Investigation of the Flow Field Morphology of Film Cooling in Supersonic Flow

Film cooling is widely implemented in highly thermally stressed gas turbine components. Its performance has been extensively investigated for several decades and many results are available in the literature. In conventional gas turbines, regions of supersonic flow are not prevalent and should generally be avoided. For this reason, results relative to film cooling in supersonic flow are limited. Nevertheless, a new interest related to Rotating Detonation Combustors (RDC) and supersonic turbines is growing. The implementation of those engine components in a gas turbine is likely to need film cooling for thermal protection. In this context, it becomes crucial to gain an understanding of how the film interacts with the freestream when operated in a supersonic flow. This paper investigates the effect caused by the injection of film cooling on the morphology of the supersonic flow field. Results obtained by means of schlieren imaging indicated that the coolant injection acts as a wedge inside the flow, determining the local formation of an oblique bow shock around each film cooling hole. The shape, inclination, and strength of the oblique shock showed a dependency on the fundamental dimensionless parameters considered for the characterization of the operating conditions of film cooling. Furthermore, as the amount of mass injected was increased, the inclination of the generated shocks increased and the impingement location of the reflected shock moved upstream along the injection plate. The fluid dynamics of this interaction affected the local pressure distribution on the injection plate, measured by means of Pressure Sensitive Paint (PSP). Different film cooling geometries and main flow conditions were tested at multiple operating conditions. The relative impact of the different parameters is presented, providing useful information for the design of a film cooled engine component exposed to a supersonic flow.