Abstract: An intercomponent seal arrangement includes a plurality of wall segments extending circumferentially. Adjacent wall segments from the plurality of wall segments include opposing first and second end wall portions defining a gap therebetween. The gap includes a saddle portion facing radially outwards and including a first sealing face on the first end wall portion and a second sealing face on the second end wall portion. The intercomponent seal arrangement further includes an intercomponent seal located within the saddle portion and including at least a first contact point contacting the first sealing face, a second contact point contacting the second sealing face, and a third contact point radially inwards from the first and second contact points and contacting one of the first and second sealing faces.

Abstract: A needle apparatus for making a pinning hole in a composite material, the apparatus having a needle with a longitudinal axis, a collar fixed about the needle, a motor having an axis of rotation, and being fixed to the needle such that the needle can be rotated by the motor, the motor and the collar being situated within a motor housing, the motor housing having a first bush, and a second bush, wherein the collar is positioned between the first bush and second bush so as to prevent the needle from moving away from or towards the motor, but allow the needle to rotate freely around its longitudinal axis, wherein each needle apparatus further comprises a linear actuator connected to the motor such that the motor can be moved along the axis of rotation of the motor by the linear actuator.

Abstract: A method of operating a gas turbine engine and a gas turbine engine includes a fuel delivery system arranged to provide fuel, a combustor arranged to combust at least a proportion of the fuel, a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough, and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. A fuel viscosity is adjusted to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.

Abstract: A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is EI maxTO , SAF EI maxTO , FF . EImaxTO,SAF is nvPM emissions index in mg/kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles includes sustainable aviation fuel. EImaxTO,FF is nvPM emissions index in mg/kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles is fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio is less than 1.

Abstract: There is provided a method of monitoring the health of an asset. The method comprises obtaining a measured value of an input parameter representing an operational state of the asset, determining, based on the value, a residual that represents the behaviour of the asset, providing the determined residual to an encoder model to map the determined residual to a data point in a feature space, determining a distance of the data point relative to a cluster of data points in the feature space, wherein the cluster of data points are determined by the encoder based on residuals representing nominally healthy assets, and in response to the determined distance exceeding a predetermined distance threshold, generating an abnormal asset behaviour prediction.

Abstract: There is provided a method of monitoring the health of an asset. The method comprises providing an input representing the behaviour of the asset to an encoder model to map the input to a data point in a feature space, wherein the encoder model is configured to cause data points relating to measured values of input parameters acquired from the asset at a similar time to cluster in feature space, determining a distance of the data point relative to a cluster of data points in feature space, where the cluster of data points are determined by the encoder based on inputs representing the asset over a first time period, and in response to the determined distance exceeding a predetermined distance threshold, generating an abnormal asset behaviour prediction.

Abstract: There is provided a method of training an encoder model to map inputs representing nominally healthy assets to a cluster of data points. The method comprises providing an input representing a behaviour of an asset at a first point in time to the encoder model undergoing training to map the input to a data point in the feature space, wherein the encoder model is intended to cause data points relating to measured values of input parameters acquired from the same asset at similar time to cluster in feature space, providing the data point to a decoder model, to determine a reconstructed input, determining a difference between the input, and the reconstructed input, and training the encoder model based on the determined difference to map inputs representing nominally healthy assets acquired at similar times to a cluster of data points.

Abstract: A method of operating a gas turbine engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a combustor arranged to combust a fuel; a fan located upstream of the engine core; an oil system arranged to circulate oil; and a heat exchange system comprising at least one fuel-oil heat exchanger arranged to transfer heat from the oil to the fuel, wherein the method comprises: determining at least one fuel characteristic of the fuel arranged to be provided to the combustor; and modulating the heat exchange system so as to adjust the fuel temperature on entry to the combustor at cruise to a set level, the set level being based on the at least one fuel characteristic.

