Film cooling of 1st stage rotor blades strongly relates to the aerodynamic performance and life expectation of gas turbine. In this piece of work, a parametric study was carried out based on 3D RANS CFD methods to systematically investigate both cooling performance and aerodynamic loss of various film holes’ arrangements on the pressure side (PS) of a realistic rotor blade. In each CFD case, cylindrical film holes were arranged in one row with P/D around 4.0. More than one hundred CFD cases were carried out with various dimensionless streamwise locations (0.1, 0.2, ..., and 0.7), compound angles (−60°, −30°, ..., and 90°), and relative coolant massflows (0.2%, 0.3%, 0.45%, 0.7%, and 1.0% of mainstream massflow).
Detailed distributions of adiabatic film cooling effectiveness on blade surface and aerodynamic loss with film cooling were computed for each case. It was found that coolant massflow was the decisive factor of aerodynamic efficiency φ. For near-LE film holes, −60° compound angle gave higher and more uniform η distributions at various coolant massflows. And the separation near LE could be suppressed, improving local cooling at near-hub on PS. For downstream film holes, 60° and −60° compound angles provided equivalent cooling performances, which were also better than other cases. The cooling characteristics of film holes with −30° compound angle were more susceptible to both coolant massflow and the location of ejection, which was mainly because of the blow-off and reattachment of coolant jets. For the cases with 90° compound angle, the distributions of η were generally similar, and the values increased monotonically as coolant massflows get larger.