A compressor rotor includes a first disk and a conical section connected to the first disk. The conical section includes at least one flow hole. A bore cavity is defined between the conical section and the first disk. The bore cavity is arranged in fluid communication with the at least one flow hole. An anti-vortex tube is disposed within the at least one flow hole of the conical section and includes at least one feature arranged in contact with a surface of the conical section to restrict movement of the anti-vortex tube out of engagement with the conical section.

Disclosed herein is a system for cooling a gas turbine. The system for cooling a gas turbine cools a turbine disk unit by individually supplying cooling air to each of a plurality of turbine disks.

A compressor cooling apparatus includes: a blade row mounted for rotation about a centerline axis; a stationary diffuser located downstream of, and in flow communication with, the blade row; an inducer disposed between the diffuser and the blade row, the inducer having an inlet in flow communication with the diffuser, and having an outlet oriented to direct flow towards the blade row.

The present disclosure is directed to a system for mitigating rotor bow at a turbine engine. The system includes a casing circumferentially surrounding a rotor assembly in which a heat pipe is attached to the casing and extended circumferentially around the rotor assembly.

A system includes an inducer assembly configured to receive a fluid flow from compressor fluid source and to turn the fluid flow in a substantially circumferential direction into the exit cavity. The inducer assembly includes multiple flow passages. Each flow passage includes an inlet configured to receive the fluid flow and an outlet configured to discharge the fluid flow into the exit cavity, and each flow passage is defined by a first wall portion and a second wall portion extending between the inlet and the outlet. The first wall portion includes a first surface adjacent the outlet that extends into the exit cavity.

In one aspect, the present disclosure is directed to a cooling circuit for a gas turbine engine. The cooling circuit includes a rotor blade having a connection portion and a rotor disc having a first axial side and a second axial side. The rotor disc defines a connection slot and a cooling passage extending between the first axial side and the second axial side. The connection slot receives the connection portion to couple the rotor blade to the rotor disc. Cooling air flows through the cooling passage.

The rotor disk (3B) for a compressor comprises, relative to the rotational axis of the disk: a radial web (4), blades (8) at the outer periphery of the web, a bore (5) at the inner periphery of the web, and a cylindrical side wall (12) extending the web in the vicinity of the ter periphery of same and having an air supply port (18), anda centripetal air collection device (15). Advantageously, the device (15) comprises a cylindrical support (23) and at least one air supply tube (16), the inlet of which is turned towards the port (18) and the outlet of which is turned towards the bore (5) in the web, the disk comprising an inner radial flange (40) extending from the cylindrical side wall (12), the cylindrical support (23) of the device (15) being attached to said inner radial flange (40), and a ring (30) extending from the web (4), the cylindrical support (23) being centred on the ring (30).

An apparatus for a turbine engine comprising an outer casing, an engine core provided within outer casing and having a at least one set of blades, and through which gasses flow in a forward to aft direction, an outer drum located within the outer casing to define an annular cavity. A set of seals extending between the first surface and the second surface to define at least one cooled cavity within the annular cavity.

A flow diverter for an air separator of a gas turbine includes a cylindrical body configured to fit within a cooling hole of the air separator. One or more air flow vents are defined through and around a partial circumference of the cylindrical body. A bottom panel closes the cylindrical body at one end. A mounting flange surrounds the cylindrical body at an open end and extends radially outward from the cylindrical body. When the flow diverter is installed, air flows through the open end in a radial direction and exits through the air flow vents in an axial direction. The cylindrical body may include a collapsible region that collapses to engage the air separator and prevent the flow diverter from being dislodged. The air separator may additionally or alternately include supplemental cooling holes in a recessed area proximate to its mounting flange.

An apparatus and method for cooling a portion of a turbine engine comprising an outer casing defining an axial centerline, a turbine section through which a flow of combustion gasses flows in a forward to aft direction, an outer drum located between the outer casing and the turbine section defining an annular cavity therebetween. A set of seals extends between the outer casing and outer drum to define at least one cooled cavity.