Abstract: The present disclosure concerns a heat exchanger, which may for example be utilised in a gas turbine engine or in other applications. Example embodiments include a heat exchanger comprising: an external surface for exchanging heat with an external fluid flow passing over the external surface; a first fluid passage extending through the heat exchanger from a first fluid inlet to a first fluid outlet, a first portion of the first fluid passage extending along the heat exchanger adjacent to the external surface for a first cooling fluid passing through the first fluid passage to exchange heat with the external fluid flow; and a second fluid passage extending through the heat exchanger from a second fluid inlet to a second fluid outlet located at the external surface for a second cooling fluid to pass from the second fluid inlet into the external fluid flow.

Abstract: A bladed disc for a turbine engine has disc and blade portions. The disc portion extends in a radial direction from the turbine engine central axis and has slots around its circumference with an inverse fir tree profile. The blade portion has aerofoil and root sections. The root section is configured to have a fir tree profile. The blade portion engages with a slot on the disc portion circumference with the fir tree profile of the root section of the blade portion engaging with the inverse fir tree profile of the slot within the disc portion. The fir tree profile of the root section of the blade portion and the inverse fir tree profile of the slot of the disc portion are curved. The bladed disc has a clearance plate mounted on at least one of the blades, the clearance plate extending across the slot on the disc portion.

Abstract: An aircraft including an aft-mounted boundary layer energisation system is shown. The system comprises a nacelle arranged around a tailcone of the aircraft which thereby defines a duct, the duct having, in axial flow series, an intake, a heat exchanger, and a nozzle, and no turbomachinery therein, whereby the system energises a boundary layer of the aircraft by means of Meredith effect.

Abstract: A gas turbine engine for an aircraft. The gas turbine engine comprises a rotatable shaft defining a rotational axis extending between rearward and forward ends of the gas turbine engine and a compressor drum surrounding, and coupled to, the shaft so as to define an annular gap therebetween. The gas turbine engine further comprises an intermediate case arranged axially rearward of the compressor drum and comprising a bearing support element extending into the annular gap, and a forward bearing mounted between the bearing support element and the shaft proximate a forward end of the compressor drum.

Abstract: The present disclosure relates to an electromechanical system comprising a toothed rotor and a stator segment. The stator segment comprises an armature coil defining a coil interior. The stator segment further comprises first and second stator portions, each stator portion comprising connected inner and outer radially extending poles. Each of the inner poles passes through the coil interior and each of the outer poles is provided outside of the coil interior. The stator segment further comprises a bridge spacing the stator portions and the bridge comprises a magnetic field generator arranged to generate a magnetic field between the stator portions.

Abstract: A vehicle propulsion system comprises a first propulsor and a second propulsor. The first and second propulsors are coupled to an electric machine comprising a stator and first and second rotors. Each of the first and second rotors are electromagnetically coupled to the stator. The first rotor is coupled to the first propulsor and the second rotor is coupled to the second propulsor.

Abstract: An electromechanical system comprising: a rotor comprising a plurality of teeth spaced about a circumference of the rotor, wherein the teeth of the rotor are equally spaced about the rotor according to a tooth spacing angle, and are skewed in an axial direction such that the circumferential positions of the teeth of the rotor vary along the axial length of the rotor, and first and second stator segments, each extending partway about the rotor and comprising a field winding and poles arranged to magnetically interact with the teeth of the rotor such that an alternating current (AC) back-emf is induced in the field winding upon rotation of the rotor. The poles of the first stator segment are angularly displaced about the rotor from the poles of the second stator segment such that the back-emf induced in the field winding of the first stator segment is phase shifted with respect to the back-emf induced in the field winding of the second stator segment.

Abstract: There is disclosed wall cooling system 50 having a double-wall geometry. A first wall 55 and a second wall 60 extend over a plan area with the second wall spaced from the first wall by a gap. The first wall 55 has multiple upstanding members 65 spanning the gap and contacting the second wall 60 such that the first and second walls are mechanically and thermally connected. The first wall 55 is shaped so as to provide a two-dimensional array of crests 85 and recesses 90. The crests 85 are spaced from the second wall 60. The first wall 55 has a plurality of through-holes 70 for flow of coolant through the first wall and into the gap. The cooling system 50 is suitable for use in a gas turbine engine 10, for example in the turbine 17, 19.

Abstract: The dual rotor electric machine comprises: a stator having one or more slots and one or more stator windings. The dual rotor electric machine further comprises a first rotor arranged to rotate relative to the stator with an airgap therebetween. The first rotor comprises a first rotor excitation element having one or more magnetic pole pairs arranged to interact with the stator windings, the first rotor configured to rotate about an axis (X). The dual rotor electric machine further comprises a second rotor arranged to rotate relative to the stator with an airgap therebetween. The second rotor comprises a second rotor excitation element having one or more magnetic pole pairs arranged to interact with the stator windings, the second rotor being configured to rotate about the axis (X).

