Film Cooling Effectiveness and Heat Transfer on the Trailing Edge Cutback of Gas Turbine Airfoils With Various Internal Cooling Designs

The present study deals with trailing edge film cooling on the pressure side cutback of gas turbine airfoils. Before being ejected tangentially onto the inclined cut-back surface the coolant air passes a partly converging passage that is equipped with turbulators such as pin fins and ribs. The experiments are conducted in a generic setup and cover a broad variety of internal cooling designs. A subsonic atmospheric open-loop wind tunnel is utilized for the tests. The test conditions are characterized by a constant Reynolds number of $Rehg=250000$⁠, a turbulence intensity of $Tuhg=7%$⁠, and a hot gas temperature of $Thg=500K$⁠. Due to the ambient temperature of the coolant, engine realistic density ratios between coolant and hot gas can be realized. Blowing ratios cover a range of $0.20<M<1.25$⁠. The experimental data to be presented include discharge coefficients, adiabatic film cooling effectiveness, and heat transfer coefficients in the near slot region $(x∕H<15)$⁠. The results clearly demonstrate the strong influence of the internal cooling design and the relatively thick pressure side lip $(t∕H=1)$ on film cooling performance downstream of the ejection slot.