59111 - Aero-Thermal Characterization of Accelerating and Diffusing Passages Downstream of Rotating Detonation Combustors 

Cycle benefits of rotating detonation engines show up to five percentage points for low pressure ratio engines. To realize this  gain, we must ensure an optimal integration between the combustor and the turbine. Indeed, the rotating detonation combustor (RDC) exhausts transonic flow with shocks rotating at few kilohertz. Alas, traditional high pressure turbines are designed to operate under quasi-steady inlet conditions at Mach 0.15. Hence, diffusers tailored to reduce fluctuations and speed at minimum pressure loss are required. The objectives of this paper comprise the design of low-loss diffusers for rotating detonation combustors and a robust design tool to achieve combustor-independent diffuser geometries. The analysis of the combustor is performed via reactive unsteady Reynolds Averaged Navier-Stokes (URANS) simulations. The unsteady RANS equations are solved via CFD++ from Metacomp with a one-step reaction mechanism for H₂-air. The resolving of the boundary layer is achieved with a structured mesh of around 40 million cells. An inlet pressure of 10 bar and a fixed back pressure at the exit of 7 bar are imposed to the combustor. Consecutively, the outlet conditions of the combustor were imposed in diffuser-only simulations and heat flux, pressure loss and unsteadiness at the outlet of the diffuser are assessed for multiple diffuser geometries.