An improvement in overall efficiency and power output for gas turbine engines can be obtained by increasing the combustor exit temperature, but the thermal management of metal parts exposed to hot gases is challenging. Discrete film cooling, combined with internal convective cooling is the current state-of-the-art available to aerothermal designers of these components. To simplify the simulation problem in the aerodynamic design phase, it is common practice to replace the cooling holes with source strips applied to the blade. This could lead to inaccuracies in high pressure turbine performance prediction. This study has been carried out on a fully-featured high pressure turbine stage using high-fidelity simulations. The film cooling holes on the nozzle guide vane and on the rotor are initially modelled using a strip model approach. Then, to increase the model fidelity, the strips on the suction side of the rotor are replaced with discrete fan shaped film cooling holes. A rigid body rotation is also applied to the nozzle guide vane to vary the stage capacity and reaction. The effects of the mesh topology & resolution are also taken into account. The results obtained with these two approaches are then compared, giving the designers a better understanding on film cooling modelling and relationship between capacity, reaction and performance. The accurate prediction of the complex interaction between cavity inflows and the main-flow, still represent a challenge for the state of the art RANS solvers. Hence, an unsteady phase-lag approach has been used to overcome the RANS limitations. A validation of the unsteady solutions has been carried out with respect to experimental data.

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