Abstract: Embodiments of the invention are directed to a rotor for use in molten metal and devices including the rotor. The rotor has a rotor body and blades, wherein each blade includes a tip that is at least twice as hard as the rotor body.

Abstract: A shrouded conduit is provided for arranging, for example, in a gas path of a turbine engine. The shrouded conduit includes a tubular shroud extending longitudinally along a centerline. The shrouded conduit also includes a fluid conduit extending longitudinally in the shroud. A first portion of the fluid conduit is connected laterally to and may be formed integral with a first portion of the shroud.

Abstract: A steam turbine rotor blade includes a protrusion portion (7) which is provided on a tip end portion of a blade body (61), on which a leading edge portion (61a) is formed, in a blade height direction and protrudes from a suction-side surface (612) toward the leading edge portion (61a) side, and a transition-region seal member which is provided so as to cover at least a portion of a base end-side surface of the protrusion portion (7) and a leading edge-side transition region, which faces the leading edge portion (61a) side, of a connection portion between the protrusion portion (7) and the suction-side surface (612), the transition-region seal member being formed of a material having a hardness higher than that of the blade body (61).

Abstract: A ceramic matrix composite (“CMC”) center body may be positioned around an austenitic nickel-chromium-based superalloy attachment ring. The attachment ring may be integrally formed with a turbine engine case. The attachment ring may have a greater coefficient of thermal expansion than the center body. A plurality of pins may be inserted through apertures in the center body and coupled to the attachment ring. The pins may slide within the apertures, allowing the attachment ring to expand without applying a load on the center body.

Abstract: A compressor includes a rotor platform; a rotor blade; and a casing having an inner surface surrounding the tip and spaced radially outwardly from the tip to define a gap. A secondary air flow system includes a bleed inlet configured to remove secondary air flow from the primary air flow; an injection opening disposed in the inner surface of the casing upstream of the bleed inlet; an accessory conduit; and a plenum fluidly coupled to the bleed inlet, the injection opening, and the accessory conduit. The bleed inlet and plenum at least partially define a secondary air flow path such that a first portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the injection opening and a second portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the accessory conduit.

Abstract: Rotor assembly apparatus are disclosed. An example rotor assembly includes a twist actuator to drive a first rotation of a first shaft about a first axis, the twist actuator positioned at a center of rotation of the rotor assembly. A first gear assembly to convert the first rotation into a plurality of second rotations of a plurality of second shafts. Each of the second shafts to provide torque to a respective blade coupled to the rotor assembly.

Abstract: A stator vane support with anti-rotation features is provided. The stator vane support may comprise an inner diameter surface opposite an outer diameter surface. The stator vane support may comprise an anti-rotation lug defining a protrusion extending inward from the inner diameter surface. The stator vane support may have a first recess defining a first void on the inner diameter surface proximate a first surface of the anti-rotation lug. The stator vane support may have a second recess defining a second void on the inner diameter surface proximate a second surface of the anti-rotation lug. The anti-rotation lug may be configured to interface with a stator vane to at least partially limit circumferential movement, and each recess may be configured to allow the stator vane to thermally expand during gas turbine engine operation.

Abstract: Disclosed is a ducted and balanced wind turbine, including: a spindle, a front cross bearing bracket, a radial magnetic levitation bearing, a cross bracket, an outer rotor vortex blade, a turbine shell, an outer rotor rotating body, an outer rotor armature coil, a conductive slip ring, an axial magnetic levitation bearing cross bracket, an axial magnetic leverage bearing, a rear cross bearing bracket, a spindle rolling bearing, an output wire, a carbon brush set, a permanent magnet, an inner rotor rotating body, an inner rotor vortex blade, an outer rotor dome, and a spindle dome. The radial and axial magnetic levitation devices and the carbon brush set are mounted on the inner wall of the turbine shell, forcing the outer rotor rotating body to rotate freely in the turbine shell through the magnetic levitation bearings.

Abstract: A bladed rotor comprising a rotor disk (12) presenting two front faces (14, 15) and an outer peripheral face (16), sockets (18) being provided in the outer peripheral face (16) and opening out into at least one of the front faces (14, 15). The rotor (10) comprising blades (30), each having a root (32) whereby the blade is fastened in a socket (18), an end face (31) of the root being substantially level with the front face (14) of the disk when the blade is fastened in the socket. A coating layer (40) is deposited on the disk (12) so as to cover both at least a portion of the front face (14) of the disk and at least a portion of the end face (31) of the root (32).

Abstract: An improved system, apparatus and method for controlling vane angles in a gas turbine engine, and more specifically, for correcting vane angle error in a gas turbine engine. An active synchronization ring comprises a plurality of micro-actuators coupled to the synchronization ring to correct distortion in the synchronization ring. The micro-actuators apply a bending moment to the synchronization ring to cancel or compensate for synchronization ring distortion. The micro-actuators may be controlled open loop or closed loop. Strain sensors measure ring distortion and provide signals to the controller for closed loop control.

