Abstract: The present disclosure relates to an assembly for formation of a fan blade. The assembly comprises a suction panel; a pressure panel; and a membrane having a leading edge and a trailing edge. The membrane is sandwiched between the suction panel and pressure panel. The membrane comprises a gas entry slot extending in a radial direction, the gas entry slot having a radially outer receiving portion for receiving a pipe, and a radially inner portion. The radially inner portion of the gas entry slot has a substantially uniform width in a direction between the leading and trailing edge of the membrane.

Abstract: A turbine blade has a root portion, a platform and an aerofoil, the aerofoil is mounted on the platform and is formed by a pressure side wall and a suction side wall and has an outer surface, the pressure side wall and the suction side wall meet at a leading edge and a trailing edge, the aerofoil has an axial chord length, the suction side wall defines part of the radially outward surface of the aerofoil, the suction side wall defines an overhang, the overhang has a maximum overhang length that is between 5% and 20% of the axial chord length of the blade and is located between 5% and 50% of the suction surface length from the leading edge.

Abstract: A turbine arrangement for a gas turbine engine comprising a turbine shaft. An axial array of turbine rotors, having a first axial end and a second axial end. A drive arm coupled between the turbine shaft and the first axial end. A measurement system arranged to measure a parameter of the turbine arrangement, the measurement system positioned at the second axial end. The parameter may be rotational speed.

Abstract: A geared gas turbine engine for an aircraft includes: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.6 or lower, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 1 to around 1.3 at cruise conditions.

Abstract: An arrangement for a gas turbine engine that includes a turbine blade configured to rotate about an axis, a casing radially outside of the turbine blade, and a carrier segment mounted to the casing so as to define a first impingement space therebetween. The carrier segment is positioned radially outside the turbine blade and includes a first impingement carrier wall, a main carrier wall, and a cooling chamber. The first impingement carrier wall is adjacent to and radially inside of the first impingement space, and the first impingement carrier wall includes a first aperture. The main carrier wall is radially inside of the first impingement carrier wall. The cooling chamber is radially inside of the main carrier wall. Additionally, an intermediate chamber is disposed radially between the cooling chamber and the first impingement space. A second aperture is configured to allow ingress of air into the intermediate chamber.

Abstract: A hollow aerofoil is described having a leading edge and a trailing edge. The leading edge and trailing edge are connected by a pressure surface side (34) and a suction surface side (37) and one or more cavities are bounded by the pressure surface side (34) and/or suction surface side (37). In use, the cavity is arranged to receive coolant from a coolant source. The trailing edge has an apogee (36) where the pressure surface side (34) and suction surface side (37) meet. In an embodiment, a row of holes (32) is provided to a pressure surface side of a centreline of the apogee (36), the holes (32) being in fluid communication with cavity. The arrangement of the holes is such that outlets (33) to the holes extend from the apogee (36) and onto an adjacent part of the pressure surface side (34). A method for the manufacture of the aerofoil is also described.

Abstract: A Ni-based superalloy powder having a freezing range of not more than 80° C. is disclosed, with a composition, expressed in weight percent, of: Ni; 9-25 Co; 3-15 Cr; 0-5 Mo; 0-12 W; 3-9 Al; 0-6.5 Nb; 0.3-1 C; 0-6.5 Ta; 0-3 Ti, and incidental impurities. It is disclosed in methods for manufacturing or repair of gas turbine engine components such as gas turbine blades using additive layer manufacturing techniques.

Abstract: An electrical machine including a moveable permanent magnet configured to generate a first magnetic field. A stator including windings configured to excite a second magnetic field. A moveable inter-pole component located between the permanent magnet component and the stator, the inter-pole component comprising an array of magnetic inter-pole pieces. The speed of movement of the inter-pole component is controlled to set a magnetic gear ratio between the first and second magnetic fields. Also a gas turbine engine, propulsor or thruster incorporating the electrical machine.

Abstract: An apparatus for manufacturing an article from powder material includes a canister, a sorter, a plurality of hoppers and at least one valve. The canister has a predetermined internal shape to define the shape of the powder metal article. The sorter sorts the powder material by the size of the powder particles, the shape of the powder particles and/or the flow characteristics of the powder particles. The hoppers contain powder material with different sizes of powder particles, different shapes of powder particles and/or powder particles with different flow characteristics. The hoppers are arranged to supply the sorted powder material to the canister. The at least one valve controls the proportions of the different powder materials supplied from the one or more of the different hoppers into the canister to control the packing density of the powder material in the canister at all positions in the canister.

