Abstract: A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Abstract: Described is a nozzle guide vane assembly may comprise: a plurality of nozzle guide vane passages arranged in an annular array, each nozzle guide passage comprising inner and outer radial platforms, a suction surface of a first vane, and a pressure surface of a second vane, wherein the nozzle guide passages are made from a plurality of layers of ceramic fibers held in a ceramic matrix, wherein a first nozzle guide passage includes at least one attachment lug attaching the nozzle guide passage to either or both of an engine casing and an adjacent second nozzle guide passage, the attachment lug being integral with the first nozzle guide passage.

Abstract: A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Abstract: A method of constructing the nozzle guide passage may include the steps of: providing a tubular pre-form having a plurality of layers of unconsolidated CMC fibers in which the internal walls of the tubular preform provide either working surfaces which define the main gas path walls of the assembled nozzle vane, or extraneous wall portions which do not provide working surfaces of the assembled nozzle guide vane arrangement; cutting the wall portions to provide one or more flaps which remain attached to the pre-form via a connecting portion of wall; and, displacing the flap from a first pre-cut position, to a second consolidation position to provide one or more ancillary features.

Abstract: A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Impingement tubes direct cooling air onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine.

Abstract: A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Impingement tubes direct cooling air onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine.

Abstract: A gas turbine engine assembly adapted to separate a high pressure zone from a low pressure zone includes a seal assembly configured to block gasses from passing through the interface of two adjacent components. The seal assembly includes a rod.

Abstract: The invention concerns an article (20) formed of a ceramic matrix composite structure having a plurality of ceramic fiber layers (22) and a binder material (24) interspersed throughout said layers. The ceramic matrix composite material may be sintered. The ceramic fiber layers undulate relative to one or more outer surfaces (38;40) of the article. Thus support features (48) within the article are able to share a load in use between a plurality of layers. The invention may be suited to engine components such as turbine seal segments in a gas turbine engine.

Abstract: A gas turbine engine assembly adapted to separate a high pressure zone from a low pressure zone includes a seal assembly configured to block gasses from passing through the interface of two adjacent components. The seal assembly includes a rod.

Abstract: This disclosure describes a wall arrangement for a gas turbine engine, comprising: an annular wall comprising a plurality of circumferential wall segments, adjacent wall segments having opposing first and second end wall portions which define a separating gap therebetween, the gap including a saddle portion which faces radially outwards and comprises a first sealing face on the first end wall portion and a second sealing face on the second end wall portion; and, a longitudinal seal member having a curved sealing surface located within the saddle portion, wherein the curved sealing surface contacts the either or both first and second sealing faces along the length of the saddle portion in use, wherein the seal member includes axially compressible segments which allow the seal member to bend longitudinally.

Abstract: This document discloses a wall arrangement for a gas turbine engine, comprising: an annular wall comprising a plurality of circumferential wall segments, adjacent wall segments having opposing first and second end wall portions which define a separating gap therebetween, the gap including a saddle portion which faces radially outwards and comprises a first sealing face on the first end wall portion and a second sealing face on the second end wall portion; and, a longitudinal seal member having a curved sealing surface located within the saddle portion, wherein the curved sealing surface contacts either or both first and second sealing faces along the length of the saddle portion in use, wherein the seal member includes one or more axial restraint features which engage with a corresponding feature of the wall segment or a supporting structure thereof.

Abstract: Described is a wall arrangement for a gas turbine engine, comprising: an annular wall comprising a plurality of circumferential wall segments, adjacent wall segments having opposing first and second end wall portions which define a separating gap therebetween, the gap including a saddle portion which faces radially outwards and comprises a first sealing face on the first end wall portion and a second sealing face on the second end wall portion; and, a longitudinal seal member having a curved sealing surface located within the saddle portion, wherein the curved sealing surface contacts the either or both first and second sealing faces along the length of the saddle portion in use, in which the seal member further includes a plurality of flow passages extending from an outboard side to an inboard side.

Abstract: A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Impingement tubes direct cooling air onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine.

Abstract: A method of constructing the nozzle guide passage may include the steps of: providing a tubular pre-form having a plurality of layers of unconsolidated CMC fibres in which the internal walls of the tubular preform provide either working surfaces which define the main gas path walls of the assembled nozzle vane, or extraneous wall portions which do not provide working surfaces of the assembled nozzle guide vane arrangement; cutting the wall portions to provide one or more flaps which remain attached to the pre-form via a connecting portion of wall; and, displacing the flap from a first pre-cut position, to a second consolidation position to provide one or more ancillary features.

Abstract: A seal segment is positioned, in use, radially adjacent the rotor. The seal segment has first and second circumferentially spaced passageways each of which extends in the fore and aft direction. In use, a first support bar is contained within the first passageway, and a second support bar is contained within the second passageway. The first and second support bars being mountable to complementary formations provided by the casing of the engine. The first passageway is configured such that the seal segment is fixed relative to the first support bar in the radial and circumferential directions. The second passageway is configured such that the seal segment is fixed relative to the second support bar in the radial direction but allows relative movement of the seal segment and the second support bar in the circumferential direction to accommodate differential thermal expansion of the seal segment and the casing.

Abstract: Described is a nozzle guide vane assembly may comprise: a plurality of nozzle guide vane passages arranged in an annular array, each nozzle guide passage comprising inner and outer radial platforms, a suction surface of a first vane, and a pressure surface of a second vane, wherein the nozzle guide passages are made from a plurality of layers of ceramic fibres held in a ceramic matrix, wherein a first nozzle guide passage includes at least one attachment lug attaching the nozzle guide passage to either or both of an engine casing and an adjacent second nozzle guide passage, the attachment lug being integral with the first nozzle guide passage.

Abstract: A seal segment is provided for a shroud ring of a rotor of a gas turbine engine. The seal segment is positioned, in use, radially adjacent the rotor. The seal segment includes a body portion having front and rear sides which each contain a respective slot that extends in the circumferential direction so that a respective hook formation is insertable into each slot for attaching the seal segment to the engine casing. The body portion is formed from continuous fiber reinforced ceramic matrix composite, reinforcing fibers of the composite wrapping around each slot so that, at the base of the slot, the reinforcing fibers are parallel to the surface of the body portion.

Abstract: A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Abstract: A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Abstract: A component of a gas turbine engine is formed from a continuous fiber reinforced ceramic matrix composite (CMC). The component has a sealing portion which, in use, makes sealing contact with an adjacent component of the engine. The sealing portion comprises a recess formed in the CMC and filled with a finer grade ceramic relative to the surrounding CMC, a metal or an intermetallic.