Conventional deflagrarive rocket propulsion has reached a perfomance plateau such that technical development has shifted from performance improvement to improvements in system reliability and manufacturability. Engine cycles utilizing detonative propulsion have been studied extensively for airbreathing engines, first focusing on pulsed detonation engines (PDEs) and more recently on rotating detonation engines (RDEs). A preliminary idealized thermodynamic analysis of RDEs for rocket propulsion applications is presented. Our initial analysis has shown an ideal performance gain of approximately 10% in the event of equivalent initial pressure ratios, which is in agreement with past ideal cycle analysis. Additionally, our preliminary results have shown that detonative combustion can be used to produce work with the same efficiency as deflagrarive systems while reducing initial pressure ratios by a factor of 5-8, allowing for system simplification that could reduce dry mass and overall cost.
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