Abstract

Endwall contouring is used to increase the aerodynamic efficiency of both compressor and turbine stages in industrial gas turbines and aeroengines. The complex interaction between the secondary air-leakage, used to cool the turbine disc, and the mainstream gas path, leads to an unsteady flow field that is challenging to compute. Current endwall designs have shown sensitivity to the introduction of secondary air, with stage efficiency improvements being reduced, or in the limit, eliminated altogether.

A computational study of an engine-representative turbine stage was conducted using an unsteady RANS solver. Previously published computations of the baseline axisymmetric endwall were validated with experimental data from a geometrically similar test rig. Understanding from this prior study was used to inform the design process for contoured endwalls, namely through the identification of three key geometric features: the leading-edge feature; the suction-side trough; the pressure-side trough.

The baseline axisymmetric endwall showed periodic unsteadiness, large secondary flow features and an egress plume which dominated the aerodynamics of the stage. The implementation of a suction-side trough (i.e. making the endwall non-axisymmetric) reduced the magnitude of the unsteadiness by controlling the path of the egress plume. The trough also reduced the span-wise migration of the egress plume through the passage and provided modest control over pitchwise position. In corroboration with the findings of other authors, the introduction of a leading-edge feature was also used to reduce the leading edge horseshoe vortex,. The pressure-side trough enabled the prominence of the leading-edge feature to be enhanced, however it increased the span-wise migration of the egress plume. Insight generated from computations of the three distinct geometric features resulted in an improved endwall concept; the improved endwall demonstrated a 0.4% net efficiency gain for the stage relative to the cylindrical baseline.

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