Abstract: A method provides for measuring an assembled component including two or more parts, the method including the steps of: providing the component with a plurality of unique markers for detection during a scanning process; defining a first of the unique markers and aligning the first defined unique marker with a first element of the component; defining a second of the unique markers and aligning the second defined unique marker with a second element of the component; scanning the component and unique markers to capture a digitised version of the component and the relative positions of the first and second defined unique markers; and aligning a nominal CAD model of a part of the component with the equivalent part of the digitised version of the component using the relative positions of the first and second defined unique markers.

Abstract: There is disclosed a cutting mechanism for severing elongate fibre reinforcement material in composite material lay-up equipment, the cutting mechanism comprising a cutting element and a counteracting element which cooperate to sever fibre reinforcement material extending through the nip between them, the cutting element being mounted on an elongate arm which is pivotable about a pivot axis A spaced from the nip, to displace the cutting element relatively to the counteracting element to perform a cutting stroke, the mechanism further comprising guide means arranged to guide the fibre reinforcement material through the nip in a feed direction transverse to the pivot axis A, and an actuation device for driving the elongate arm in a cutting stroke, the actuation device acting on the elongate arm at a position away from the cutting element. There is also provided composite material lay-up equipment comprising the cutting mechanism.

Abstract: Engine has core compressor, combustor and turbine, fan located upstream of core and supersonic intake for slowing down incoming air at inlet formed by intake, bypass duct surrounding engine core, fan generates airflow to engine core and bypass airflow through bypass duct. Engine has mixer for exhaust gas flow exiting engine core, bypass airflow exiting bypass duct, thrust nozzle for discharging mixed flows, and controller for thrust produced by engine. To change level of engine thrust between transonic push operation and supersonic cruise operation, controller adjusts one or more components which vary relative areas available for hot exhaust gas flow and cold bypass airflow at mixer while holding fan inlet non-dimensional mass flow w ?{square root over (?)} T/P substantially constant, where w is mass flow of incoming air at fan inlet, T is stagnation temperature of incoming air at fan inlet and P is stagnation pressure of incoming air at fan inlet.

Abstract: An aircraft defining longitudinal, lateral and vertical directions the aircraft comprising: a main wing and a tail, each being pivotable about the lateral direction (B); a plurality of main propellers mounted to the main wing, and configured to pivot with the main wing; at least one cruise propeller mounted to the tail, and configured to pivot with the tail; each main propeller being stowable from a deployed position to a stowed position; wherein each main propeller has a fixed pitch, and each cruise propeller has a variable pitch.

Abstract: A gas turbine engine for an aircraft, comprising: an engine core; a covering configured to cover at least part of the engine core; and an engine accessory that interacts with the engine core or a component of the engine core; wherein the engine accessory is outside the covering.

Abstract: A gas turbine engine for an aircraft is provided. The engine includes an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The engine further includes a core casings surrounding the engine core. The engine further includes one or more engine accessories mounted adjacent to and vertically beneath the core casings. The engine further includes a heatshield positioned between the one or more engine accessories and the core casings. The heatshield contains one or more ventilation holes for channelling convectively driven flows of ventilation air around the one or more engine accessories on engine shutdown.

Abstract: A flap for a variable area exhaust nozzle of a gas turbine engine, having a support structure and a gas shield having a plurality of gas shield segments connected to the support structure and adjacently arranged along a longitudinal direction of the flap corresponding to flow through the nozzle. Each of the gas shield segments has a flowpath section arranged to cooperate with a corresponding flowpath section of an adjacent gas shield segment to define a flowpath surface of the gas shield to bound flow through the nozzle. The gas shield segments are configured to permit the respective flowpath section to move longitudinally relative the support structure in response to thermal expansion of the gas shield.

Abstract: A rear outer discharge nozzle for a gas turbine engine, the rear outer discharge nozzle forming a loop around an axis and comprising: a first part; a second part wrapped around the first part; and a radial slit, having first and second side walls; wherein the first part is shaped to provide the first side wall of the radial slit, and the second part is shaped to provide the second side wall of the radial slit; and wherein the second part comprises a break, splitting the second part in the axial part.

Abstract: A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

Abstract: A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

Abstract: A gas turbine engine combustion arrangement includes an annular combustion chamber, an annular discharge nozzle, an array of turbine nozzle guide vanes and an annular casing. The nozzle is between the downstream end of an outer annular wall of the annular combustion chamber and the array of turbine nozzle guide vanes. The annular casing surrounds the annular combustion chamber, the annular discharge nozzle and the array of turbine nozzle guide vanes. The upstream end of the annular discharge nozzle has an annular axially extending slot and the downstream end of the outer annular wall of the annular combustion chamber locates in the annular axially extending slot in the upstream end of the nozzle. The upstream ends of the radially outer platforms of the turbine nozzle guide vanes are adjacent the downstream end of the annular discharge nozzle and a support structure is connected to the nozzle and the annular casing.

