Experimental investigation of film cooling characteristics in the supersonic region of a turbine blade suction surface

Film cooling in the supersonic region of turbine blades is investigated experimentally in a transonic turbine cascade. A measurement approach based on a ternary oxygen mixture combined with binary pressure sensitive paint is employed to obtain film cooling effectiveness and surface pressure distributions under non-unity density ratio conditions, allowing pressure effects to be decoupled from jet-mainstream mixing behavior. Experiments are performed over a range of mainstream pressure ratios, blowing ratios, and density ratios. The results show that mainstream shock impingement does not alter the fundamental influence pattern of the blowing ratio on film cooling, and the streamwise evolution of coolant-mainstream mixing remains qualitatively similar across different pressure ratios. Jet injection in the supersonic region significantly modifies the blade surface pressure field. A bow shock forms upstream of the hole, while the downstream shocks are weakened and displaced. In addition, jet-shock interaction redistributes the recovery temperature on the blade surface by altering shock location and strength, resulting in changes in the presence and position of local recovery-temperature extrema. These results provide experimental insight into film cooling mechanisms in the supersonic region and the associated thermal loading on turbine blades.