Abstract: A fuel supply system for fuel injectors of a multi-stage combustor of a gas turbine engine is provided. The system has a metering and splitting arrangement which receives a pressurised fuel flow and controllably meters and splits the received flow into metered pilot and mains flows for injecting respectively at pilot and mains fuel discharge orifices of the injectors to perform staging control of the combustor. The system further has pilot and mains fuel distribution pipeworks respectively distributing fuel from the metering and splitting arrangement to the pilot and mains discharge orifices. The system further has a controller which is configured to command the metering and splitting arrangement to provide the metered flows, to select the pilot distribution pipework and deselect the mains distribution pipework for pilot-only operation.

Abstract: A gas turbine engine for an aircraft. The gas turbine comprises a combustor, a fuel injection system connected with a source of fuel and configured to inject fuel into the combustor, a water injection system connected with a source of water and which is configured to inject water into the combustor, and a control system. The control system is configured to identify an atmospheric condition; determine a water-fuel ratio for injection into the combustor of the gas turbine engine in response to the atmospheric condition; and control injection of fuel and water by the fuel injection system and the water injection system according to said water-fuel ratio to control an soot emissions caused by combustion of fuel therein.

Abstract: A gas turbine engine for an aircraft. The gas turbine comprises a staged combustion system having pilot injectors and main injectors, a fuel metering system configured to control fuel flow to the pilot injectors and the main injectors, and a fuel system controller. The controller is configured to identify an atmospheric condition, determine a ratio of pilot fuel flow rate for the pilot injectors to main fuel flow rate for the main injectors in response to the atmospheric condition, and inject fuel by the pilot injectors and the main injectors in accordance with said ratio to control an index of soot emissions caused by combustion of fuel therein.

Abstract: The blades for a rotor of a gas turbine engine are all manufactured to the same design. However, manufacturing tolerances mean that in practice each individual blade is different to the others. It is proposed to arrange the blades around the circumference of the rotor in a manner that limits excessive stress being induced in the blades due to differences in the vibration response between a given blade and its two neighbouring blades.

Abstract: An aerodrome system for an aerodrome is provided. The aerodrome system comprises a plurality of light emitting elements configured to emit light from a surface of the aerodrome upon which aircraft may take-off, land and manoeuvre; and a controller operatively coupled to each of the light emitting elements so as to selectively control the light emitting elements, wherein the light emitting elements form pixels of a display and the controller is configured to control an image displayed by the display so as to display and demarcate at least one runway for aircraft to take-off or land, wherein the light emitting elements are spaced such that the controller may display the runway with a variable orientation and the controller is further configured to change the image displayed by the display so as to change the orientation of the runway.

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of backswept blades (211,212) each of which have a blade exit angle (?2) of from ?50 to ?70 degrees, and a fully vaneless diffuser (203).

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of blades, and a fully vaneless diffuser (203). The radius of the inlet (r0) is from 25 to 50 percent of the radius of the impeller (r2).

Abstract: A gas turbine engine includes an engine core and a fan located upstream of the engine core. The engine core includes: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. The first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor. A fan diameter ratio of: first ? ? flange ? ? radius fan ? ? diameter is equal to or greater than 0.125.

Abstract: A method of upgrading a modular gas turbine engine, which includes: a first fan module with a fan having plurality of fan blades; a first engine core module including an engine core and a gearbox providing drive to the fan; and a first fan case module with a fan case arranged to enclose the fan blades, the method including: disassembling the gas turbine engine, replacing one of the first fan module, first engine core module or first fan case module with a replacement fan module, a replacement engine core module or a replacement fan case module; and reassembling the gas turbine engine using the replacement module, which is compatible with the others of the first fan module, first engine core module or first fan case module; and the replacement module designed to different parameters to one of the first fan module, first engine core module or first fan case module.

Abstract: The present disclosure relates to a method of manufacturing an aerofoil structure for a gas turbine engine, wherein the aerofoil structure includes a root configured to be received in a rotor disc of the gas turbine engine, the method including: providing a pre-moulded polymer insert; adding the polymer insert into a mould for forming the aerofoil structure; adding a composite constituent into the mould; and heating the composite constituent in the mould to bond the polymer insert to the composite constituent, the polymer insert being provided at a shoulder of the aerofoil structure root. The composite constituent is pre-impregnated with a resin and heating the composite constituent in the mould thermosets the resin. Heating the composite constituent in the mould to bond the polymer insert to the composite constituent forms an intermediate part and the method further comprises machining the intermediate part to remove excess material and form the aerofoil structure.

