A tangential on-board injector (TOBI), comprising: a body defining an annular passageway to receive cooling air, the TOBI defining a plurality of discharge nozzles; a rotating component mounted for rotation relative to the body about an axis of rotation; a seal extending between the body and the rotating component; a plurality of vanes circumferentially distributed about the axis of rotation and located downstream of the plurality of discharge nozzles relative to a flow of the cooling air circulating toward the seal from the plurality of discharge nozzles and upstream of the seal; and flow passages defined between the plurality of vanes, a flow passage of the flow passages extending along a passage axis, the passage axis having a tangential component at an outlet of the flow passage that is different than a tangential component of an exit flow axis of a nozzle of the plurality of discharge nozzles.

A blade for use in a gas turbine engine, the blade comprising a blade portion and a fir tree root portion, the blade portion and the root portion having a connected passage for allowing cooling air to flow within the blade, the fir tree root portion having an air intake on its leading edge, the air intake allowing cooling air to enter the cooling passage and wherein the fir tree root portion comprises a plurality of projections, including at least a base projection and a top projection; and wherein the air inlet located in the base projection of the fir tree root portion and wherein the air inlet comprises at least 50% of the face of the base projection of the fir tree root portion.

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.

A gas turbine engine, has: a compressor; a turbine having a rotor; and an inertial particle separator located upstream of the turbine downstream of the compressor, the inertial particle separator having: an intake conduit in fluid flow communication with the compressor and defining an elbow, a splitter, a leading edge of the splitter located downstream of the elbow, the splitter located to divide a flow into a particle flow and an air flow, and an inlet conduit and a bypass conduit located on respective opposite sides of the splitter, the inlet conduit receiving the air flow, the inlet conduit in fluid flow communication with a cavity containing the rotor for cooling the rotor of the turbine section, the bypass conduit receiving the particle flow, the bypass conduit in fluid flow communication with an environment outside the gas turbine engine while bypassing the cavity containing the rotor.

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.

A turbine for a turbomachine has a first rotor and a second rotor configured to pivot about an axis in two opposite directions of rotation. The first rotor includes a radially outer drum from which blades extend radially inwards. The turbine has cooling means attached relative to the outer drum. The cooling means include a support plate with at least one first orifice and a calibration plate that is attached to the support plate and located radially inside the support plate. The calibration plate has at least one second calibration orifice facing the outer drum to allow the passage of cooling air radially from the outside to the inside through the first and second orifices.

A gas turbine engine having a pre-swirler adjustability is presented. The pre-swirler includes a pre-swirler insert installed in a component enclosed by a cover. The component includes an inner compressor exit diffusor enclosed by an outer casing or a shaft cover enclosed by the inner compressor exit diffusor. The pre-swirler is adjustable by replacing the pre-swirler insert. An access port including an access window is arranged on the cover. The access port gives access to the pre-swirler insert for replacement through the access window. The access window includes a manhole or combustor assembly installation hole on the outer casing, or a cutout on the inner compressor exit diffusor. The access port allows adjusting the pre-swirler by replacing the pre-swirler insert installed in the component without lifting the cover enclosing the component.

A moving blade of a turbo machine, having a leading edge, a flow trailing edge, flow conduction faces, and a blade root mounting the moving blade to a hub body. The blade root is fir tree-like with projections spaced apart from one another. An inner shroud is arranged between the blade leaf and the blade root. A cooling passage is integrated in the blade leaf and the blade root for a cooling medium. An inlet of the cooling passage is formed on the blade root formed of a first inlet passage portion and a second inlet passage portion and a material web extends there between. The first inlet passage portion and the second inlet passage portion merge into a unifying passage portion arranged radially outside the uppermost projection of the blade root and radially inside of the inner shroud.

A turbine assembly for use with a gas turbine engine includes a bladed wheel assembly, a vane assembly, and an inner seal. The bladed wheel assembly is adapted to interact with gases flowing through a gas path of the gas turbine engine. The vane assembly is located upstream of the bladed wheel assembly and adapted to direct the gases at the bladed wheel assembly. The inner seal is configured to block gases from passing around the vane assembly.

A damper seal for a turbine blade of a gas turbine engine, the damper seal having: an upper portion; a first downwardly curved portion; and a second downwardly curved portion, the first downwardly curved portion and the second downwardly curved portion extend from opposing end regions of the upper portion, the upper portion having a length extending between the opposing end regions of the upper portion and a width transverse to the length, wherein the upper portion is curved along the entire width as it extends along the length.