The rocket ejector refers to a core component of a rocket-based combined cycle (RBCC) engine. The ignition is of critical significance for rocket ejection. Reliable and stable ignition crucially determines the normal operation of the engine. In this paper, a thrust chamber with coaxial swirl injector for the RBCC rocket ejection was developed and tested. Gas oxygen (GOX) and kerosene acted as propellants. As revealed from the test results, the process of ignition pressurizing comprised four phases. The oxygen prefilling time before ignition slightly impacted the ignition time, whereas it affected the peak pressure of ignition. In a confined range, the peak pressure decreased as the prefilling time was extended. The ignition was simulated by building a numerical model, and the results well complied with the experimentally achieved results. The numerical model is capable of specifically indicating the position of the kernel of fire and the process of flame propagation. The simulation results reveal that the propellant could form a combustible condition within 4 ms. The kernel was 6 mm away from the injector, located at the oxygen and kerosene mixing interface and approaching the upper wall. The above results reflected the vital role of the central recirculation zone formed by the prefilled oxygen. The ignition energy was transported near the injector under the convection effect, which ignited the stoichiometric mixture, and the entire ignition could reach a stable state within 20 ms. The numerical model which was developed in this paper can help clarify the combustion mechanism.
Analysis of a transpiration cooled LOX/CH4 rocket thrust chamber