Abstract: An aerofoil assembly includes a platform and a plurality of aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. Each aerofoil has a leading edge proximal to the first edge and a trailing edge distal to the first edge. A pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface. A mid-pitch location is defined midway along the pitch spacing. The platform defines one or more recesses disposed between the leading edges of the plurality of aerofoils and the first edge. Each of the one or more recesses is disposed proximal to the mid-pitch location between adjacent aerofoils.

Abstract: An aerofoil component (10) defining an in use leading edge (16) and a trailing edge (18), the leading edge (16) comprising at least one serration (22) defining a sawtooth profile projecting in a generally upstream direction, each serration (22) comprising a plurality of rectangular-wave projections (26) extending in an upstream chordwise direction from the sawtooth profile.

Abstract: An engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox. The gearbox is an epicyclic gearbox and comprises a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20×109 N/m, and/or the tilt stiffness of the planet carrier is greater than or equal to 6.00×108 Nm/rad. A method of operation of such an engine is also disclosed.

Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting the turbine and compressor; a fan upstream of the engine core including a plurality of fan blades; and a gearbox that receives an input from a gearbox input shaft portion of the core shaft and outputs drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox including a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon, and wherein the sun gear receives input from the core shaft. At cruise conditions the torque on the core shaft is greater than 10,000 Nm and a ratio of core shaft stiffness to core shaft torque is within a specified range.

Abstract: A method of detecting an anomaly in a crystallographic structure, the method comprising: illuminating the structure with x-ray radiation in a known direction relative to the crystallographic orientation; positioning the structure such that its crystallographic orientation is known; detecting a pattern of the diffracted x-ray radiation transmitted through the structure; generating the simulated pattern based on the known direction relative to the crystallographic orientation; comparing the detected pattern to a simulated pattern for x-ray radiation illuminating in the known direction; and, detecting the anomaly in the crystallographic structure based on the comparison.

Abstract: A gas turbine engine for an aircraft has an engine core having a turbine, compressor, and core shaft connecting the turbine and compressor; a fan upstream the engine core, the fan having fan blades; and a gearbox. The gearbox receives an input from a core shaft and outputs drive to a fan to drive the fan at a lower rotational speed than the core shaft. The gearbox is an epicyclic gearbox and has a sun gear, planet gears, ring gear, and planet carrier on which the planet gears are mounted. The gearbox has a gear mesh stiffness between the planet gears and the ring gear and a gear mesh stiffness between the planet gears and the sun gear. The gear mesh stiffness between the planet gears and the ring gear divided by that between the planet gears and the sun gear is in the range from 0.90 to 1.28.

Abstract: An aerofoil assembly includes a platform and one or more aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. The one or more aerofoils are disposed between the first edge and the second edge. Each of the one or more aerofoils has a leading edge proximal to the first edge and a trailing edge distal to the first edge. The platform defines one or more recesses disposed between the leading edge of each of the one or more aerofoils and the first edge.

Abstract: The present disclosure relates to a variable stator vane and a method of fabricating the variable stator vane of a gas turbine engine. The method includes providing at least one fibre sheet. The method further includes rolling the at least one fibre sheet around a mandrel to form a spindle section of the variable stator vane. An excess of material of the at least one fibre sheet remains after forming the spindle section. The method further includes using the excess of material of the at least one fibre sheet to form the at least one aerofoil section of the variable stator vane.

Abstract: A nacelle for housing a fan within a gas turbine engine having a longitudinal centre line includes a leading edge and a trailing edge. A nacelle length (Lnac) is defined as an axial distance between the leading edge and the trailing edge. An azimuthal angle (?) is defined about the longitudinal centre line. The nacelle length (Lnac) varies azimuthally. The nacelle length (Lnac) decreases azimuthally from an inboard end of the nacelle to an outboard end of the nacelle.

Abstract: The present disclosure relates to a thermal barrier coating for coating a substrate. The thermal barrier coating may comprise an inner ceramic layer (e.g. 7YSZ) having a columnar grain structure and a first outer ceramic layer (e.g. 7YSZ) having a branched grain structure. The thermal barrier coating further comprises a nucleation layer (which may comprise alumina or tantala), interposed between the inner ceramic layer and the first outer layer. The layers can be deposited by PVD using substantially contact deposition parameters because the nucleation layer induces branching in the first outer ceramic layer.