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 gas turbine spacer disk includes a disk portion, a rim portion, a first fillet, and a second fillet. The disk portion is disposed about a rotational axis. The rim portion is disposed about the disk portion. An outer face of the rim portion defines a plurality grooves extending circumferentially about the rotational axis. The first fillet transitions from the rim portion to a first side of the disk portion. The second fillet transitions from the rim portion to a second side of the disk portion. The plurality of grooves includes a pair of first grooves having a first diameter and a pair of second grooves having a second diameter that is less than the first diameter. A first one of the first grooves overlaps in an axial direction with the first fillet. A second one of the first grooves overlaps in the axial direction with the second fillet.

When a closed type impeller is manufactured from one block, a rough cutting process of cutting a flow path region of a block using a rough cutting tool, and a residue-cutting process of cutting a cutting residue in the cutting process using a residue-cutting tool are executed. The residue-cutting tool has a tool main body having a blade formed on an outer periphery thereof, and a handle having the tool main body fixed to a distal end thereof. A maximum outer diameter of the tool main body is larger than a minimum outer diameter of the handle. Further, the tool main body has a rear blade directed in a direction including a tool rear side component.

An assembly of turbine blades or vanes includes a first and a second airfoil extending span-wise from a first and a second platform respectively. The first and the second platform respectively have a first and a second mate face that interface along a platform splitline. The first mate face is proximal to a suction side of the first airfoil the second mate face is proximal to a pressure side of the second airfoil. The first mate face is chamfered or filleted along an aft portion thereof. The chamfered or filleted portion of the first mate face lies in a region in a flow path between the first and second airfoils where a mean velocity of the working medium is directed from the second platform to the first platform.

A rotor disk for an exhaust turbocharger is mounted in a housing of the exhaust turbocharger able to rotate about an axis of rotation. The rotor disk has a disk hub comprising a disk back and a disk front remote from the disk back. A plurality of rotor disk blades are formed on the disk hub in a manner extending between the disk back and the disk front. A balancing mark is arranged in a blade channel formed between a first blade of the plurality of rotor disk blades and a second blade, arranged adjacent to the first blade, of the plurality of rotor disk blades. A width of the balancing mark is less than a length of the balancing mark. The disclosure also relates to an exhaust turbocharger comprising such a rotor disk and to a method for balancing a rotor assembly for such an exhaust turbocharger.

A shrouded rotor constructed through a hybrid additive manufacturing process. The shrouded rotor including a hub at a radial center having an outer surface forming an inner wall of a flow path, a shroud at a radial outer side having an inner surface forming an outer wall of the flow path, and vanes extending within the flow path between the hub and the shroud. The outer surface of the hub, the inner surface of the shroud, and all surfaces of the vanes have a surface roughness average of less than 32R.sub.a.

In a turbine rotor blade assembly 1 of the present invention, each turbine rotor blade 10 includes a platform 11 having a blade root 12 fixed to a turbine disk 30, a profile 13 rising from the platform 11, and a shroud 14 provided at a top end of the profile 13. The shroud 14 of the present invention includes a first contact end part 15 that comes into contact with an adjacent shroud adjacent to one end side in a circumferential direction, a second contact end part 16 that comes into contact with an another adjacent shroud adjacent to the other end side in the circumferential direction, and a main body part disposed between the first and second contact end parts 15 and 16. One or both of the first and second contact end parts 15 and 16 are lower in rigidity than the main body part.

A hybrid additive manufacturing process is utilized for creating a shrouded rotor with the shrouded rotor having a hub at a radial center, a shroud at a radial outer side, and vanes extending therebetween. The hybrid additive manufacturing process includes forming the shrouded rotor in stages, with a first stage being formed by depositing material in an axial direction through a first stage of the hub, machining an outer surface of the first stage of the hub to smooth the outer surface, depositing material on the first stage of the hub in a radial direction through a first stage of the vanes and the shroud, and machining all surfaces of the first stage of the vanes and an inner surface of the first stage of the shroud to smooth the surfaces. Subsequent stages of the shrouded rotor are formed similarly to the first stage.

The present invention enables machining to be easily performed and reduces stress concentration on a machined hole that is formed in a cylindrical member. Recesses (2) recessed in the depth direction of a machined hole (1) are formed on circumferential side-portions of the machined hole 1 formed in a cylindrical member (10). In each of the recesses (2), a part of the opening edge is formed to be a circular arc portion (2a) that has a circular arc shape, the bottom is formed to be gradually shallowed by an inclined surface (2c) toward an opened portion (2b) in which the circular arc shape is opened, from a part along the circular arc portion (2a), and the circular arc portion (2a) is disposed toward the machined hole (1).

A method for restoring a blade or vane platform of a gas turbine assembly configured for a power plant by: providing a blade or a vane having a platform with an edge deterioration zone; removing the deterioration zone electro discharging machining technology; and rebuilding a removed zone by additive manufacturing technology. The removing can be performed to create a recessed plane along a platform edge, the recessed plane being connected to a platform plane by an enter inclined plane and an exit inclined plane arranged opposed along the platform edge.