The impact on compressor flow range of circumferential casing grooves of varying groove depth, groove axial location, and groove axial extent is assessed against that of a smooth casing wall using computational experiments. The computed results show that maximum range improvement is obtained for a groove depth of approximately a tip clearance and of an axial width of approximately 0.175 axial chord (approximately the near stall tip clearance vortex core size for the smoothwall compressor examined). It was found that for a groove of specified depth and axial width there are two axial locations, one at approximately 0.15 axial chord and one at approximately 0.6 axial chord from the rotor blade leading edge, to position the groove for maximum flow range improvement; this result is in accord with recent experimental observations. In addition, a method of extracting body forces is used to show that the force is finite at the blade tip and non-vanishing in the tip clearance.
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