Abstract: A gas turbine engine for an aircraft, includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft, wherein the fan assembly has fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub having slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the fan blades is within the range of around 0.45 to around 0.7.

Abstract: A method of controlling at least part of a start-up or re-light process of a gas turbine engine, the method comprising: determining when a flame in a combustion chamber of a gas turbine engine is extinguished, during a start-up process or re-light process or during operation; purging the combustion chamber by controlling rotation of a low pressure compressor using a first electrical machine, and controlling rotation of a high pressure compressor using a second electrical machine, the combustion chamber downstream of the low pressure compressor and high pressure compressor; and controlling rotation of the low pressure compressor using the first electrical machine, and controlling rotation of the high pressure compressor using the second electrical machine to restart the start-up process or perform the re-light process.

Abstract: A method of controlling at least part of a start-up or re-light process of a gas turbine engine, the method comprising: increasing an angular velocity of a low pressure compressor; determining if an exit pressure of the low pressure compressor is equal to or greater than a first threshold pressure; in response to determining that the exit pressure of the low pressure compressor is equal to or greater than the first threshold pressure, controlling rotation of the low pressure compressor using a first electrical machine and controlling rotation of a high pressure compressor using a second electrical machine, to increase angular velocity of the high pressure compressor; determining if an exit pressure of the high pressure compressor is equal to or greater than a second threshold pressure.

Abstract: A method of controlling at least part of a start-up or re-light process of a gas turbine engine, the method comprising: controlling ignition within a combustion chamber of the gas turbine engine; controlling rotation of a low pressure compressor using a first electrical machine, and controlling rotation of a high pressure compressor using a second electrical machine, the combustion chamber downstream of the low pressure compressor and high pressure compressor; determining if an exit pressure of the high pressure compressor is equal to or greater than a self-sustaining threshold pressure; and in response to determining that the exit pressure of the high pressure compressor is equal to or greater than the self-sustaining threshold pressure, ceasing controlling rotation of the low pressure compressor using the first electrical machine, and/or the high pressure compressor using a second electrical machine.

Abstract: A geared turbofan engine include a gas turbine engine for an aircraft, including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan assembly to drive the fan assembly at a lower rotational speed than the core shaft, the hub including a plurality of slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, the rim having a minimum radial thickness between a base of each slot and an internal cavity within the hub, wherein the minimum radial thickness is within a range of around 0.5% to around 1.1% of the outer fan diameter.

Abstract: The present disclosure concerns apparatus and methods for reinforcing composite materials by insertion of reinforcement pins. Example embodiments include an apparatus for reinforcing a composite preform, comprising: a carriage for introducing reinforcement pins into the composite preform, the carriage comprising a drill module, a pin feeder module and a tamper module, the assembly being translatable relative to the composite preform, wherein the drill module and pin feeder module are exchangeable between a common actuation position on the carriage.

Abstract: A nose cone assembly is provided comprising a nose cone, a support ring on which the nose cone is mounted, and an insulating layer disposed between the nose cone and the support ring, the insulating layer at least partially comprising glass fibres. A composite nose cone is also provided comprising an attachment region A, in which composite fibres are oriented substantially perpendicular to an axis of rotation of the nose cone, and an impact region I, in which composite fibres are oriented substantially parallel to the axis of rotation of the nose cone. Also provided is a support ring for a nose cone assembly, the support ring comprising an annular main body and an annular projecting attachment flange, the attachment flange terminating in a plurality of circumferentially spaced attachment tabs, wherein the attachment flange forms with the main body a cross section having at least two C curves.

Abstract: A turbine case cooling system comprises a manifold (21) radially adjacent a portion of a radially outer surface of the turbine case (26) and in fluid communication with one or more of radially inwardly directed outlets (25). The manifold (21) has a first inlet (22) and a second inlet (23). The first inlet (22) is obstructed by a first flow restrictor (22a) and the second inlet (23) is obstructed by a second flow restrictor (23a). The first inlet (22) includes a valve (24) upstream of the first flow restrictor (22a) and the valve is adjustable to control flow of fluid supply entering the first inlet (22).

Abstract: An aerofoil structure (26) for a gas turbine engine (10), the aerofoil structure comprising a carbon composite aerofoil portion (38) and a tip portion (48), wherein the tip portion comprises a tip surface (60) configured to face a corresponding casing structure (30), the tip portion further comprising a ridge line (62) extending along at least a portion of the tip surface, wherein the tip surface comprises a first surface (64) and a second surface (66) provided either side of the ridge line, the ridge line being defined by the intersection of the first and second surfaces, wherein the ridge line is configured to cut into the casing structure during an interaction between the aerofoil structure and casing structure.

Abstract: An aircraft fuel pumping arrangement comprises a first fuel pump, an electrically powered motor operable to drive the first fuel pump to deliver fuel to an outlet, a second fuel pump, a gas driven turbine operable to drive the second fuel pump to deliver fuel to the outlet, and a controller operable to control the operation of the gas driven turbine to determine the rate at which the second fuel pump is driven.

