The aerodynamic force coefficients of five linear cascade of existing film cooling
turbine blades are evaluated numerically. The blade is geometrically identical to the
first rotor blades of the high pressure (HP) turbine of the F-100-PW-220 military
turbofan. Cascade turbine blade test rig has been designed, constructed, and calibrated
to introduce experimental work for the same flow conditions of the numerical solution
to validate correctness of the numerical results. The numerical simulation shows
acceptable agreement with experimental. Also it was found experimentally that both
lift and drag coefficients are increased slightly with add of film cooling.
The local Mach number distributions outside the boundary layer on both blade
sides of the cascade blade are evaluated numerically and compared with the results of
well known CFD code (Fine/Turbo) for existing gas turbine rotor stage of identical
blade. The computational results obtained for both cases show that the Mach number
distributions trend along both blade sides for rotor stage and cascade are approximately
the same, and the values of Mach number of rotor stage are higher than that for the
corresponded cascade. Also it was found that the Mach number distributions on both
blade sides are reduced in values by the addition of air cooling, and the local Mach
numbers for the cascade case is reduced in values among the rotor stage for the two
cases with and without film cooling on both blade sides.