1. Aerospace propulsion is intrinsically related to the passage of supersonic flows across a throat. It is, however, essential to address how we achieve the supersonic flow from an initial no-flow state. This process is called starting; as we transition from no flow into supersonic, a moving normal shock must be swallowed across the convergent passage. To ensure steady supersonic operation, one should avoid sonic conditions in the convergent section upstream of the throat. Self-starting is the process in which a passage may transition from subsonic to supersonic without any geometry change. It is essential to understand the underlying physics of this process to be able to reduce related losses significantly. For fast geometry generation and its assessment of flow starting a new design tool is required to evaluate this transient process efficiently.
2. Experimental validation of the transient models needs to be provided. Therefore an experimental facility needed to be designed to be able to assess transient flow phenomena in the transonic and supersonic regime. A shock generator was designed and built to be able to imitate the oblique shock occurring in a rotating detonation engine.
3. The evaluation of aerodynamic losses is an important challenge in the aerodynamic design of virtually any flow device. For the development of reduced loss models in supersonic flows, the loss creation in the boundary layer, the tip leakage vortex, or other secondary flow structures need to be analyzed. Also, it is required to analyze losses in a transient flow field to understand loss-creating mechanisms in a turbine exposed to the outflow of a rotating detonation engine. This way the performance of two different geometries can be compared and judged. Hence, the formulation of a physically consistent loss definition for aerodynamic losses derived from computational fluid dynamics CFD results was realized. Based on these loss analysis reduced models can be derived and implemented in a 1D Euler solver.
4. Rotating detonation engines and organic Rankine cycles would profit from efficient turbomachinery, which can perform a subsonic supersonic flow transition. To suit this need, a baseline geometry for the first turbomachinery stage has to prove the design concept. Following, a design optimization will explore the limits of this concept. For further efficiency gains and to allow for integration of the novel concept in existing engines a second stage needs to be designed to allow for the creation of the right outflow conditions.
Conclusions:
1. The underlying process of the normal shock movement during the starting process could be identified. This knowledge was used to design a reduced model for the rapid design of passages with improved startability.
2. An experimental facility that allows the testing of supersonic passages taking advantage of the hydraulic analogy was designed and built. The startability of several passages could be assessed confirming the numerically found trends. This way the in 1. designed reduced-order model could be validated. Furthermore, a shock generator was designed and built for the same facility. This shock generator allows us to imitate the occurring oblique shock in a rotating detonation combustor.
3. Aerodynamic and Aerothermal losses were defined based on the entropy transport equation. This allows the analysis of irreversible flow losses locally and at each time instance based on CFD results. This way complex CFD results with transient boundary conditions as the ones used in simulations of rotating detonation combustors can be assessed and the efficiency of turbines can be computed. A method to calculate the efficiency in an experimental setting was also proposed.
4. A new turbine technology was designed and optimized to take advantage of the pressure gain in a rotating detonation combustor. The technology was numerically assessed under transient conditions and showed dominant efficiency compared to known designs for this application.
