2026-09-20T08:20:40Zhttps://dspace.mit.edu/server/oai/requestoai:dspace.mit.edu:1721.1/596752022-01-13T07:54:11Zcom_1721.1_7582com_1721.1_7581col_1721.1_131023
Choon S. Tan and Edward M. Greitzer.
Hanley, Brian K. (Brian Kyle)
Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics.
Massachusetts Institute of Technology. Department of Aeronautics and Astronautics
2010-10-29T18:09:16Z
2010-10-29T18:09:16Z
2010
2010
http://hdl.handle.net/1721.1/59675
668214452
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2010.
Cataloged from PDF version of thesis.
Includes bibliographical references (p. 77).
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.
by Brian K. Hanley.
S.M.
77 p.
eng
Massachusetts Institute of Technology
M.I.T. theses are protected by
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http://dspace.mit.edu/handle/1721.1/7582
Aeronautics and Astronautics.
Effect of circumferential groove casing treatment parameters on axial compressor flow range
Thesis
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<Abstract>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.</Abstract>
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