A Novel Design Methodology for Compact Turbine Stages Operating Downstream of Rotating Detonation Engines

Abstract During last decades, increased environmental awareness and stricter regulations have pushed aviation industry to adopt measures for reducing aircraft emissions. A promising advancement is Pressure Gain Combustion (PGC), which improves propulsion efficiency and a forthcoming implementation of this technology is Rotating Detonation Engine (RDE). RDEs burn fuel via a supersonically traveling, azimuthally rotating, detonation wave, offering potential performance benefits in efficiency, specific impulse and thrust compared to traditional deflagrative engines. However, the highly unsteady and non-uniform flow field brings challenges especially when retrofitted to turbomachinery, which can suffer performance losses due to RDE’s detonation cycles. Open Literature is limited to designs of large engines characterized by a significant axial extension which hinder additional compactness befits RDEs can offer. The current paper aims at designing a smaller, more compact and short length supersonic turbine, ensuring the self-starting capability of the cascade with minimum shock losses. A high-fidelity total pressure loss model has been developed to delineate an accurate loss budgeting. Numerical simulations, including 2D and 3D URANS studies, examine steady and unsteady stage performance. Specific analysis has been devoted to the effects of inlet unsteadiness on a novel unstarting mechanism due to bow shock coalescence.