Performance Calculation of Rocket-Based Combined Cycle Engine

After a thorough investigation into the working principles of RBCC components, this paper established mathematical models for performance analysis of each component. A one-dimensional RBCC engine characteristic calculation program was developed based on integral equations, with the core calculation method being iterative flow balance. The computational results of the program were validated through comparison with relevant literature and found to be reasonable. Appropriate engine design point parameters were selected, including an incoming Mach number of 6.0, an inlet mass flow rate of 2.0 kg/s, an inlet exit Mach number of 3.0, a rocket total pressure of 6 MPa, a rocket mass flow rate of 1.0 kg/s, and an equivalent fuel-to-air ratio of 1.0 in the combustion chamber. Throttle characteristics calculations were completed from the injection-ram mode to the subsonic/supersonic combustion mode (Mach 3~6). In the injection-ram mode (Mach=3.0), the engine thrust and specific impulse were jointly affected by the rocket mass flow rate and equivalence ratio in the combustion chamber. As the equivalence ratio increased, the thrust and thrust gain increased, and the specific impulse and specific impulse gain also increased. However, increasing the rocket mass flow rate (with a maximum value of 1.0 kg/s) resulted in a decrease in thrust gain, specific impulse, and specific impulse gain. Under the subsonic/supersonic combustion mode, the engine thrust and unit thrust performance increased with the increase of the equivalence ratio, while the specific impulse performance reached its maximum value around an equivalence ratio of 0.8. The design of maximum thrust control law for the engine and maximum specific impulse control law for the injection-ram mode were completed.