Abstract: A centrifugal compressor includes an impeller and a casing, the casing includes a scroll part forming a scroll flow passage and a diffuser part forming a diffuser flow passage. The diffuser part includes: a first diffuser portion belonging to a first angular range in a circumferential direction of the impeller; and a second diffuser portion belonging to a second angular range downstream of the first angular range in a flow direction of the scroll flow passage, of the angular range in the circumferential direction of the impeller, the second diffuser portion having an outer radius R2 which is defined along a reference circle centered at a rotational center of the impeller. An outer radius R1 of the first diffuser portion in the first angular range is smaller than the outer radius R2 of the second diffuser portion in the second angular range.

Abstract: A rotor assembly of a gas turbine engine may be spoked and includes a rotor and a shell. The rotor has a rotor disk and a plurality of blades each having a platform attached to the rotor disk and with a first channel defined radially between the platforms and the rotor disk. The shell projects aft of the rotor and includes inner and outer walls with a passage defined therebetween. The passage is in fluid communication with the first channel and, together, form part of a secondary flowpath for cooling of adjacent components. The rotor assembly may further include a structure located radially inward of the rotor disk and shell. The structure defines a supply conduit for flowing air from the passage and into a rotor bore defined at least in part by adjacent rotor disks. The entering air, being pre-heated when flowing through the channel and passage, warms the bore and reduces thermal gradients, thus thermal fatigue, across the rotor disk.

Abstract: A turbine center frame for a gas turbine having a plurality of first components and a plurality of second components alternately mutually adjacently disposed in circumferential direction and bounding a hot gas-conducting flow channel in radial direction, the first components each having two first overlapping portions, and the second components each having two second overlapping portions. In a transition from a first component to a second component, one of the first overlapping portions and one of the second overlapping portions overlap. A centering element, against which, the first components and the second components are braced along the circumferential direction and in the radial direction, centers the first components and the second components relative to a central axis of the turbine center frame.

Abstract: A structure configured to rotate about an axis may include a fastener opening defined in a surface of the structure. The fastener opening may extend through the surface and include a circular portion, a first slot portion, and a second slot portion. The circular portion may include a first radius of curvature. The first slot portion may extend from the circular portion circumferentially relative to the axis. The second slot portion may also extend from the circular portion circumferentially relative to the axis.

Abstract: A nozzle assembly includes a first nozzle cavity, a second nozzle cavity, a rib positioned between the nozzle cavities, a rib cap positioned on the rib, a first cavity insert, and a second cavity insert. The rib cap is wider than the rib, such that the rib cap extends outwardly into those portions of the nozzle cavities immediately adjacent the rib. The first cavity insert and the second cavity insert include a longitudinal surface offset from the rib. The offset surfaces include rib cap-interface surfaces that are joined to the rib cap. A method of modifying the nozzle assembly is achieved by installing a rib cap having a width greater than that of the rib and by installing modified cavity inserts having a surface offset from the rib, and by joining rib cap-interface surfaces of the modified cavity inserts to the rib cap.

Abstract: A multi-spool gas turbine engine comprises a low pressure (LP) spool and a high pressure (HP) spool. The LP spool and the HP spool are independently rotatable about an axis. The LP pressure spool has an LP compressor and an LP turbine. The HP spool has an HP turbine and an HP compressor. An accessory gear box (AGB) is axially mounted at one end of the engine. The LP compressor is axially positioned between the HP compressor and the AGB. The AGB is drivingly connected to the HP spool through the center of the LP compressor.

Abstract: Aspects of the disclosure are directed to an inlet cap, comprising: a first lug slot located proximate a perimeter of the inlet cap, and a second lug slot located proximate the perimeter of the inlet cap and located peripherally adjacent to the first lug slot, where the first lug slot includes a first ply, the second lug slot includes a second ply that is different from the first ply, and at least a third ply is common to the first lug slot and the second lug slot.

Abstract: A nacelle for an aircraft propulsion system includes a core cowl portion, a bifurcation portion and an extension portion. The core cowl portion extends about a centerline to the bifurcation portion. The bifurcation portion is connected to and extends radially between the core cowl portion and the extension portion. The extension portion projects out from the bifurcation portion and circumferentially extends over the core cowl portion. The extension portion includes an acoustic panel and a structural reinforcement. The acoustic panel includes a cellular core between a perforated face skin and a back skin, wherein the face skin is radially inboard of the back skin. The structural reinforcement is bonded to the back skin and structurally reinforces the acoustic panel.

Abstract: Aspects of the disclosure are directed to a seal comprising: a first fitting configured to couple to a first disk, and a curvic joint including curvic teeth, where a first distance between the first fitting and the curvic teeth is equal to or greater than a first thickness of the first disk. In some embodiments, an engine comprises: a compressor disk, a turbine disk, and a seal including: a first fitting configured to couple to the turbine disk, a second fitting configured to couple to the compressor disk, and a curvic joint including curvic teeth, where a first axial distance between the first fitting and the curvic teeth is greater than or equal to a first radial thickness of the turbine disk.

Abstract: A shear web for a wind turbine blade is described. The shear web comprises a panel and a web head. A longitudinal edge of the panel is received within a slot defined in the web head. A plurality of discrete spring features are attached to the longitudinal edge of the panel. The spring features are mutually spaced apart at intervals along the length of the longitudinal edge. The spring features compress against a base of the slot when the longitudinal edge region of the panel is inserted into the slot.