Abstract: A gas turbine engine that comprises an engine core having a compressor, a turbine, a combustor, and a rotation axis. The engine further comprising a fan case wall having a radially-inner surface, a by-pass duct between the fan case wall and around the engine core, a plurality of struts circumferentially distributed and extending across the by-pass duct, a surface cooler adjacent to the radially-inner surface of the fan case and configured to be exposed to a by-pass air flow through the by-pass duct during operation of the gas turbine engine, the surface cooler fluidly communicating with a fluid circuit of the engine requiring cooling for of a fluid, and a recirculation conduit extending between an inlet in the radially-inner surface of the fan case disposed downstream of the surface cooler and an outlet in the radially-inner surface of the fan case wall disposed upstream of the surface cooler.

Abstract: A compressor diffuser for a gas turbine engine which includes at least one diffuser pipe having a tubular body with an inner surface defining an internal flow passage extending therethrough. The tubular body includes a first portion that extends in a first direction and defines a throat therein, a second portion that extends in a second direction different from the first direction, and a curved portion fluidly linking the first portion and the second portion. A plurality of vortex generators are disposed within the diffuser pipe and extend from the inner surface into the internal flow passage. The vortex generators are disposed downstream of the throat in the first portion of the diffuser pipe and upstream of the curved portion. In operation, the vortex generators engage fluid flow in the internal flow passage to generate downstream vortices.

Abstract: A gas turbine engine compressor includes a rotor defining a central axis of rotation and a plurality of blades which project into an annular compressor gas flow passage, and a shroud circumferentially surrounding the rotor and having a radially inner surface adjacent to the blade tips. Bleed holes extend through the shroud adjacent the blade tips, each of the bleed holes having an inlet end disposed in the shroud radially inner surface and an outlet end disposed in a shroud radially outer surface. Bleed air removed from the annular gas flow passage flows through the bleed holes from the inlet to the outlet ends. The outlet end of each bleed hole is located circumferentially upstream of the inlet end relative to a direction of rotational flow in the annular gas flow passage driven by a direction of rotation of the rotor.

Abstract: A centrifugal compressor having at least a diffuser is disclosed. The diffuser has an annular diffuser body having circumferentially spaced apart diffuser passages defining fluid paths through the diffuser body. The diffuser also has a plurality of diffusion members mounted to the annular diffuser body. Each diffusion member has a member inlet in fluid communication with a diffuser passage and a member outlet. Each diffusion member defines an aerodynamic throat disposed between the member inlet and the member outlet. The diffuser also has a fluid injection assembly with multiple injection conduits. Each injection conduit extends between a conduit inlet configured to receive a flow of compressible fluid from a supply and a conduit outlet communicating with a corresponding diffusion member downstream of the aerodynamic throat. The compressible fluid is injected through the conduit outlet into the diffusion members.

Abstract: A compressor for a gas turbine engine having a bleed air recirculation system includes a plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips. The bleed holes have a closed outer perimeter along their complete length. An annular bleed cavity surrounds the shroud and is in communication with outlet openings of the bleed holes. The bleed holes provide communication between the main gas flow passage and the bleed cavity. The bleed cavity includes exit passages having outlets disposed in the shroud at a second axial location which is upstream of both the first axial location and the leading edge of the blades of the rotor. Bleed air is passively bled from the main gas flow passage via the bleed holes, recirculated through the bleed cavity and re-injected back into the main gas flow passage at the second axial location.

Abstract: The centrifugal compressor described includes an impeller shroud which encloses the impeller and has a curved shroud surface that extends between an inducer portion and an exducer portion. The compressor includes one or more circumferential grooves in the shroud body within the exducer portion. Each groove has opposed wall segments spaced apart therefrom. The wall segments are inclined at a nonzero groove angle relative to a normal of the shroud surface in a direction opposite the fluid flow path along the shroud surface.

Abstract: A compressor diffuser for a gas turbine engine which includes at least one diffuser pipe having a tubular body with an inner surface defining an internal flow passage extending therethrough. The tubular body includes a first portion that extends in a first direction and defines a throat therein, a second portion that extends in a second direction different from the first direction, and a curved portion fluidly linking the first portion and the second portion. A plurality of vortex generators are disposed within the diffuser pipe and extend from the inner surface into the internal flow passage. The vortex generators are disposed downstream of the throat in the first portion of the diffuser pipe and upstream of the curved portion. In operation, the vortex generators engage fluid flow in the internal flow passage to generate downstream vortices.

Abstract: A compressor for a gas turbine engine having a bleed air recirculation system includes a plurality of bleed holes extending through the shroud at a first axial location thereon substantially adjacent the blade tips. The bleed holes have a closed outer perimeter along their complete length. An annular bleed cavity surrounds the shroud and is in communication with outlet openings of the bleed holes. The bleed holes provide communication between said main gas flow passage and the bleed cavity. The bleed cavity includes exit passages having outlets disposed in said shroud at a second axial location which is upstream of both the first axial location and the leading edge of the blades of the rotor. Bleed air is passively bled from the main gas flow passage via the bleed holes, recirculated through the bleed cavity and re-injected back into the main gas flow passage at the second axial location.

Abstract: A rotor for a gas turbine engine includes a plurality of radially extending blades, each having a remote blade tip defining an outer tip surface, and a leading edge defined between opposed pressure and suction side airfoil surfaces. A shroud circumferentially surrounds the rotor, and a radial distance between an inner surface of the shroud and the outer tip surface of the blades defines a radial tip clearance gap therebetween. The tip of each of the blades has a pressure side edge formed at the intersection between the outer tip surface and the pressure side airfoil surface, and a suction side edge formed at the intersection between the outer tip surface and the pressure side airfoil surface. The suction side edge has a larger radius of curvature than the pressure side edge, thereby reducing the amount of tip leakage flow through the radial tip clearance gap.