Abstract: A method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method includes: exposing the surface of a piezoelectric crystal to the fuel; measuring a vibration parameter of the piezoelectric crystal; and determining one or more fuel characteristics of the fuel based on the vibration parameter. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.

Abstract: A strut for a bearing assembly of a gas turbine engine. The bearing assembly has a first bearing structure and a second bearing structure. The strut includes a first mounting end portion, a second mounting end portion spaced apart from the first mounting end portion, and a shaft portion extending between the first mounting end portion and the second mounting end portion along a longitudinal axis. The first mounting end portion is configured to be attached to the first bearing structure thereby forming an interference fit between the first mounting end portion and the first bearing structure. The second mounting end portion is configured to be attached to the second bearing structure. The first mounting end portion includes at least one relief feature.

Abstract: A fuel system for a hydrogen fuelled gas turbine engine. The fuel system having a main hydrogen fuel line configured to provide hydrogen fuel to a combustor of the gas turbine engine and a closed loop heat exchange fluid line having a heat exchange fluid. The fuel system further having a first heat exchanger configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line, a second heat exchanger downstream of the first heat exchanger in heat exchange fluid flow configured to exchange heat between the heated heat exchange fluid and hydrogen in the main hydrogen fuel line, and a third heat exchanger downstream in main hydrogen fuel flow of the second heat exchanger in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.

Abstract: A fuel system for a hydrogen fuelled gas turbine engine. The fuel system having a main hydrogen fuel line configured to provide hydrogen fuel to a combustor of the gas turbine engine and a closed loop heat exchange fluid line having a heat exchange fluid. The fuel system further having a first heat exchanger configured to exchange heat between gas turbine engine compressor bleed air and heat exchange fluid in the heat exchange fluid line, a second heat exchanger downstream of the first heat exchanger in heat exchange fluid flow configured to exchange heat between the heated heat exchange fluid and hydrogen in the main hydrogen fuel line, and a third heat exchanger downstream in main hydrogen fuel flow of the second heat exchanger in main hydrogen fuel flow configured to exchange heat between a further engine fluid and heated hydrogen fuel.

Abstract: An 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 located upstream of the engine core, the fan having 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 method for scanning a plurality of components includes providing an imaging beam source, an imaging beam receiver, and a support platform. The support platform is configured to rotate and/or revolve relative to the imaging beam source and the imaging beam receiver about one or more axes. The method further includes disposing the plurality of components on the support platform. The plurality of components is positioned in a gap between a first geometrical figure and at least one second geometrical figure in a configuration that reduces the variation in material thickness penetrated at different relative angles of rotation and/or revolution. The first geometrical figure and the at least one second geometrical figure are concentric.

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, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Abstract: A method of operating a gas turbine engine for an aircraft is disclosed. The aircraft includes a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system arranged to supply oil to the gearbox; a heat exchange system having an air-oil heat exchanger; a fuel-oil heat exchanger through which the oil and the fuel flow; and a valve arranged to allow a proportion of the oil sent via at least one of the heat exchangers to be varied. The method includes controlling the valve such that, under idle conditions, an oil flow ratio of: rate ? of ? oil ? flow ? into ? air - oil ? heat ? exchanger r ? ate ? of ? oil ? flow ? into ? fuel - oil ? heat ? exchanger is in the range from 0.62 to 5.29.

Abstract: A method of operating a gas turbine engine. The gas turbine engine includes an accessory gearbox arranged to power one or more engine accessories, and a combustor in which fuel is combusted and a fuel-oil heat exchanger arranged to transfer heat between oil and fuel that is provided to the combustor. An average temperature of lubricating oil is at least 180° C. on exit from the accessory gearbox at cruise conditions. The method includes transferring heat from the oil to the fuel before the fuel enters the combustor so as to raise the fuel temperature to an average of at least 135° C. on entry to the combustor at cruise conditions. Also provided is a gas turbine engine.

Abstract: Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a bearing.

Abstract: Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a bearing.

Abstract: A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.