Abstract: There is described a core for an investment casting process, comprising: a core passage which extends between a first point and a second point along a tortuous path having length L, wherein the first point and second point are separated by a direct line of sight distance, S, wherein L is greater than S; and, a core bridge which extends between the first and second points away from the core passage.

Abstract: A gas turbine, including: axially adjacent first and second components each having a gas path facing surface within an engine casing, inter-component cavity located between first and second components on the gas path outboard side, an axial-restrictor within the inter component cavity for restricting movement of the first or second component relative to at least one other first and second component or engine casing, wherein the axial-restrictor is radially located in the engine casing or other first and second component and includes a radially facing surface; a retaining element, including a body having circumferential length extending around the gas turbine engine and axial length extending between axially adjacent first and second components when in use; an obstructing portion between the radially facing surface of the axial-restrictor and opposing second radial surface of the other first or second component or engine casing to restrict radial movement of in use axial-restrictor.

Abstract: Described is an abradable component for a gas turbine engine that includes a base having an outboard side which receives a supply of cooling air in use and an inboard side with a plurality of walls thereon. The walls intersect one another to define an abradable network of open faced cells on a gas washed surface thereof, and at least one of the walls includes one or more through-holes for providing a flow of cooling air from the outboard side to the gas washed surface of the abradable network of open faced cells, when in use.

Abstract: Described is a shroud arrangement for a gas turbine engine, comprising: a seal segment for bounding a hot gas flow path within the gas turbine engine, the seal segment being attached to a casing of the engine via at least one fore attachment and at least one aft attachment, the fore and aft attachment restricting radial movement of the seal segment relative to the engine casing; and, an axial restrictor which prevents axial movement of the seal segment relative to the engine casing, the axis being the principal rotational axis of the engine, wherein the fore and aft retention features are slidably engaged with a carrier segment or engine casing from a common direction. Also described is a gas turbine engine having the shroud arrangement.

Abstract: A seal segment of a shroud arrangement for bounding a hot gas flow path within a gas turbine engine, including: a plate having an inboard hot gas flow path facing side and an outboard side; a bulkhead extending from the outboard side of the plate which defines a fore portion and an aft portion; a first cooling circuit within the plate for cooling a first portion of the plate; a second cooling circuit within the plate for cooling a second portion of the plate; wherein the first cooling circuit is in fluid communication with the fore portion and the second cooling circuit is in fluid communication with the aft portion and the first and second cooling circuits are fluidically isolated from one another. Also described is a method of cooling a seal segment in a gas turbine engine.

Abstract: There is described a core for an investment casting process, comprising: a core passage which extends between a first point and a second point along a tortuous path having length L, wherein the first point and second point are separated by a direct line of sight distance, S, wherein L is greater than S; and, a core bridge which extends between the first and second points away from the core passage.

Abstract: A gas turbine, including: axially adjacent first and second components each having a gas path facing surface within an engine casing, inter-component cavity located between first and second components on the gas path outboard side, an axial-restrictor within the inter component cavity for restricting movement of the first or second component relative to at least one other first and second component or engine casing, wherein the axial-restrictor is radially located in the engine casing or other first and second component and includes a radially facing surface; a retaining element, including a body having circumferential length extending around the gas turbine engine and axial length extending between axially adjacent first and second components when in use; an obstructing portion between the radially facing surface of the axial-restrictor and opposing second radial surface of the other first or second component or engine casing to restrict radial movement of in use axial-restrictor.

Abstract: Described is an abradable component for a gas turbine engine, comprising: a base having an outboard side which receives a supply of cooling air in use and a plurality of walls on an inboard side thereof, the walls adjoining one another to provide an abradable network of open faced cells at a gas washed surface thereof; wherein at least one wall includes one or more through-holes for providing a flow of cooling air from the outboard side to the gas washed surface of the abradable network of open faced cells, when in use.

Abstract: A seal segment of a shroud arrangement for bounding a hot gas flow path within a gas turbine engine, including: a plate having an inboard hot gas flow path facing side and an outboard side; a bulkhead extending from the outboard side of the plate which defines a fore portion and an aft portion; a first cooling circuit within the plate for cooling a first portion of the plate; a second cooling circuit within the plate for cooling a second portion of the plate; wherein the first cooling circuit is in fluid communication with the fore portion and the second cooling circuit is in fluid communication with the aft portion and the first and second cooling circuits are fluidically isolated from one another. Also described is a method of cooling a seal segment in a gas turbine engine.

Abstract: Described is a shroud arrangement for a gas turbine engine, comprising: a seal segment for bounding a hot gas flow path within the gas turbine engine, the seal segment being attached to a casing of the engine via at least one fore attachment and at least one aft attachment, the fore and aft attachment restricting radial movement of the seal segment relative to the engine casing; and, an axial restrictor which prevents axial movement of the seal segment relative to the engine casing, the axis being the principal rotational axis of the engine, wherein the fore and aft retention features are slidably engaged with a carrier segment or engine casing from a common direction. Also described is a gas turbine engine having the shroud arrangement.

Abstract: An interstage seal is provided for installation between forward and aft stage discs of a turbine section of a gas turbine engine. The seal has a forward annular portion, which extends forward to contribute to an air seal at the outer periphery of the forward stage disc, and an aft annular portion, which extends rearward to contribute to an air seal at the outer periphery of the aft stage disc. The seal has a peripheral slot into which static vane inner rails are loadable. The seal includes a first component, which provides the forward annular portion, and a second component, which provides the aft annular portion. The second component is separably connectable to the first component such that the first and second components, when connected, cooperate to form the peripheral slot. Loading the static vane inner rails into the slot prevents the second component separating from the first component.

Abstract: An interstage seal is provided for installation between forward and aft stage discs of a turbine section of a gas turbine engine. The seal has a forward annular portion, which extends forward to contribute to an air seal at the outer periphery of the forward stage disc, and an aft annular portion, which extends rearward to contribute to an air seal at the outer periphery of the aft stage disc. The seal has a peripheral slot into which static vane inner rails are loadable. The seal includes a first component, which provides the forward annular portion, and a second component, which provides the aft annular portion. The second component is separably connectable to the first component such that the first and second components, when connected, cooperate to form the peripheral slot. Loading the static vane inner rails into the slot prevents the second component separating from the first component.