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An experimental study on a film cooled nozzle guide vane has been conducted in a transonic annular sector to observe the influence of suction and pressure side film cooling on aerodynamic performance. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at an exit reference Mach number of 0.89. The aerodynamic results using a five hole miniature probe are quantified and compared with the baseline case which is uncooled. Results lead to a conclusion that the aerodynamic loss is influenced substantially with the change of the cooling flow rate regardless the positions of the cooling rows. The aerodynamic loss is very sensitive to the blowing ratio and a value of blowing ratio higher than one leads to a considerable higher loss penalty. The suction side film cooling has larger influence on the aerodynamic loss compared to the pressure side film cooling. Pitch-averaged exit flow angles around midspan remain unaffected at moderate blowing ratio. The secondary loss decreases (greater decrease in the tip region compared to the hub region) with inserting cooling air for all cases compared to the uncooled case.

This paper is about the modern mathematical models of working process in the whole flow passage of aviation gas turbine engines. These models are referred as high-level models, based on real 3D geometry of engine flow passage. They allow to simulate steady and unsteady processes in 1D, 2D and 3D formulations, calculate engine performances, determine propagation of radial and circular parameter no uniformities in engine flow passage and predict influence of main parameters on engine efficiency. Typical examples of working process simulation in whole engine and their components are presented below.

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