An integrally bladed rotor, including: a plurality of blades integrally formed with a hub as a single component, each of the plurality of blades having a blade body extending from the hub to an opposed blade tip surface along a longitudinal axis, wherein the blade body defines a pressure side and a suction side, and wherein the blade body includes a cutting edge defined between the blade tip surface of the blade body and the pressure side of the blade body, wherein the cutting edge is configured to abrade a seal section of an engine case. A method for manufacturing an integrally bladed rotor includes: forming a plurality of airfoils integrally with a hub to form a single component, each of the plurality of airfoils having an opposed tip surface with respect to the hub extending along a longitudinal axis, wherein each of the plurality of airfoils defines a pressure side and a suction side; and forming a cutting edge between the tip surface and the pressure side of each of the plurality of airfoils, wherein the cutting edge is configured to abrade a seal section of an engine case.

A turbine rotor blade includes an airfoil portion having an airfoil defined by a pressure surface and a suction surface; and a squealer rib on a tip surface of the turbine rotor blade, the squealer rib extending from a leading-edge side (toward a trailing-edge side. The squealer rib has a ridge extending in an extending direction of the squealer rib. The turbine rotor blade is configured to provide a clearance between the tip surface of the turbine rotor blade and an inner wall surface of a casing of a turbine such that the inner wall surface of the casing of the turbine faces the tip surface of the turbine rotor blade and the clearance has a local minimum value on the ridge. The clearance is greater than the local minimum value at both sides of the ridge in a width direction of the squealer rib.

Baffles for gas turbine engines are provided. The baffles include a baffle body extending between a first end and a second end, a chamfered surface formed at at least one corner of the baffle body, wherein the chamfered surface extends from the first end to the second end, and a plurality of baffle holes formed in the chamfered surface.

A turbocharger (18) includes a shaft (20), a mixed flow turbine wheel (40) including a wheel hub (44) and blades having tips (42), and a heat shield (11). The heat shield (11) has a side wall, an end (19), and a contoured front edge (13) that connects the sidewall (16) and the end (19). The front edge (13) of the heat shield (11) defines a slope that forms an imaginary line that is angled relative to the sidewall (16) and the end (19), and intersects an axis of rotation of the shaft (20). The heat shield (11) resides at a position that is between a bearing housing (22) of the turbocharger (18) and the turbine wheel (40), and axially inward relative to an axially-facing surface of a turbine volute (24), and between the bearing housing (22) and the turbine wheel (40).

An axial flow rotating machine includes: a rotor; a plurality of vanes; and a medium flow modification member. Each of the vanes has an inner shroud and one or more seal fins. An annular groove, which is recessed toward a radially inner side, in which the inner shroud and the seal fins are placed in a non-contact manner, is formed at a rotor shaft. A distance from an end of the inner shroud on a furthest axially downstream side to a downstream-side groove side surface is a distance (L). A distance (Lf) from a most-downstream seal fin to a medium flow modification surface is equal to or less than the distance (L) in the axially downstream side.

A gas turbine engine for an aircraft includes: a flow path boundary, which delimits the flow path through the engine radially on the outside, and a lining, which lines the flow path boundary on the inside, at least along an axial section. Here, the lining includes a plurality of panels, which, in the circumferential direction of the flow path boundary, adjoin each other and which together line a circumferential area of 360, wherein each panel has two end faces, which each adjoin an end face of an adjacent panel. The panels are of beveled design at their end faces, such that two mutually adjoining panels form a V-shaped gap between them, the minimum clearance of which is realized at the inside of the panels. The panels can be sound-absorbing panels. Also disclosed is a panel for a gas turbine engine.

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.

A method for modifying an airfoil shroud located at a tip of an airfoil is provided. A locating step locates a reference location in a first end edge. The reference location includes a portion of a fillet. A seal rail extends circumferentially along a radially outer surface of the airfoil shroud. The fillet extends along the radially outer surface and is positioned directly adjacent to the seal rail. A forming step forms a compound relief cut in the fillet without performing a weld process on the airfoil shroud to remove the reference location. The compound relief cut has a first cut having a first radius extending from the seal rail towards the first end edge, and a second cut having a second radius extending from an end of the first cut to the first end edge. Modifying the airfoil shroud is complete following forming the compound relief cut.

A tip portion of a compressor blade faces a casing with a clearance therebetween. The tip portion has an upstream-side region including a blade leading edge, and a downstream-side region including a blade trailing edge. The upstream-side region has a small clearance formation portion including a part in which the clearance is a minimum in the tip portion. The downstream-side region forms a large clearance formation portion having a clearance larger than a clearance of the small clearance formation portion throughout the entire region of the downstream-side region.

A turbine blade, including: an airfoil and a platform, which has an upper face on which the airfoil is arranged; and at least one lateral face, the lateral face including a slot for insertion of a sealing strip. The transition between the upper face of the platform and the at least one lateral face includes a stepped portion and a beveled portion is provided.