Abstract: A combustion chamber assembly comprises an annular combustion chamber casing, a combustion chamber, a fuel injector and a tubular seal. The combustion chamber comprises an upstream end wall having an aperture extending there-through. A fuel injector head of the fuel injector is locatable in the aperture in the end wall and the fuel injector head has an axis. The tubular seal is positioned between the fuel injector head and the aperture in the end wall and the tubular seal has a flange and an aperture extends there-through. The tubular seal is movable radially and circumferentially with respect to the axis of the casing. The fuel injector head is locatable in the tubular seal. The aperture in the upstream end wall has a major radial dimension and a minor circumferential dimension relative to the axis of the casing to allow insertion or removal of a lean burn fuel injector head.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a support and a turbine vane arranged around the support. The support is made of metallic materials. The turbine vane is made of ceramic matrix composite materials to insulate the metallic materials of the support.

Abstract: A combustion chamber comprises an upstream end wall, at least one fuel injector and at least one seal. Each fuel injector is arranged in a corresponding aperture in the wall. Each seal is arranged in a one of the apertures in the wall and around one of the fuel injectors. Each seal has a first portion, a second portion and a third portion. The second portion abuts the corresponding fuel injector. The third portion is arranged at the downstream end of the seal and increases in diameter in a downstream direction. The first portion is arranged upstream of the second portion and has a plurality of coolant apertures extending there-through. The coolant apertures extend through the first portion with axial, radial and tangential components. The seals produce better mixing between the coolant and the fuel and air mixture from the fuel injectors to reduce emissions produced in the combustion chamber.

Abstract: A thermal management device for dissipating heat from one or more electrical components in an aircraft, the thermal management device including: a heat spreader with a plurality of layers of annealed pyrolytic graphite, the heat spreader having higher thermal conductivity in a plane defined by the layers than in a direction perpendicular to the layers; a flow path for a cooling fluid defined through the heat spreader, such that at least a portion of the flow path is surrounded by the heat spreader, wherein the flow path is helical, around a central axis extending through the plurality of layers; and a housing encasing the heat spreader, the housing arranged to have one or more electrical components mounted there-on and to form a thermal interface between the heat spreader and one or more electrical components.

Abstract: A boundary layer ingestion engine comprising a fan, the fan comprising a disc and a plurality of blades integrally formed with the disc.

Abstract: A method of spray coating a substrate is disclosed, the method comprising: a step of spray coating metal particles onto a substrate; and a step of induction heating the coating; wherein the step of induction heating comprises performing the induction heating in a vacuum.

Abstract: A method of manufacturing an abrasive media article, the method comprising the steps of: preparing an abrasive media composition comprising a binder material and ceramic based abrasive particles, wherein the binder material comprises an epoxy resin, wherein the particles form 50-90% by mass of the composition, and wherein the particles have an average size in the range of 5 ?m to 15 ?m; casting the composition into a mold; curing the composition in the mold; and removing the cured composition from the mold.

Abstract: A method of manufacturing a rotary abrasive machining tool, the rotary abrasive machining tool including a hub and a plurality of abrasive segments mounted to the hub, the method including the steps of: mounting each abrasive segment on the hub; machining an abrading edge on each abrasive segment while the abrasive segment is mounted on the hub.

Abstract: A combustion chamber comprising an upstream end wall, at least one annular wall, at least one fuel injector and at least one seal. The at least one annular wall being secured to the upstream end wall. The upstream end wall having at least one aperture. Each fuel injector being arranged in a corresponding one of the apertures in the upstream end wall and each seal being arranged in a corresponding one of the apertures in the upstream end wall and around the corresponding one of the fuel injectors. Each seal having an inner surface facing the corresponding one of the fuel injectors and an outer surface facing away from the corresponding one of the fuel injectors. Each seal abutting the corresponding one of the fuel injectors. The downstream end of each seal increasing in diameter in a downstream direction and the upstream end of each seal having a radially extending flange.

Abstract: An aerofoil member that extends between radially inner and radially outer endwalls which define a gas annulus of the compressor. Aerofoil member has a leading edge, a trailing edge, a pressure surface and a suction surface such that successive cross-sections through the aerofoil member transverse to the radial direction provide respective aerofoil sections. The external shape of the aerofoil member is defined by the stacking of the aerofoil sections on a stacking axis which passes through a reference point common to each aerofoil section. The aerofoil member has lean produced by the projection of the stacking axis onto a plane normal to the engine axis intersecting a first one of the endwalls at an angle of from 5° to 25° to the circumferential direction direction such that the pressure surface faces the first endwall.

Abstract: An annular array of turning vanes 200 is provided in a duct 100 of a gas turbine engine 10. The annular array of turning vanes 200 comprises aerodynamic vanes 220 and strut-vanes 240. The strut-vanes 240 have greater chord length and extend further axially downstream than the aerodynamic vanes 220. The leading edge of the strut-vanes 240 is upstream of the trailing edge of the aerodynamic vanes 220. The strut-vanes provide flow turning. The arrangement allows the duct 100 to be axially short.