INVESTIGATION OF SCRAMJET PERFORMANCE DURING COMBUSTION MODE TRANSITION

Abstract

Combustion mode transition is a key challenge in dual-mode ramjet engines (DMRJ) due to its direct effect on the engine’s performance and operation. DMRJs are designed to operate both in a ramjet mode at lower supersonic speeds and in a scramjet mode at higher speeds. The investigation of the engine’s performance during the combustion mode transition is desirable to ensure continuous engine operability over the transitional Mach numbers. Depending on the flight speeds, the combustion mode can switch from subsonic combustion mode, commonly known as thermally choked mode, to supersonic combustion mode and vice versa. This study investigates and quantifies the effects of these transitions on combustor stability and propulsion performance metrics.The mode transition is expected to occur near the start of the hypersonic flight regime. Combustion mode transition experiments were conducted using a laboratory-scale, direct-connect test facility that employed a vitiated air heater to match the enthalpy of Mach 5 flight conditions at an altitude of 19 km. A convergent-divergent supersonic nozzle was used to establish a Mach 2 isolator entrance flow, which defined the upstream boundary conditions. The combustor model featured one fuel injector upstream of the cavity flame-holder and three more fuel injectors downstream. Previous studies have shown that natural mode transition occurs at or near the critical equivalence ratio in a series of sporadic transitions back and forth over a relatively substantial length of time spanning around 400~600 msec. This study investigates and quantifies the effects of these transitions on combustor stability and propulsion performance metrics. High-frequency dynamic pressure measurements and axial distribution of wall pressure were measured at various equivalence ratios near the mode transition conditions. The results showed that combustion instabilities are observed during the natural mode transition process at or near the critical equivalence ratio of 0.24. The dominant oscillations were observed at 957 ± 10 Hz with a peak amplitude of 0.2 psi. Also, a cycle analysis was performed to investigate the effect of mode transition on instantaneous thrust. Using previously obtained velocity data and wall pressure measurements, local thrust function was computed, and the cycle analysis was extended into an assumed pressure-matched exit nozzle. The cycle analysis results revealed a substantial effect of combustion mode transition on expected thrust level, changing the amplitude by more than 13%. Finally, a mode transition control strategy is discussed that can not only avoid the combustion instability but also reduce the sudden disruption caused by thrust change associated with the mode transition.