Abstract: A centrifugal compressor includes an impeller and a diffuser, the diffuser having a plurality of descrete D-shaped passages intersecting each other.

Abstract: An apparatus and method is provided for improving efficiency of a transonic gas turbine engine compressor by bleeding off a shockwave-induced boundary layer from the gas flow passage of the compressor using an array of bleed holes having a downstream edge aligned with a foot of an oblique shock wave which originates on the leading edge of an adjacent transonic rotor blade tip.

Abstract: A centrifugal compressor includes an impeller and a diffuser. The impeller has an inner integral hub with vanes thereon, is adapted to rotate within an outer shroud about a central longitudinal axis, and has a defined hub-to-shroud distribution of fluid exit angles. The diffuser, downstream from the impeller, comprises a plurality of circumferentially spaced discrete passages at least partially defining fluid paths through the diffuser, and angled such that adjacent discrete passages intersect each other to form an annular semi-vaneless diffuser inlet space. The discrete passages downstream of the semi-vaneless space each have an inlet therefrom and an outlet with a greater cross-sectional area than the inlet. The intersection of the annular semi-vaneless space and each discrete passage defines a leading edge thereof.

Abstract: A compressor for a gas turbine engine, comprising a casing having an inner surface defining a plurality of circumferentially spaced-apart grooves adjacent the tips of a row of rotor blades. Each groove as a depth which is less than half the width thereof in order to increase surge margin at off design points without affecting the performance of the compressor under normal operating conditions.