A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall each includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at a joining line that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

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 variable capacity turbocharger includes a nozzle flow path which allows a gas to pass from a scroll flow path toward a turbine impeller, a shroud side ring and a hub side ring which face each other in a rotation axis direction of the turbine impeller and form a nozzle flow path therebetween, a bearing hole which is provided in the shroud side ring, a bearing hole which is provided in the hub side ring, and a nozzle vane which is disposed in the nozzle flow path and is supported by both bearing holes. A center axis line of the bearing hole is located on the inside in a radial direction in relation to a center axis line of the bearing hole at a room temperature, and the center axis line is located on the outside in the radial direction in relation to the center axis line during operation.

An exhaust baffle component for an aircraft starter includes a plurality of frames arranged with one another as a single component. At least one of the plurality of frames may include an attachment flange extending therefrom and defining a receiving aperture. Each frame of the plurality of frames may include one or more louvers stacked relative to one another and spaced from one another to define an air opening therebetween. The attachment flange may extend from the frame at a substantially central region of an upper portion of the frame to substantially align the receiving aperture with a starter aperture of a starter component to facilitate securement of the attachment flange to the starter component.

A shrouded impeller and a method of additively manufacturing the same are provided. In one example aspect, the shrouded impeller includes a hub and a shroud spaced from the hub. The shrouded impeller also includes a plurality of vanes extending between and connecting the hub and the shroud. The vanes are spaced circumferentially apart from one another. Flow passages are defined between adjacent vanes. In some implementations, the shrouded impeller is additively manufactured. During printing, one or more support structures are formed within and fill a portion of one or more of the flow passages to support the unsupported walls of the shrouded impeller, e.g., the shroud. Further, the support structures are removable from the shrouded impeller, e.g., after the shrouded impeller has been printed.

A method for the assembly of a counter-rotating radial turbine includes: preparing a central case; pre-assembling a first turbine unit and a second turbine unit, each including: a half-case delimiting a housing; a rotating unit with a shaft housed and rotatably supported in the housing and that is free to rotate about an axis of rotation with respect to said housing; a rotor disc joined to and overhanging a distal end of the shaft and having a front bladed face facing the opposite side with respect to the half-case. The method includes coupling the first pre-assembled turbine unit and the second pre-assembled turbine unit to the central case so as to arrange the front bladed face of the first rotor disc in front of the front bladed face of the second rotor disc, wherein, following the coupling process, the first half-case and the second half-case laterally close the central case.

The invention relates to a diffuser (5) of a radial or hybrid compressor (2) comprising an internal shroud (20), an external shroud (30) and a series of vanes (10), the trailing edge (12) of each vane (10) being curved in such a way that a distance (di) between the leading edge (11) and the trailing edge (12) of the vane (10) at the mid-height point (h/2) of the vane (10) is 5% to 15% shorter than a distance (d2) between the leading edge (11) and the trailing edge (12) of the vane (10) at the level of the internal shroud (20), the trailing edge (12) also having an evolving cross section comprising a first part forming a low wall (14) in the vicinity of the external shroud (20) and having a height (H) equal to at most 10% of the height (h) of the vane (10), a curved second part and a third part forming a low wall (14) substantially identical to the first part.

Circumferentially-split diffuser assemblies utilized within compression systems, such as centrifugal and mixed-flow compression systems employed within gas turbine engines, are provided. In embodiments, the diffuser assembly includes flow passages, which extend through the diffuser assembly and which include diffuser flow passage sections. Diffuser airfoils are interspersed with the diffuser flow passage sections. The diffuser airfoils include inboard and outboard airfoil segments distributed around a diffuser assembly centerline. The inboard and outboard airfoil segments are contained in and, thus, defined by inner and outer annular diffuser structures, respectively. The outer annular diffuser structure circumscribes the inner annular diffuser structure. In certain cases, the inboard airfoil segments and at least a portion of inner annular diffuser structure are composed of a first material, while the outboard airfoil segments and at least a portion of outboard annular diffuser structure are composed of a second material different than the first material.

A variable nozzle unit includes a variable nozzle vane which is disposed between a hub wall surface and a shroud wall surface and is rotatable around a rotation axis parallel to a rotation axis of a turbine impeller inside a gas flow passage. The variable nozzle vane includes a leading edge, a trailing edge, a hub end surface facing a hub wall surface, and a shroud end surface facing a shroud wall surface. The variable nozzle vane is twisted around a twist center located between the trailing edge and the rotation axis so that the hub end surface protrudes to a radial outside of a rotation axis in relation to the shroud end surface at the leading edge side and the hub end surface protrudes to a radial inside of the rotation axis in relation to the shroud end surface at the trailing edge side.

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.