Abstract: The present disclosure relates to a combustor assembly for a gas turbine engine. The combustor assembly comprises a combustor liner, a combustor head, a cowl and a fastener. The combustor liner and the cowl define a cavity. The combustor liner has an integral lug that extends into the cavity from a wall of the combustor liner. The fastener extends into the combustor liner lug to fasten the combustor head to the combustor liner.

Abstract: A combustor assembly for a gas turbine engine comprises a combustor casing. The combustor casing has a diametral height D and an axial length L. An Aspect Ratio parameter S is defined as: S = D L and the S parameter lies in the range of 0.30 to 1.00.

Abstract: A combustor assembly for a gas turbine engine comprises a combustor casing, and an integer quantity N of fuel swirl nozzles. The combustor casing defines a total internal volume of the combustor casing V (cm3) together with a Fuel Swirl Nozzle Density ratio (m?3) that is defined as: D F ? S ? N = N V and the DFSN parameter is a value in the range of 200 to 1,500.

Abstract: An integral guide vane combustor assembly for a gas turbine engine comprises an annular array of radially extending guide vane combustors. The gas turbine engine comprises, in axial flow sequence, a compressor assembly, the integral guide vane combustor assembly, a turbine assembly, and an exhaust assembly. Each of the guide vane combustors comprises, in axial flow sequence, a guide vane portion, and a combustor body portion. The guide vane portion is formed integrally with the combustor body portion, and the guide vane portion is configured to direct a gas flow exiting the compressor assembly into the combustor body portion.

Abstract: A component for a hot section of an engine comprises a substrate and an infrared-reflective layer configured to reflect infrared radiation away from the substrate and in some cases the infrared-reflective layer has: a spectral reflectance R_? of no less than about 0.5 when measured by spectrophotometry, according to ISO 15368:2021, using incident electromagnetic radiation having a wavelength ? no less than about 500 nm and no greater than about 1 mm; and a spectral emissivity ?_? of less than about 0.4, for example no greater than about 0.3, when measured using incident electromagnetic radiation having a wavelength ? no less than about 500 nm and no greater than about 1 mm.

Abstract: There is described a fuel injector for a gas turbine engine. The fuel injector comprises an air passageway having an inlet region and an outlet region in fluid communication with the inlet region at an air passageway interface, the inlet region being configured to receive a flow of air from a compressor of the gas turbine engine at an air inlet, the outlet region being configured to receive the flow of air from the inlet region via the air passageway interface and discharge the flow of air to a combustor head of the gas turbine engine. The fuel injector also comprises a fuel passageway having a fuel outlet configured to discharge a flow of fuel into the combustor head. A width of the inlet region in a direction perpendicular to a centreline of the air passageway decreases continuously from the air inlet to the air passageway interface.

Abstract: Disclosed is a method of manufacturing a component, the method comprising: determining a predicted direction and extent of distortion of a portion of the component; designing a support structure based on the predicted direction and extent of distortion of the portion of the component; additive manufacturing, by melting and fusing powder from a build plate, the component and the support structure. The support structure comprises, a hook structure comprising an elongate arm and a hook protruding from the elongate arm. The hook is deposited to abut a distortion surface of the component to restrain distortion of the component. The distortion surface is the furthest in the direction of distortion of the component. The elongate arm is deposited to extend from the hook in a direction having a vector component opposing the direction of distortion such that shrinkage of the elongate arm provides a force to minimise distortion of the component.

Abstract: There is disclosed a method of manufacturing a component. The method comprises melting and fusing powder in layers from a build plate to form the component and a disposable support structure, heat treating the component; and removing the support structure. Each layer of the support structure and the component are formed such that there is at least a 10-micron gap between them, to ensure that the component and the support structure are not fused together.

Abstract: Disclosed is an apparatus for the additive manufacture of a component comprising: a bed for supporting the component; a powder depositor; an energy beam source for emitting a beam of energy towards a beam target location; and control means for adjusting the beam target location; wherein the apparatus is configured to provide rotation of the beam target location and/or the powder depositor relative to the bed, the relative rotation being about a rotational axis. Also disclosed are an additive manufacture apparatus comprising a substantially annular bed, a method for additive manufacture, a gas turbine engine, and an aircraft.

Abstract: Disclosed is an apparatus for the additive manufacture of a component comprising a substantially annular bed for supporting the component, a powder depositor, an energy beam source for emitting a beam of energy towards a beam target location, and control means for adjusting the beam target location. The apparatus is configured to provide rotation of the beam target location and/or the powder depositor relative to the bed, the relative rotation being about a rotational axis. Also disclosed are a method for additive manufacture, a gas turbine engine, and an aircraft.

