A nozzle ring (10) for a radial turbine is described. The nozzle ring includes a disc-shaped main body (11) having a central opening (12) for passing a shaft there through. Additionally, the nozzle ring includes guide vanes (14) disposed circumferentially in a radially outer portion on a first surface (11A) of the main body (11). Two or more bores (17) are provided in a radially inner portion el of the main body (11). Further, a groove (171) is provided in the first surface (11A) of the main body (11), the groove (171) connecting at least two of the two of more bores (17).

The invention relates to an air conditioning system for an aircraft comprising a turbine engine comprising at least one radial turbine and a compressor, said radial turbine extending along a central axis (10) and comprising a wheel (20) equipped with vanes (21) mounted so as to be able to rotate about said central axis (10), a nozzle (30) arranged at the periphery of said wheel (20) comprising a plurality of variable-pitch blades (31) arranged around said central axis (10), and a volute (12) comprising an air inlet (12a) and an air outlet (12b) opening onto said plurality of variable-pitch blades (31), characterized in that said variable-pitch blades (31) are arranged around said central axis (10) between two nozzle casing walls (11a, 11b) delimiting an air passage section of said nozzle between them, each variable-pitch blade (31) further comprising means for axial displacement of the blade between the two casing walls (11a, 11b).

A gas turbine engine comprising: a combustor configured to initiate combustion; and a turbine comprising a stator vane ring defining a plurality of passageways between adjacent vanes; wherein at least one of the passageways is provided with a restrictor which defines a temporary gas washed surface for the stator vane ring and is configured to be ablated upon initiation of combustion to reveal an operational gas washed surface of the stator vane ring. A method of starting a gas turbine engine is also described.

A shroud configured to be disposed around an impeller of a centrifugal compressor, the shroud has a wall extending circumferentially around a central axis, the wall having an inner face oriented toward a gaspath and an outer face opposed to the inner face, a bleed slot defined in the wall and extending along at least a portion of a circumference thereof, the bleed slot defining a bleed direction from the inner face and away from the gaspath, the bleed direction at the inner face of the wall being either parallel to the central axis or oriented toward the central axis. A method of manufacturing a shroud is provided.

A turbomachine defining a flowpath therethrough at which a fluid is compressed is provided, in which the turbomachine includes a first compressor in serial flow arrangement upstream of a second compressor. The second compressor includes a port at the flowpath and is configured to receive at least a portion of the fluid from the flowpath from the second compressor. The first compressor includes a vane positioned at the flowpath and the vane includes an opening at the flowpath configured to egress the portion of the fluid from the port into the flowpath.

A system includes a divided volute turbocharger having variable turbine geometry. The turbocharger includes a turbine housing, first and second volutes separated by a wall having a first and second tongue, and a turbine housing outlet. The system also includes a turbine wheel disposed in the turbine housing and a vane ring disposed in the turbine housing between the turbine wheel and the volutes. The vane ring includes a plurality of rotatable vanes disposed on a vane ring surface of an annular disk in an asymmetric or asymmetric vane pattern that receives the turbine wheel therewithin and a pair of aerodynamic spacers spaced circumferentially outward of the plurality of vanes with each spacer positioned adjacent to a respective one tongue, with the spacers and vanes directing and controlling the flow of exhaust from the volutes into the turbine wheel with generally equal flow while significantly reducing HCF forcing function.

A compressor of a gas turbine engine includes an impeller having a plurality of impeller blades. The compressor includes a diffuser downstream from the impeller that has a plurality of diffuser blades. Each diffuser blade extends from a hub to a shroud in a spanwise direction, and a leading edge of each diffuser blade is spaced apart from an impeller trailing edge of each of the plurality of impeller blades by a vaneless gap. Each diffuser blade includes a cutback region that extends from proximate the leading edge toward a trailing edge. The cutback region reduces a thickness of each of the diffuser blades such that a throat area defined between adjacent diffuser blades increases in the spanwise direction from the hub to the shroud and the vaneless gap increases in the spanwise direction from the hub to the shroud.

Flow control devices and structures designed and configured to improve the performance of a turbomachine. Exemplary flow control devices may include various flow guiding channels, ribs, diffuser passage-width reductions, and other treatments and may be located on one or both of a shroud and hub side of a machine to redirect, guide, or otherwise influence portions of a turbomachine flow field to thereby improve the performance of the machine.

The invention relates to an air conditioning system for an aircraft comprising a turbine engine comprising at least one radial turbine and a compressor, said radial turbine extending along a central axis (10) and comprising a wheel (20) equipped with vanes (21) mounted so as to be able to rotate about said central axis (10), a nozzle (30) arranged at the periphery of said wheel (20) comprising a plurality of variable-pitch blades (31) arranged around said central axis (10), and a volute (12) comprising an air inlet (12a) and an air outlet (12b) opening onto said plurality of variable-pitch blades (31), characterized in that said variable-pitch blades (31) are arranged around said central axis (10) between two nozzle casing walls (11a, 11b) delimiting an air passage section of said nozzle between them, each variable-pitch blade (31) further comprising means for axial displacement of the blade between the two casing walls (11a, 11b).

A radial inflow turbine includes a scroll flow passage, a turbine wheel disposed radially inward of the scroll flow passage, a plurality of variable nozzle vanes disposed on a flow passage extending from the scroll flow passage toward the turbine wheel, at a radial position between the scroll flow passage and the turbine wheel, a nozzle mount rotatably supporting each of the plurality of variable nozzle vanes, a nozzle plate arranged to face the nozzle mount and forming the flow passage with the nozzle mount, and a swirl generating member disposed, radially outward of the plurality of variable nozzle vanes, on the nozzle plate in a height range which is smaller than that of a vane height of each of the plurality of variable nozzle vanes. A position of an end part of the swirl generating member on a side of the nozzle mount is farther away from the nozzle mount than a position of an end part of each of the plurality of variable nozzle vanes on the side of the nozzle mount in an axial direction.