Embodiments of a stage for a turbomachine have been provided herein. In some embodiments, a stage for a turbomachine may include a bearing having a housing, the bearing defining an interior cavity; an outer ring disposed radially outward from the housing; and a plurality of airfoils disposed between the housing of the bearing and the outer ring.

Seals for gimbaling and/or fixed rocket engine nozzles, and associated systems and methods are disclosed. A representative rocket propulsion system includes a rocket engine having an exhaust nozzle, a seal plate carried by the exhaust nozzle, and a seal engaged with the seal plate. The seal includes at least one support, multiple pivotable first flaps, carried by the at least one support and positioned to contact the seal plate, and multiple pivotable second flaps, with an individual second flap positioned to shield a corresponding individual first flap. At least one forcing element is operatively coupled to at least one of the individual first flap or the individual second flap, to apply a pivoting force to the at least one of the individual first flap or the individual second flap.

A TOBI for a gas turbine engine having a TOBI body, a first TOBI airfoil having a radially extending portion extending from a leading edge and an axially extending portion extending toward a trailing edge, and a second TOBI airfoil circumferentially adjacent to the first TOBI airfoil, the second TOBI airfoil having a radially extending portion extending from a leading edge and an axially extending portion extending toward a trailing edge. An entrance is defined between the leading edges of the adjacent TOBI airfoils and an exit is defined between the trailing edges of the TOBI airfoils, wherein airflow entering the entrance enters in a radial direction relative to the TOBI body and airflow exiting the exit exits in a circumferential direction relative to the TOBI body.

Various embodiments include a steam turbine drum nozzle, along with a related assembly and steam turbine. Particular embodiments include a nozzle having: an airfoil; a radially inner sidewall coupled with a first end of the airfoil; and a radially outer sidewall coupled with a second end of the airfoil, the second end opposing the first end, wherein the radially outer sidewall includes: a first section radially outward of the airfoil; a thinned section coupled with the first section; and a second section coupled with the thinned section radially outward of the airfoil, the second section having a radially outer face and a circumferentially facing side abutting the radially outer face, wherein the second section includes a circumferentially extending slot, and wherein the second section includes a relief slot extending into a body of the second section from the circumferentially facing side.

A thermostructural composite material part including carbon or ceramic fiber reinforcement densified by a matrix having at least one thin portion in which: the thickness of the part is less than 2 mm, or indeed less than 1 mm; the fiber reinforcement is made as a single thickness of multilayer fabric made of spread yarns having a weight of not less than 200 tex; the fiber volume ratio lies in the range 25% to 45%; and the ratio between the number of layers of the multilayer fabric and the thickness in millimeters of the part is not less than four.

A variable displacement exhaust turbocharger is provided with a plurality of nozzle vanes that are rotatably supported by a nozzle mount, a drive ring that is interlocked with an actuator and fits to a guide part of the nozzle mount, a lever plate having one end coupled to the drive ring and the other end coupled to the nozzle vane, and a variable nozzle mechanism that varies the vane angle of the nozzle vane by rotating the drive ring. A bulge suppressing portion is provided for absorbing bulging occurring on an outer peripheral surface of the guide part, the bulging being caused by the press-fitting of a nail pin into a press-fitting hole in a direction slightly toward the rotational axis with respect to the guide section and following the axial direction of the nozzle mount.

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 vane arc segment includes an airfoil fairing that has first and second platforms and an airfoil section. The platforms have first and second openings that open into the airfoil section. The platforms each define first and second circumferential mate faces, forward and aft sides, a gaspath side, a non-gaspath side. The first platform has a first flange that projects radially from the non-gaspath side aft of the first opening and a second flange that projects radially from the non-gaspath side forward of the first opening. The first and second flanges are exclusive flanges on the first platform. The second platform has a third flange that projects radially from the non-gaspath side aft of the second opening.

The invention relates to a distributor for a turbomachine radial turbine, comprising an annular grill (26) extending about a central axis (10) and comprising a plurality of variable-pitch blades (31), defining between them an air passage cross section, characterized in that each blade is rotatably mounted about a pivot shaft (32), itself moveable in a translation direction, comprising at least one radial component, such that said blade may, upon actuation of control means (40), be pivoted about the pivot shaft and/or moved in relation to the central axis in said translation direction so as to be able to modify the air passage cross section by respectively controlling the metal angle (α3) and the radial spacing (ΔR).

A flowpath apparatus for a gas turbine engine includes: a plurality of ducts arranged in an array, each duct including a peripheral wall structure having a closed perimeter that defines a flow channel from an upstream end to a downstream end thereof; and a support structure positioning a the plurality of ducts in an array configuration.