.0$，在这种情况下观察到一种新的燃烧模式（不稳定射流尾流稳定超燃冲压发动机模式）。腔稳定超燃冲压发动机模式和非定常射流尾流稳定超燃冲压发动机模式之间的模式转变归因于放热产生的背压。背压逐渐将反应区推向上游。在这种情况下$d/H=7.5$，腔体没有直接暴露在膨胀风扇或再附着冲击波中。然而，这种情况下的点火过程更为复杂。燃烧室先后经历了上述三种燃烧模式，最终进入冲压喷气发动机模式。最后两种模式之间的模式转换是热扼流的结果。当燃烧器被堵塞时，火焰迅速向上游传播。