A case for a gas turbine engine includes a case wall and a boss that extends from the case wall. The boss includes a perimeter step.

Turbine assemblies for a turbine of a gas turbine engine are disclosed herein. The turbine assemblies include a carrier and a blade track. The blade track extends around blades of a turbine wheel assembly to block gasses passed along a gas path from moving over the blades without causing the blades to rotate during operation of the turbine assembly. The blade track includes a plurality of blade track segments arranged circumferentially adjacent to one another about a central axis to form a ring. A retention ring engages both the carrier and the blade track segments to facilitate coupling and sealing.

An apparatus relating to a component for a turbine engine which generates a hot gas fluid flow, and provides a cooling fluid flow. The component having a wall separating the hot gas fluid flow from the cooling fluid flow and having a heated surface along which the hot gas fluid flow flows and a cooled surface facing the cooling fluid flow. At least one cooling hole including a connecting passage extending from a first inlet located at the cooled surface, to a first outlet located at the heated surface, the connecting passage defining a downstream flow direction from the first inlet to the first outlet, the connecting passage comprising a metering section fluidly coupled to the first inlet and defining a metered centerline and a diffusing section, downstream of the metering section, defining a diffused centerline, and fluidly coupling the metering section to the first outlet.

A propeller shaft assembly for an aircraft engine includes a shaft having a bore extending through the shaft at a front end thereof, the front end of the shaft having an outer surface facing radially outwardly from the shaft and an inner surface spaced apart from the outer surface and facing radially inwardly to the bore. The shaft has a front flange extending radially outwardly on the outer surface, the front flange having a base merging with the outer surface of the shaft. A sleeve is coupled to the shaft within the bore by an interference fit between the sleeve and the shaft, at least part of the sleeve axially aligned with the front flange. The sleeve axially extends from a front to a rear sleeve end, the rear sleeve end axially offset from the engine side surface of the front flange at the base of the front flange.

Variable bleed valves with struts for aerodynamic stability are described. An apparatus described herein includes a variable bleed valve port, and a strut disposed within the variable bleed valve port, the strut defining a plane and a plurality of holes disposed perpendicularly to the plane, the plurality of holes disposed in a pattern to dampen an acoustic response associated with the variable bleed valve port.

A sealing element for a turbomachine, in particular an aircraft engine, with a housing for the at least one rotating structural component is provided. The sealing element comprises a honeycomb structure for arrangement inside the housing, wherein the honeycomb structure extends in a first direction, wherein support structures are connected to the honeycomb structure in one piece and/or in a pattern-like manner and extend at least partially into the honeycomb structure, and the support structures extend at least partially or completely in a second direction that is different from the first direction. The support structures have planar portions, which are formed by at least one partially or completely closed cell together with the honeycomb structure.

A case for a gas turbine engine includes a case wall and a boss that extends from the case wall. The boss includes a perimeter step.

In a rotary machine, platforms are each in a shape of a flat surface extending in an axial line direction and each includes damper abutting surfaces extending to be close to each other as approaching the outer side in the radial direction while opposing each other in the peripheral direction in each of the platforms adjacent to each other, and the damper pin forms a regular polygonal prism extending in the axial line direction and includes a damper pin main body in which an angle formed by two side surfaces among a plurality of side surfaces corresponds to an angle formed by the damper abutting surfaces adjacent to each other.

A blade assembly for use in a gas turbine engine. The blade assembly includes a blade, a platform distinct from the blade and configured to extend around the blade, and a pin that couples the platform with the blade.

An igniter seal arrangement for a combustion chamber. The combustion chamber has a wall having first surface and second surfaces. A boss projects from the first surface, and has a platform on its remote end. The platform has an inner surface spaced from the first surface of the wall to define a chamber between the first surface of the wall and the inner surface of the platform. The platform has an outer surface facing away from the first surface of the wall and an aperture extends through the wall from the outer surface of the platform of the boss to the second surface. First and second L-shape rails extend from the platform and a sealing member has a first edge and a second edge locatable between the outer surface of the platform and the first and second L-shape rails and the sealing member has an aperture to receive an igniter.