Abstract: A combustion chamber comprising a plurality of circumferentially arranged cassette segments coupled to a combustor head at one end and a wall section at the other end, each cassette segment extending the full length of the combustion chamber, and wherein the combustor head has an annular tongue structure on a mating surface, the tongue structure engages with a groove portion present in each of the cassettes so that when assembled the groove portions in the each of the plurality of cassette segments forms a substantially continuous groove.

Abstract: A combustion liner for a gas turbine engine. The combustion liner defines a bypass direction which extends between an upstream portion and a downstream portion of the combustion liner in use, the combustion liner. The combustion liner comprises a liner wall for defining at least a portion of a combustion chamber, the liner wall having an outer surface and an inner surface and a depth between the outer and inner surfaces of the combustion liner in use. The combustion liner comprises a chute formed through the liner wall for conveying fluid from the outer surface through the liner wall and ejecting fluid from an exhaust hole of the chute on the inner surface, the exhaust hole having a length L along the bypass direction.

Abstract: 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.

Abstract: A fuel injection system for a gas turbine engine comprises; a pilot fuel injector section and a main airblast fuel injector section, the main airblast fuel injector section having an aft end facing a combustion chamber. A surface of the injection system exposed to air flow through an injection system is non-axisymmetric or non-planar in a reference circumferential plane and/or is configured to generate controlled and varying acoustic impedance at or adjacent the aft end where, in use, the air flow collides with an oncoming acoustic wave.

Abstract: A fuel injector comprises an elongate fuel passage (31) having an elongate axis (31a) extending from an upstream inlet end to a downstream outlet end. A plurality of outlets (33) is arranged at the outlet end, each outlet extends obliquely with respect to the elongate axis (31a). The elongate fuel passage is defined by an inner skin of a double skinned pipe, the double skinned pipe defines a first annular cavity (34) between the inner skin and an outer skin. The inner skin and the outer skin meet adjacently upstream of the one or more outlets to close an end of the first annular cavity (34). The injector has a nose section (32) at a downstream end, the nose section (32) being convergent and fluted. The flutes (38) are arranged between the outlets (33) and extend towards the downstream end of the nose section (32) whereby to guide an air stream (A) passing over the injector (30) to form single jet at the downstream end of the nose section (32).

Abstract: A gas turbine engine combustion chamber comprises upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. The upstream end of each combustion chamber segment comprises a surface having a plurality of circumferentially spaced radially extending holes and the upstream ring structure having a plurality of circumferentially spaced holes extending radially through a portion abutting the surface of the upstream end of each combustion chamber segment. Each combustion chamber segment being removably secured to the upstream ring structure by a plurality of fasteners locatable in the holes in the combustion chamber segment and corresponding holes in the upstream ring structure.

Abstract: A combustion chamber assembly comprises a combustion chamber and a plurality of nozzle guide vanes. Each nozzle guide vane comprises an inner platform, an outer platform and an aerofoil. The combustion chamber comprises an annular wall which includes at least one box like structure. An outer wall of each box has a plurality of apertures for the supply of coolant into the box and the interior of the box is divided into at least two regions. The upstream end of each box has apertures to supply coolant from a first region of its interior onto an inner surface of the inner wall to form a film of coolant. The downstream end of each box has apertures to supply coolant from a second region of its interior onto a surface of the inner or outer platform of the nozzle guide vanes to form a film of coolant.

Abstract: A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. A frame structure at the downstream end of each segment has multiple spaced radially extending holes. The downstream end of each segment has an axially upstream extending groove and the downstream ring structure has an annular axially upstream extending hook which locates in the groove of each segment. A portion of the downstream ring structure abuts the frame structure of each segment. The downstream ring structure has multiple holes through the portion abutting the frame structure and segment is removably secured to the downstream ring structure by multiple fasteners locating in the holes in the segments and the downstream ring structure.

Abstract: A combustion chamber includes at least one annular wall which includes at least one box like structure and each box like structure includes an inner wall, outer wall, upstream end wall and downstream end wall. The inner wall is spaced radially from the outer wall and the outer wall has a plurality of apertures for the supply of coolant into the box like structure. The inner wall, the outer wall, the upstream end wall and the downstream end wall are integral. The upstream end of the annular wall has features to secure the annular wall to an upstream ring structure and a downstream end of the annular wall has features to mount the annular wall on a downstream ring structure. The inner wall has at least one slot extending through the full thickness of the inner wall to accommodate differential thermal expansion between the inner wall and the outer wall.

Abstract: A fuel injection system for a gas turbine engine (10) comprises; a pilot fuel injector section (22, 23) and a main airblast fuel injector section (25, 26, 27, 28, 29), the main airblast fuel injector section having an aft end (29) facing a combustion chamber (30). A surface of the injection system exposed to air flow through an injection system is non-axisymmetric or non-planar in a reference circumferential plane and/or is configured to generate controlled and varying acoustic impedance at or adjacent the aft end where, in use, the air flow collides with an oncoming acoustic wave.