Patent ID: 7147426

Claim:
A transonic gas turbine engine compressor comprising: a rotor having a central axis of rotation and a plurality of blades extending into a gas flaw passage through said compressor, each of said blades having a blade tip and a leading edge defined between opposed pressure and suction surfaces, said rotor being rotatable about said axis of rotation at a speed such that gas flow adjacent said blade tips becomes supersonic, creating oblique shock waves originating at said leading edge and terminating at a shock foot on said suction surface of an adjacent blade; an outer shroud surrounding said rotor, said outer shroud defining a radially outer boundary of said gas flow passage; a plurality of bleed holes extending through at least a portion of said outer shroud adjacent said blade tips to provide gas flow communication between said gas flow passage and an outer bleed passage, said plurality of bleed holes defining a bleed bole array, and said bleed hole array defining a downstream edge thereof substantially aligned in a gas flow direction with said shock foot and extending upstream thereof, said bleed holes in said array being selected in size, number and location to bleed at least a portion of a shockwave-induced boundary layer from said gas flow passage adjacent said outer shroud.