Large tip clearances in the region of six percent span exist in the high pressure stages of compressors of industrial gas turbines. The over tip clearance flow causes significant blockage and accounts for the largest proportion of loss in the high pressure compressor. This paper examines large tip clearances in two different compressor cascades. The first was a cascade of controlled-diffusion blades and the second cascade had geometry more representative of modern engine practise. The second cascade also featured realistic inlet boundary layer conditions delivered by an upstream injection system. Increasing the understanding of such flows will allow for improvements in the design of such compressors. The key conclusions of this paper are: a) At large tip clearances the tip clearance is the primary variable influencing the flow pattern, blade geometry is a secondary consideration; b) The blade geometry has a significant influence on loss generation through the cascade; c) With increasing tip clearance the loss generation through the cascade can diminish but the flow turning is much reduced; and d) The blade loading at the tip can increase at large tip clearances.

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