A combustor assembly for a gas turbine engine includes a TOBI module that includes a TOBI housing. The TOBI housing has a slot. A vane includes a dovetail removably received within the slot. The TOBI housing includes an axially extending TOBI nozzle array. The TOBI housing includes a plenum. A cooling passage fluidly connects the plenum to the vane. The TOBI housing includes passages configured to provide cooling fluid to the vane. The TOBI housing includes a first passageway further connecting the TOBI nozzle to the plenum.

A rotor of a turbine cooled by compressed air supplied from a compressor of a gas turbine prevents combustion gas from flowing between platforms and guides compressed air discharged outside the platform toward the blade airfoil. The rotor includes a disk having an outer circumferential surface; a plurality of blade airfoils arranged around the disk; and a plurality of platforms coupled to the outer circumferential surface of the disk, each platform having an outer circumferential surface to receive a corresponding blade airfoil and side-facing surfaces which are spaced apart from each other on adjacent platforms of the plurality of platforms to form a cooling passage through which the compressed air flows outward in a radial direction of the rotor, the cooling passage including a bend in a rotational direction of the rotor. The bend directs the compressed air in a direction opposite to the rotational direction of the rotor.

The present disclosure relates to a rotor assembly for a gas turbine engine, the rotor assembly comprising a first rotor stage having a first disc portion with a peripheral first rim portion and a second rotor stage, the second rotor stage having a second disc portion with a peripheral second rim portion. The second rotor stage is axially adjacent and downstream of the first rotor stage and the second rim portion has an axial extension extending towards the first rim portion such that the axial extension of the second rim portion defines a rotor drum cavity between the first and second disc portions. The second rotor stage further comprises a drive arm extending within the drum cavity to the first disc portion, the drive arm being connected to the first disc portion by at least one connector. The drive arm divides the drum cavity into radially outer rim cavity portion and a radially inner main cavity portion. The rotor assembly further comprises a rim seal located between the axial extension of the second rim portion and the first rim portion, and a pressure equalisation path extending from the rim cavity portion to the main cavity portion.

A rotor blade for disposition in a hot gas duct of a turbomachine, the rotor blade having an airfoil and an inner platform radially inwardly thereof, the inner platform having a chamber formed therein which is radially outwardly bounded by an inner shroud of the inner platform and is axially bounded between a forward chamber wall and a rearward chamber wall of the inner platform. A forward sealing plate is disposed at the forward chamber wall and extends axially forwardly away from the chamber, and a rearward sealing plate is disposed at the rearward chamber wall and extends axially rearwardly away from the chamber. Furthermore, an axial partition is disposed in the chamber, which axial partition divides the chamber into a radially inner chamber wall region and a radially outer chamber wall region and which, when viewed in an axial cross section, extends into the forward chamber wall at the level of the forward sealing plate and into the rearward chamber wall at the level of the rearward sealing plate.

A turbine rotor blade root is additively manufactured and includes a shank having a radially extending chamber defined therein. A blade mount is at a radial inner end of the shank. The blade mount has a hollow interior defined therein with the hollow interior in fluid communication with the radially extending chamber. A lattice support structure is disposed within the hollow interior of the blade mount.

A blade coupling structure and a turbine system having the same securely couple a blade to a rotor disk. The blade coupling structure includes a root elastic member disposed between a root end of a blade root of the blade and an inner end of a coupling slot formed in the rotor disk; and a wedge member having a wedge body fitted between the root elastic member and the inner end of the coupling slot. The wedge member and the root elastic member press each other and press the root end so that the bade root is fixedly coupled to the coupling slot. A flat portion of the wedge member contacts the root elastic member, and an inclined portion of the wedge member facilitates the fitting of the wedge body. A wedge passage is formed in non-contact regions of the wedge body and passes a cooling fluid to the blade.

A turbine rotor blade is additively manufactured and includes an airfoil body including a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges. A shank is at a radial inner end of the airfoil body, and at least one angel wing extends laterally from at least one side of the shank. A coolant transfer passage is defined through the at least one angel wing. The coolant transfer passage fluidly couples a first wheel space portion defined between the shank and a first adjacent shank of a first adjacent turbine rotor blade and a second wheel space portion defined between the shank and a second adjacent shank of a second adjacent turbine rotor blade. The coolant transfer passage allows coolant to pass between wheel space portions of adjacent turbine rotor blades.

A sealing plate for sealing cooling passages in a cooled turbine blade of a gas turbine engine and directing and controlling air flow through cooling passages is provided.

A gas turbine engine includes a fan, a compressor, a combustor, and a turbine. The compressor compresses gases entering the gas turbine engine. The combustor receives the compressed gases from the compressor and mixes fuel with the compressed gases. The turbine receives the hot, high pressure combustion products created by the combustor by igniting the fuel mixed with the compressed gases. The turbine extracts mechanical work from the hot, high pressure combustion products to drive the fan and compressor.

An air seal may comprise an annular ring defined by at least a proximal surface, a distal surface, an aft side and a forward side. A channel may be disposed in the forward side of the air seal and/or the aft side of the air seal and may extend between the proximal surface and the distal surface. An additional channel extending from at least one of the forward side or the aft side may be disposed in the distal surface. The channel and the additional channel may be circumferentially in line. The channels may define a flow path for direction cooling air from a proximal side of the air seal to a distal side of the air seal. The radial channel may interface with the axial channel at an edge of the air seal.