A rotating machine includes a hub portion, wherein the hub portion comprises a forward face and an aft face. The rotating machine further includes a cooling channel formed on either the forward face or the aft face and configured to direct cooling air to a location on the rotating machine, wherein the cooling channel extends from a radially inner location along the face to a radially outer location along the face, and wherein the cooling channel is configured as a recess formed into an outer surface of the face.

The present disclosure is directed to a nozzle cooling system for a gas turbine engine. An impingement plate is positioned radially inwardly from a radially inner surface of an inner side wall of a nozzle. The impingement plate and the inner side wall collectively define an inner chamber. The impingement plate includes a first portion defining one or more impingement apertures and a second portion defining one or more post-impingement apertures. A duct plate encloses the first portion of the impingement plate. The duct plate, the first portion of the impingement plate, and inner side wall collectively define an outer chamber in fluid communication with the inner chamber through the one or more impingement apertures. Compressed air from the outer chamber flows through the one or more impingement apertures into the inner chamber and exits the inner chamber through the one or more post-impingement apertures.

Described is a heat-protection element (50) for a gas turbine (10), in particular an aircraft gas turbine, the heat-protection element (50) being adapted to at least partially surround a bearing chamber (60) of the gas turbine (10) and having at least one connecting portion (52) which is disposed in an axially forward region (VB) and connectable or connected by a material-to-material bond to a protective element (54) of a seal carrier, in particular a seal carrier with a carbon seal, at least one supporting portion (58) which is disposed in an axially central region (MB) and adapted to support the heat-protection element (50) radially on the bearing chamber (60), an end portion (64) which is disposed in an axially rearward region (HB) and forms a free end (66) of the heat-protection element (50) and which is configured such that the end portion surrounds (64) the bearing chamber (60) in a contactless manner.

The rotor assembly can have a first disc having a first body extending circumferentially and radially around the axis, a first set of circumferentially distributed blades protruding radially from the first disc, and a male spline extending axially relative the first body, the male spline extending around and along the axis, and a second disc having a second body extending circumferentially and radially around the axis, a second set of circumferentially distributed blades protruding radially from the second disc, and a female spline extending around and along the axis, the female spline receiving the male spline in a spline engagement.

An apparatus and method for an engine component for a turbine engine. The engine component having an outer wall defining an interior and extending between a root and a tip to define a radial direction, a tip wall spanning the first side and second sides to close the interior at the tip. A tip rail extending from the tip wall and having an inner tip rail surface, an outer tip rail surface extending from at least one of the first or the second side, and radially terminating in an upper tip rail surface connecting the inner tip rail surface and the outer tip rail surface. A tip rim formed in at least one of the outer surface or the inner tip rail surface and spaced from the upper tip rail surface in the radial direction, and multiple cooling passages formed in the outer wall and fluidly coupling the at least one cooling conduit to the tip rim at corresponding passage outlets.

A turbine airfoil cooling system including a low pressure cooling system and a high pressure cooling system for a turbine airfoil of a gas turbine engine is disclosed. In at least one embodiment, the low pressure cooling system may be an ambient air cooling system, and the high pressure cooling system may be a compressor bleed air cooling system. In at least one embodiment, the compressor bleed air cooling system in communication with a high pressure subsystem that may be a snubber cooling system positioned within a snubber. A delivery system including a movable air supply tube may be used to separate the low and high pressure cooling subsystems. The delivery system may enable high pressure cooling air to be passed to the snubber cooling system separate from low pressure cooling fluid supplied by the low pressure cooling system to other portions of the turbine airfoil cooling system.

The present invention relates to a turbine assembly (10) of a turbine engine (1), comprising at least: a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged in series, the rotors (12, 14) being mounted on a shaft (2); a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first recess (C1) arranged between the first rotor (12) and the stator (13), from a second recess (C2) arranged between the stator (13) and the second rotor (14); and pressure-reducing means (300, 31) positioned inside the first recess (C1), the assembly being characterised in that said pressure-reducing means (300, 31) comprise a plurality of substantially radial recompression fins (300) extending into the first recess (C1).

A rotor wheel for an engine includes a plurality of impeller vanes and a plurality of fluid passages defined by adjacent impeller vanes. The fluid passages are radially disposed across at least a portion of the rotor wheel. One or more impeller inserts may be disposed within one or more of the plurality of fluid passages, respectively. The impeller inserts define an impeller passage with a passage shape that controls a flow of fluid through the one or more of the plurality of fluid passages.

Disclosed herein is a gas turbine disk that includes a cooling target, and a disk unit having a main passage that is open to supply cooling air to the cooling target, and a plurality of unit passages that are open at an end of the main passage while each having a predetermined size.

A gas turbine has at least one rotor and inner housing part to form an annular chamber therebetween. The annular chamber is fluidically connected to a compressor portion at one end and expansion turbine portion at the other, and is supplied with cooling fluid. First and second swirl supply lines supply the annular chamber with cooling fluid. The cooling fluid is supplied to the surface of the rotor with a tangential flow component, and a first seal element in the annular chamber acts as a flow resistor. A discharge line in the rotor between the first seal element and expansion turbine portion receives and discharges cooling fluid from the second swirl supply line. No bypass lines are provided from the first swirl supply line such that the cooling fluid is conducted around the second swirl supply line in order to be returned to a location of the annular chamber.