The present study investigates the film cooling performance over a C3X gas turbine blade under thermodynamic conditions close to engine operating conditions. Three different cooling configurations are investigated: (1) cylindrical film cooling holes without trailing edge discharge slots, (2) cylindrical film cooling holes along with trailing edge discharge slots, and (3) fan-shaped film cooling holes along with trailing edge discharge slots. The trailing edge discharge slots are incorporated to enhance convective cooling in the trailing edge region. Numerical simulations are performed using k−ω shear stress transport turbulence model. The blade material used is a nickel-based single-crystal alloy (CMSX-4), which is widely used in high-temperature turbine blade applications. Exit Mach number (Ma) of flow passage is 0.89. A parametric study is performed to compare all three geometries by varying the coolant-to-mainstream inlet gas pressure ratio. The analysis is conducted at both low and high mainstream gas inlet temperatures. The analysis quantifies the improvement in film cooling effectiveness due to trailing edge discharge slots and fan-shaped cooling holes. The results indicate that the fan-shaped film cooling hole with trailing edge discharge slots provides superior film cooling effectiveness in both the streamwise and spanwise directions compared to the other film cooling hole configurations, even at high mainstream temperature and high coolant-to-mainstream pressure ratio. The findings offer valuable insight into optimizing turbine blade cooling strategies to enhance heat transfer and prevent thermal failures such as creep and thermomechanical fatigue.
Singh et al. (2025) studied this question.