A multi-stage turbine (16) is designed as an induction turbine with vapour induction in at least one intermediary stage. It is more particularly conceived as a radial-outward-flow type multi-stage turbine, with an axial main vapour inlet port (82) and an annular secondary vapour inlet port (84), which is arranged in the turbine (16) so as to annularly induce, in an intermediary stage of said turbine, a secondary vapour stream into an already partially expanded radial main vapour stream. The annular secondary vapour inlet port (84) comprises as a ring-zone (92) with through holes (94), which radially surrounds said axial main vapour inlet port (82) in a first turbine housing part (80). The axial vapour inlet port comprises a first tubular vapour inlet connection (82). The annular vapour inlet port comprises a second tubular vapour inlet connection (84) surrounding the first tubular vapour inlet connection (82), so as to define with the latter an annular space (86), wherein the ring-zone (92) with through holes (94) is arranged in this annular space (86).

A multi-stage turbine (16) is designed as an induction turbine with vapour induction in at least one intermediary stage. It is more particularly conceived as a radial-outward-flow type multi-stage turbine, with an axial main vapour inlet port (82) and an annular secondary vapour inlet port (84), which is arranged in the turbine (16) so as to annularly induce, in an intermediary stage of said turbine, a secondary vapour stream into an already partially expanded radial main vapour stream. The annular secondary vapour inlet port (84) comprises as a ring-zone (92) with through holes (94), which radially surrounds said axial main vapour inlet port (82) in a first turbine housing part (80). The axial vapour inlet port comprises a first tubular vapour inlet connection (82). The annular vapour inlet port comprises a second tubular vapour inlet connection (84) surrounding the first tubular vapour inlet connection (82), so as to define with the latter an annular space (86), wherein the ring-zone (92) with through holes (94) is arranged in this annular space (86).

A fixture assembly includes an inner diameter wall displaced from a main body by a first end wall with a convex surface and a second end wall with a concave surface. A method of machining a gas turbine engine component with an Abrasive Flow Media (AFM) process includes restricting a flow of media adjacent to an outer sidewall of an outer airfoil to be generally equal between each of a multiple of airfoils of the component.

A turbine rotor includes a base and a plurality of blades. A central nose is radially inward of the blades and defines an axis of rotation. A plurality of cooling manifolds is disposed within the turbine rotor and includes impingement cooling jets extending through a rear surface of the turbine rotor. An internal cooling manifold extends radially inward of the impingement cooling jets and extends between the base and the rear surface of the turbine rotor. A central nose cooling manifold extends into the central nose and is fluidically connected to the internal cooling manifold. A base cooling manifold is fluidically connected to the central nose manifold and extends radially outward from the central nose cooling manifold. A blade cooling manifold is fluidically connected to the base cooling manifold and extends within the blade. Trailing edge jets extend from the blade cooling manifold and through the trailing edge of blades.

An exhaust gas turbine is provided. The exhaust gas turbine includes a first turbine housing part having insulating material extending along an interior surface and a second turbine housing part having insulating material extending along an interior surface, the second turbine housing part coupled to the first turbine housing part to form a volute directing exhaust gas to a turbine wheel.

A compressor wheel for an internal combustion engine may include a plurality of blades which in a circumferential direction of the wheel are spaced from one another. The plurality of blades may respectively include an inflow edge which during operation are subject to an inflow of a compressible fluid substantially axially to the wheel axis. The plurality of blades may include a surface layer of locally distinct materials to adapt to locally distinct loads during operation.

A method of manufacturing a component of a rotary machine, the component has at least one inner passage that extends from a center up to a boundary surface of the component and is at least partly closed, and a blank is provided that includes the boundary surface and a top surface. The Method includes a first subtractive machining step that is carried out in which a part of the passage that at least includes an opening of the passage into the boundary surface as well as a cut-out in the top surface are manufactured by machining production, and subsequently the passage is completed by build-up production on the blank.

An assembly is provided for a turbine engine. This assembly includes a static structure and an impeller rotor housed within the static structure. The impeller rotor includes a vane structure and a shroud. The vane structure includes a first sidewall, a second sidewall and a plurality of vanes arranged circumferentially about a rotational axis. The vanes include a first vane. The first vane includes a first portion, a second portion and a third portion. The first portion is axially between the first sidewall and the second sidewall. The second portion is radially between the first sidewall and the shroud. The third portion is radially between the second sidewall and the shroud. The shroud circumscribes the vane structure. A gap is formed by and extends between the shroud and the static structure. A dimension of the gap changes as the gap extends along the shroud.

An impeller associated with a compressor of a gas turbine engine includes a plurality of impeller blades. Each impeller blade of the plurality of impeller blades has a leading edge and a trailing edge opposite the leading edge in a streamwise direction. Each impeller blade of the plurality of impeller blades extends in a spanwise direction from a hub at 0% span to a tip at 100% span, and each impeller blade of the plurality of impeller blades has a plurality of mean camber lines that each extend from the leading edge to the trailing edge at a respective spanwise location. The leading edge includes a partially swept leading edge surface defined between 70% span to 100% span that extends in the streamwise direction between 3% to 15% of a mean camber line at 100% span.

A turbocharger turbine wheel can include a hub that includes a rotational axis, a backdisk and a nose, where the rotational axis defines an axial coordinate (z) in a cylindrical coordinate system that includes a radial coordinate (r) and an azimuthal coordinate (0) in a direction of intended rotation about the rotational axis; and blades that extend outwardly from the hub, where each of the blades includes a hub profile, a shroud edge, a leading edge, a trailing edge, a pressure side, and a suction side, where the hub profile includes a global maximum radius and a global minimum radius, and where, between the global maximum radius and the global minimum radius, in an axial direction from the backdisk to the nose, the hub profile includes a local minimum radius at a first axial coordinate position and a local maximum radius at a second axial coordinate position.