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The prevention of separation and flow reversal in the corners of compressor blade cascades

Published online by Cambridge University Press:  04 July 2016

B. S. Stratford
B. S. Stratford*
Affiliation:
Rolls-Royce (1971) Ltd.
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Summary

From tests on tunnel cascades one of the sources of loss in axial compressors could be corner stall, i.e. local separation of the flow in and near the corners between the annulus walls and the convex surfaces of the blades.

Tests on a tunnel cascade at entry Mach numbers of 0·5 and 0·6 show that, at least in a tunnel, the stalling may be cured by boundary layer control. The present tests obtain control by suction through a corner slot. The tests show that removal through the slot of 1·3% of the flow decreases the total losses of the cascade by about 25%.

A more detailed examination of the losses in the tunnel cascade without suction suggests that about three-quarters of the end losses are due to the separation and reverse flow. Much of the remaining end loss appears to result from the fact that the low energy stream tubes in the entry boundary layer would experience a disproportionately large increase in cross-sectional area during diffusion, even in attached two-dimensional flow. The suction corner control can eliminate the greater part of the end losses, by eliminating the separation.

Type
Technical notes
Information
The Aeronautical Journal , Volume 77 , Issue 749 , May 1973 , pp. 249 - 256
Copyright
Copyright © Royal Aeronautical Society 1973 

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References

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