Abstract: A rigid electrical panel has both a flexible printed circuit and an electromagnetic protection layer embedded therein. The flexible printed circuit has an integral electromagnetic shield. The integral electromagnetic shield and the electromagnetic protection layer are electrically isolated, so as to provide independent grounding paths. The independent grounding paths are individually tested for safe operation.

Abstract: A combustion chamber wall is hollow, has a first surface and a second surface and includes a plurality of polyhedron shaped chambers defined by a matrix of integral interconnected walls. The walls of the chambers in a first layer define the first surface of the combustion chamber wall and the walls of the chambers in a third layer define the second surface of the combustion chamber wall. The chambers are fluidly interconnected by apertures extending through the integral interconnected walls of the chambers for the flow of coolant there-between. The walls of the chambers in the first layer have apertures extending there-through to supply coolant into the first layer and the walls of the chambers in the third layer have apertures extending there-through to supply coolant from the third layer into the combustion chamber. The combustion chamber wall is manufactured by additive layer manufacture.

Abstract: A vane for an exhaust system duct in a gas turbine engine including a vane plate, a guide plate, and a baffle. The vane plate includes a leading edge, a first leg and a second leg, with the first and second legs respectively extending on opposing sides of the leading edge to form a substantially U-shaped profile. The baffle connects respective distal ends of the first and second legs. The guide plate is accommodated within the vane plate, and includes a first aperture extending along the guide plate and aligned with the leading edge. The vane further includes a fluid inlet arranged, in use, to direct a fluid flow from outside the duct into an interior of the vane.

Abstract: A gas turbine engine and air intake assembly comprises an intake passage extending between an inlet highlight at a first end and an upstream face of a fan at a second end and comprising in flow series an intake lip, a most upstream portion of which defines the intake highlight and a most downstream portion of which defines a throat, a diffuser with sectional area broadening towards the fan; and a straight conditioning duct, arranged immediately upstream of the fan. The camber line of the intake passage intersects the engine main axis at an intersecting point upstream of the fan at an intersecting point and the camber line is parallel to the engine main axis in the straight conditioning duct.

Abstract: A bladed disk may be arranged to rotate about an axis of rotation and including a rotor disk having an axial length extending from a first end to a second end, along the axis of rotation and a radial thickness extending between a first radius from the axis of rotation and a second radius from the axis of rotation, less than the first radius. The bladed disk may include a plurality of blades formed integrally with the rotor disk and arranged circumferentially around the rotor disk, and a first set of divides and a second set of divides, each divide comprising a partial break in the rotor disk extending axially from the first end of the rotor disk to the second end of the rotor disk and radially from the first radius to a third radius, less than the first radius and greater than the second radius.

Abstract: A cobalt-free alloy is made essentially of: an effective concentration of up to 2.5 percent carbon; 12 to 58 percent chromium; 7 to 9 percent manganese; 4 to 5 percent silicon; 4 to 6 percent nickel; 0.08 to 0.2 percent nitrogen; the balance being iron plus incidental impurities.

Abstract: A gas turbine engine has a nacelle providing an external skin of the engine, a casing structure radially inwards of the external skin and providing an outer surface of an air flow duct of the engine, and plural engine components located in the bay formed between the casing structure and the nacelle. The components are part of either a first set of components or a second set of components. The components of the first set are all the components from the bay which are potential ignition sources, and the components of the second set are all the remaining components from the bay. The engine further has a container located in the bay, the container being fireproof capable of withstanding the application of heat by a standard flame for 15 minutes and the container containing the components of the first set.

Abstract: A blade or vane for a gas turbine engine comprising a pressure surface and a suction surface. The pressure surface or the suction surface comprises a roughness zone which is configured to provide greater flow resistance in a direction along the blade or vane than in a direction across the blade or vane.

Abstract: A method of forming a blade or vane for a gas turbine engine comprising: attaching a first outer layer to a first two-dimensional array of attachment areas on a first surface of an intermediate layer; attaching a second outer layer to a second two-dimensional array of attachment areas on a second surface of the intermediate layer opposite the first surface; and increasing a separation between at least a portion of the first and second outer layer to thereby deform the intermediate layer into a corrugated structure having corrugations in first and second directions.

Abstract: A containment arrangement for a gas turbine engine comprises: a radially outer annular casing (12); a fan track liner (28) radially within the casing comprising: an impact region (30) comprising a cellular material having a first compressive strength; at least one elongate ridge portion (26) extending through the impact region in a direction having an axial component, the ridge portion having a second compressive strength higher than the first compressive strength; and a gas-washed layer (38) radially within the fan track liner.