Abstract: An aerofoil and an aerofoil assembly, in particular an aerofoil with improved stagnation zone cooling and an aerofoil assembly comprising such an aerofoil. The aerofoil is an aerofoil for a gas turbine engine comprising a pressure surface, a suction surface, a leading edge, a trailing edge, a stagnation zone located in the region of the leading edge, and an elongate channel running along the leading edge at the stagnation zone.

Abstract: A vane cooling system for a gas turbine engine comprises a vane (21) arranged on a stator and having a chamber (23) extending continuously from a radially inner end to a radially outer end of the vane. The vane (21) has; a radially inner inlet (24) and a radially outer inlet (25), a first cooling fluid feed (39) in communication with the radially inner inlet (24) and a second cooling fluid feed (28) in communication with the radially outer inlet (25), The first cooling fluid feed (39) has a higher pressure than the second cooling feed (28). A flow adjustment device (30) is arranged for adjusting a flow of the second cooling fluid feed into the radially outer inlet (25).

Abstract: A nozzle for a gas turbine engine. An array of convergent petals hingedly coupled to a nozzle exit of fixed diameter. An array of divergent petals, each divergent petal hingedly coupled to one of the array of convergent petals. A cam surface associated with the array of convergent petals. A pivot point coupled to the array of divergent petals by a first linkage and to a fixed point by a second linkage. A cam follower coupled to the second linkage by a third linkage, the cam follower arranged to abut and travel in contact with the cam surface. An actuator coupled to the cam follower and arranged to translate the cam follower along the cam surface to move the convergent and divergent petals.

Abstract: An aerofoil body for a gas turbine engine is provided. The aerofoil body has leading and trailing edge portions, wherein one of the leading and trailing edge portions is a morphable edge portion having a composite layer structure. The aerofoil body further has a non-morphing central portion which forms pressure and suction surfaces of the aerofoil body between the leading and trailing edge portions. The composite layer structure includes a return spring, one or more shape memory alloy layers, and a flexible cover for the return spring and the one or more shape memory alloy layers. The flexible cover defines pressure and suction surfaces of the aerofoil body at the morphable edge portion. The one or more shape memory alloy layers are electrically heatable to deform the layers against the resistance of the return spring, and thereby alter the pitch of the aerofoil body at the morphable edge portion.

Abstract: An actuation system 100 comprises a fluid supply 130 that is fluidly connected to a position control actuator 110. The output of the position control actuator is determined by the fluid supply. The fluid supply is controlled by a piezoelectric actuator 150.

Abstract: Producing a component having an in use gas washed surface includes: obtaining a reference component having a reference shape with an in use gas washed surface; setting one or more performance threshold for the reference shape, the threshold defining an acceptable performance for the reference shape; obtaining a manufactured component made to the reference shape; measuring the manufactured component and determining a displacement distribution indicative of the geometric deviation of the manufactured component from the reference shape; determining a performance sensitivity distribution for the reference component, the sensitivity distribution having a plurality of points, each point indicative of a performance factor for the reference component; combining the sensitivity distribution and displacement distribution to determine a performance prediction for the manufactured component; determining whether the performance prediction is within the performance threshold; accepting or rejecting the component for use if

Abstract: A fan nacelle for an aircraft gas turbine engine. The nacelle includes an aft nacelle portion including a radially outer surface and a radially inner surface, the radially outer and inner surfaces defining an internal cavity therebetween. The nacelle further includes an aft nacelle segment translatable along a translation vector having an axial component, wherein the aft nacelle segment is configured to translate between a forward deployed position in which the nacelle defines a first primary fan nozzle exit area (A1) and a clean position in which the nacelle defines a second primary fan nozzle exit area (A2) less than the first primary fan nozzle exit area A1, wherein in the forward deployed position, the aft nacelle segment is at least partly located within the internal cavity.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A hub-tip ratio of the fan, defined as the ratio of the diameter of the hub to the diameter of the fan measured at the leading edge of the blades, is from 0.45 to 0.55.

Abstract: Method of reducing the thickness of a bore of a cylindrical workpiece for use as a gear. The method involves the steps of: mounting a cylindrical workpiece having a horizontal central axis and an outer diameter in a grinding machine; and grinding the bore of the cylindrical workpiece to reduce its thickness using a grinding wheel that has a diameter that is from 40% to 80% of the outer diameter of the cylindrical bore and has a direction of rotation about an axis of rotation that is parallel to the horizontal central axis of the cylindrical workpiece. The axis of rotation of the grinding wheel may be located from 90 degrees to 180 degrees, in the direction of rotation of the grinding wheel, from a plane that extends vertically through the workpiece when it is mounted in the grinding machine. The gear may be one of a planetary, sun, parallel axis or helical gear.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A plurality of outlet guide vanes are positioned aft of the fan. An afterbody is located aft of the plurality of outlet guide vanes and which tapers to an apex having an apex angle with respect to the rotational axis of between 35 and 45 degrees.