The trend towards ultra compact thermal power generation, is constrained by the unavailability of fluid-machinery adequate for supersonic flow conditions. Conventional turbine designs exhibit unacceptable performances related to large aerodynamic losses and a narrow operation range. This paper provides for the first time in the open literature the design procedure, and subsequent analysis of the turbine performance of a turbine adequate for supersonic axial pulsating flows, as those encountered in innovative combustors. The design approach considers the most adverse condition, a steady inlet axial Mach number equal to 3.5. The possible turbine families were classified by the velocity triangles and discussed. A fundamental issue in supersonic passages is to ensure the normal shock at the start of the engine is swallowed through the turbine passages, namely the turbine passage is started. To ensure self-starting capability the turning is restricted to lower values than in the conventional subsonic turbines. The design procedure was based on the method of characteristics, converting the inlet uniform flow into a vortex flow field, such that the adequate deflection is inflicted to the supersonic flow. The performance of the supersonic passage was first assessed and then compared to conventional designs. The present design procedure and analysis of unconventional supersonic turbines provides guidelines for the design and optimization of efficient high supersonic passages, suitable to future tightly packed fluid machinery

Rotating detonation combustors (RDCs) offer theoretically a significant total pressure increase, which may result in enhanced cycle efficiency. The fluctuating exhaust of RDC, however, induces low supersonic flow and large flow angle fluctuations at several kHz, which affects the performance of the downstream turbine. In this paper, a numerical methodology is proposed to characterize a supersonic turbine exposed to fluctuations from RDC without any dilution. The inlet conditions of the turbine were extracted from a three-dimensional (3D) unsteady Reynolds-averaged Navier–Stokes simulation of a nozzle attached to a rotating detonation combustor, optimized for minimum flow fluctuations and a mass-flow averaged Mach number of 2 at the nozzle outlet. In a first step, a supersonic turbine able to handle steady Mach 2 inflow was designed based on a method of characteristics solver and total pressure loss was assessed. Afterward, unsteady simulations of eight stator passages exposed to periodic oblique shocks were performed. Total pressure loss was evaluated for several oblique shock frequencies and amplitudes. The unsteady stator outlet profile was extracted and used as inlet condition for the unsteady rotor simulations. Finally, a full stage unsteady simulation was performed to characterize the flow field across the entire turbine stage. Power extraction, airfoil base pressure, and total pressure losses were assessed, which enabled the estimation of the loss mechanisms in supersonic turbine exposed to large unsteady inlet conditions.

In this paper, OpenFoam, an open source unsteady Reynolds-averaged Navier–Stokes software was carefully examined to analyze rotating detonation combustors. This tool was subsequently used to quantify the effect of the outlet conditions on five different exhaust nozzle geometries. Both detonation and deflagration of the hydrogen–air mixture were taken into account in the source terms of the species transport equation. The premixed ${\mathrm{H}}_2$–air mixture was injected at three different total pressures, ranging from 0.4 to 0.8 MPa. First, the rotating detonation combustor was described, and then five different nozzles were investigated: straight duct, conical, Bezier outer wall, Bezier inner wall, and two Bezier surfaces. All the nozzles were compared regarding their outlet total temperature, pressure gain, exit Mach number, and outlet flow angle. The use of a straight duct nozzle generated a noticeable drop in pressure gain of about 27% compared to the baseline combustor. The conical nozzle expanded the combustor outlet flow from sonic conditions into Mach 1.8 to 2.3. The Bezier outer wall underperformed compared to the conical nozzle at low combustor inlet pressures, while the situation was reversed at higher combustor inlet pressures. Using the inner wall as a divergent section showed no benefits in terms of Mach number and total pressure. The use of a nozzle with two Bezier surfaces did not provide any significant benefits in spite of the enhanced outlet-to-combustor-inlet-area ratio.

The paper focuses on the unsteady pressure field measured around the rotor midspan profile of the VKI Brite transonic turbine stage. The understanding of the complex unsteady flow field is supported by a quasi-three-dimensional unsteady Navier–Stokes computation using a k-ω turbulence model and a modified version of the Abu-Ghannam and Shaw correlation for the onset of transition. The agreement between computational and experimental results is satisfactory. They both reveal the dominance of the vane shock in the interaction. For this reason, it is difficult to identify the influence of vane-wake ingestion in the rotor passage from the experimental data. However, the computations allow us to draw some useful conclusions in this respect. The effect of the variation of the rotational speed, the stator–rotor spacing, and the stator trailing edge coolant flow ejection is investigated and the unsteady blade force pattern is analyzed.