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: An assembly for hot forming a component includes first and second location features reversibly attached to a sheet metal element that forms the component. The assembly is loaded into a high temperature forming rig by loosely locating the location features into corresponding datum features in a forming tool. As the assembly heats up, the location features locate accurately with the datum features. Thus, the assembly can be accurately located onto the tool accurately and with minimal opening of the hot forming rig, thereby improving efficiency and safety.

Abstract: A rotary friction welding process is provided, including: providing a first axisymmetric workpiece having a first annular weld surface, and a second axisymmetric workpiece having a second annular weld surface; and rotationally welding the workpeices. On the longitudinal section through the aligned workpieces, each of the first and second annular weld surfaces is flanked by radially inner and outer side surfaces, the first weld surface and its side surfaces being shaped to thermally match the second weld surface and its side surfaces across a line of initial contact of the first and second weld surfaces such the heat flows from the weld at all the side surfaces are substantially equal.

Abstract: Apparatus for insertion into a cavity of an object, the apparatus comprising: a first tube comprising a first end, a second end, a first cavity extending between the first end and the second end, and an aperture at the second end; a member positioned within the first cavity of the first tube and adjacent the aperture of the first tube, the member being configured to enable an action to be performed; a plug coupled to the second end of the first tube; and an actuator configured to move the first tube relative to the object between a first position in which the plug seals the member from the cavity of the object, and a second position in which the member is exposed to the cavity of the object.

Abstract: A fan blade containment system is arranged to surround a fan including a plurality of fan blades in a gas turbine engine for an aircraft. The fan blade containment system includes a fan case arranged to surround the fan and two debris retainers. The first debris retainer extends inwardly from the fan case and is arranged to prevent forward debris release should all or part of a fan blade become detached from the fan, the first debris retainer being located near and forward of the fan. The second debris retainer extends inwardly from the fan case and is arranged to prevent forward debris release should all or part of a fan blade become detached from the fan, the second debris retainer being located forward of the first debris retainer.

Abstract: A gas turbine engine includes a fan with a plurality of aerofoils such as fan blades. Each fan blade includes a leading edge and a leading edge zone behind the leading edge. The leading edge zone extends spanwise for the full length of the fan blade. The leading edge zone includes one or more deflected regions which locally reduce the angle of attack of the fan blade. The deflected regions extend from the leading edge to between 1% and 50% of the chordal extent of the fan blade.

Abstract: An electro-discharge machining (EDM) device comprises a housing, an erosion electrode, a ground electrode, and a dielectric inlet. The housing is configured to be positioned on a surface of a workpiece and proximal to a fastener to be eroded. The planar erosion electrode is positioned at least partially within the housing, with the erosion electrode being movable relative to the housing in a first axis. The first axis is substantially normal to a longitudinal axis of the fastener. The ground electrode is conductively connected to the fastener. The dielectric inlet is configured to deliver a dielectric fluid to a region between the erosion electrode and the fastener.

Abstract: The present disclosure relates to optimisation of a cold nozzle, or bypass exit area, for a gas turbine engine, in particular for a geared turbofan gas turbine engine. Example embodiments include a method of optimising a geared turbofan gas turbine engine for an aircraft, the method comprising: determining expected service parameters for the aircraft, the expected service parameters including an expected range of travel for the aircraft; selecting components for the geared turbofan gas turbine engine to define a first smaller cold nozzle area if the range of travel is within a first smaller range and to define a second larger cold nozzle area if the range of travel is within a second larger range.

Abstract: A fan blade containment system is arranged to surround a fan including a plurality of fan blades in a gas turbine engine for an aircraft. The fan blade containment system includes a fan case arranged to surround the fan; a debris retainer, such as a front hook, mounted on the fan case and arranged to prevent forward debris release should all or part of a fan blade become detached from the fan; and a front panel mounted on the fan case adjacent to and rearward of the front hook, and arranged to move or break when struck by a detached fan blade or fan blade part so as to allow the detached fan blade or fan blade part to engage the front hook.

Abstract: A fan blade containment system is arranged to surround a fan including a plurality of fan blades in a gas turbine engine for an aircraft. The fan blade containment system includes a composite fan case arranged to surround the fan; and a metallic insert mounted on the composite fan case and including a metallic hook arranged to prevent forward debris release should all or part of a fan blade become detached from the fan.

Abstract: A method of filling an internal cavity of a component for a machine with viscoelastic damping medium and/or insulating medium comprises pumping a plurality of capsules into the internal cavity using a peristaltic pump. Each capsule comprises a flexible skin encapsulating a viscoelastic damping medium, one or more viscoelastic damping medium precursors, an insulating medium, or one or more insulating medium precursors.