Abstract: Methods and apparatus are described for improving fit checking at the bonding interface of a first component, such as a fan blade (1), and a second component, such as an edge guard (2) for the fan blade (1), to be bonded by adhesive. A bondline thickness profile over the area of the bonding interface is determined by interposing a compressible, flexible sensor layer between the components before they are bonded. The sensor layer contains an array of piezoelectric elements which indicate local bondline thickness by signalling pressure due to compression of the layer. The bondline thickness profile can be processed by a programmed control processor to produce an adhesive application schedule prescribing the shapes of pieces of adhesive film which, when applied over the bonding interface, will build up an adhesive layer corresponding in thickness to the bondline thickness. The sensor layer can be prepared as a kit of shaped panels or as a contoured preform (55) matching the form of the bonding interface.

Abstract: A waveguide for a sensor system for mounting in a casing structure to measure the clearance between the casing structure and a rotary blade. The casing structure, which may be a turbine casing structure, includes a radially outer casing portion and a radially inner casing portion (e.g. a segment seal). The waveguide includes a radially outer portion for mounting in the radially outer casing portion and a radially inner portion for mounting in the radially inner casing portion. The two portions are sealed together and bridged by a sealing portion having a deformable portion.

Abstract: An aerofoil blade or vane for the turbine of a gas turbine engine includes: an aerofoil leading edge; an aerofoil trailing edge; an aerofoil suction side; and at least one reverse-pass coolant passage, extending within the aerofoil blade or vane; wherein the at least one reverse-pass coolant passage includes a substantial reverse-pass portion in which, in use, coolant flows in a direction away from the trailing edge and towards the leading edge.

Abstract: An aircraft (10) comprising a fuselage (12), the fuselage (10) comprising at least first and second inwardly tapering portions (20, 22) extending in a generally rearward direction from an aft end of the fuselage (12), each inwardly tapering portion (20, 22) comprising a boundary layer ingesting propulsor (36) mounted thereon.

Abstract: A turbomachine fan blade preform (10) comprising a pair of mutually opposing panels (12), and a membrane (20) sandwiched between the pair of panels (12). The membrane (20) is of a non-uniform thickness. The thickness of at least a portion of the membrane tapers along its length and/or width, between a relative thicker first region and relatively thinner second region.

Abstract: A mount assembly for attaching a heat exchanger to a casing of a gas turbine engine comprises a first attachment feature by which in use the mount assembly is attached to the casing and a second attachment feature spaced from the first attachment feature by which in use the mount assembly is attached to the heat exchanger. The first and second attachment features are joined by an elongate member. The assembly is characterised in that the elongate member is significantly larger in a length direction and a height direction than in a thickness direction, so that it is relatively stiff in the length and height directions L, H and relatively flexible in the thickness direction T; the flexibility allowing movement in use within the mount assembly to accommodate differential thermal expansion between the heat exchanger and the casing.

Abstract: A permanent magnet electrical machine has a rotor supporting a circumferential row of permanent magnets and a stator coaxial with the rotor and having a circumferential row of teeth carrying respective coil windings. The teeth provide paths for magnetic flux from the magnets, thereby electromagnetically linking the magnets and coils when the rotor rotates. The teeth have respective core portions on which the coil windings are mounted, and respective tip portions located between the core portions and the rotor. The tip portions are controllably rotatable between a first position where the tip portion and the core portion are angularly aligned to enhance magnetic flux linkage between the magnets and the coils, and a second position in which the tip portion is rotated out of angular alignment with its core portion to reduce magnetic flux linkage between the magnets and the coils. The tip portions are biased to the second position.

Abstract: The present disclosure concerns an electrical interconnect system for a vehicle such as an aircraft or marine/submarine vessel.

Abstract: A combustion chamber comprises an upstream end wall structure and inner and outer annular wall structures. The upstream end wall structure comprises an upstream wall and a plurality of circumferentially arranged heat shields secured to the upstream wall. The upstream wall has a plurality of circumferentially spaced fuel injector apertures. Each heat shield has radially outer and radially inner ends and a fuel injector aperture aligned with a corresponding fuel injector aperture in the upstream wall. The radially outer and inner ends of each heat shield have outer and inner rails spacing the heat shield from the upstream wall. The radially outer and inner ends of each heat shield have first and second pluralities of circumferentially spaced apertures extending there-through and through the associated outer and inner rails to direct coolant over the surface of the outer and inner annular wall structures to form respective films of coolant.

Abstract: There is a described an aerofoil component for a turbomachine, the aerofoil component comprising: a central core formed from a metal matrix composite; and an external layer comprising a pressure surface, a suction surface, a leading edge, a trailing edge and a root, the external layer being formed by a metal which covers the metal matrix composite of the central core. Also described is a method of manufacturing such an aerofoil component.