Abstract: A nuclear power plant comprises a nuclear reactor, the nuclear reactor comprising reactor fuel elements, a reactor vessel surrounding the nuclear reactor and a primary shield surrounding the reactor vessel. The reactor fuel elements are arranged between a first height and a second height above the first height. The primary shield comprises a base portion, an intermediate portion and a top portion. The base portion has an upper height below the first height and the base portion comprises concrete. The top portion has a lower height above the second height and the top portion comprises concrete. The intermediate portion is arranged vertically between the base portion and the top portion of the primary shield. The intermediate portion comprises at least one support structure and a matrix material containing tungsten or boron, and the least one support structure extends between the top portion and the bottom portion of the primary shield.

Abstract: An exhaust nozzle assembly for a gas turbine engine. The assembly includes concentrically arranged inner mixer and outer exhaust nozzles, the exhaust nozzle extending axially downstream of said mixer nozzle. A centre-body is axially mounted within and extends axially downstream from the mixer nozzle. A core flow duct is defined by the mixer nozzle and the centre-body, the core flow duct having a core exit area. An exhaust duct is defined at least in part by the exhaust nozzle downstream of the mixer nozzle, the exhaust duct having an exhaust exit area. The mixer nozzle includes a mixer cowl which is axially-translatable along the centre axis and the exhaust nozzle includes an exhaust cowl which is either axially-translatable along or angularly-adjustable relative to the centre axis. The assembly further includes an actuation mechanism and the mixer cowl and exhaust cowl are movable by the actuation mechanism.

Abstract: An aircraft defining longitudinal, lateral and vertical directions the aircraft comprising: a main wing and a tail, each being pivotable about the lateral direction; a plurality of main propellers mounted to the main wing, and configured to pivot with the main wing; at least one cruise propeller mounted to the tail, and configured to pivot with the tail; the main propellers defining a swept disc area (Adisc), and the main wing defines a wing area (Awing); wherein a ratio of the disc swept area to the main wing area (Adisc:Awing) is between 1 and 2.

Abstract: Apparatus for a gas turbine engine, the apparatus comprising: a low pressure compressor; a high pressure compressor; a first electrical machine configured to provide torque to the low pressure compressor; a second electrical machine configured to receive torque from the high pressure compressor; and wherein the high pressure compressor does not comprise a sub-idle bleed valve.

Abstract: An aerofoil, comprises: a wall comprising an external surface, and an internal surface defining a cavity for receiving a cooling fluid in use; a first trench formed in the external surface, and a second trench formed in the external surface; a first passageway extending from a first passageway inlet in the cavity to a first passageway outlet in the first trench, and a second passageway extending from a second passageway inlet in the cavity to a second passageway outlet in the second trench; wherein the first passageway and second passageway intersect within the wall.

Abstract: A method of diffusion bonding two components together comprises providing a first component having a first bonding surface, and a second component having a second bonding surface. Each of the first bonding surface and the second bonding surface is etched. A cold working process is applied to each of the first bonding surface and the second bonding surface. Each of the first bonding surface and the second bonding surface is then etched. The first component is positioned adjacent to the second component with the first bonding surface abutting against the second bonding surface, to define a joint surface between the first component and the second component. A peripheral edge of the joint surface is sealed. The first bonding surface is diffusion bonded to the second bonding surface.

Abstract: A method of operating a gas turbine engine compressor. The engine comprises a compressor having an environmental control system bleed port having an outlet in fluid communication with an aircraft environmental control system air duct, and an air turbine starter configured to rotate a compressor shaft of the gas turbine engine. The air turbine starter has an inlet in fluid communication with the environmental control system air duct via an air turbine valve. The method comprises determining a surge margin of the compressor, and where the surge margin of the compressor is determined to be below a predetermined minimum surge margin, opening the air turbine valve to supply air to the air turbine.

Abstract: There is disclosed a component arranged to extend across a bypass duct (22) of a gas turbine engine (10), comprising: an outlet guide vane (24); and a structural support (26) arranged to structurally support a radially inner component of the gas turbine engine (10). The outlet guide vane (24) and the structural support (26) are contiguous at a merged portion (32) which is at a radially outer region of the component.

Abstract: A gas turbine engine for an aircraft comprises: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a core casing surrounding the engine core. The gas turbine engine further comprises a core cowl surrounding the engine core and the core casing. The gas turbine engine further comprises an engine accessory gearbox driven by a take-off from the core shaft. The gas turbine engine further comprises an oil system having one or more oil pumps powered by the engine accessory gearbox for circulating lubricating oil around components of the engine including the engine accessory gearbox, and having an oil tank for receiving and storing oil scavenged from the engine components before recirculation thereto.