Abstract: Electrical power systems for distributing electrical power in arrangements including one or more gas turbine engines; one such system includes: one or more rotary electric machines coupled to a gas turbine engine spool; a set of converter circuits connected to the rotary electric machines converting between alternating current (ac) and direct current (dc), wherein the rotary electric machines and the converter circuits output a number R?2 of dc power channels, each channel having an index r=(1, . . . , R); and a group of N dc load channels connected to the R dc power channels by a switching arrangement, wherein N>R and each dc load channel has an index n=(1, . . . , N). For each respective N load channels, a current limiting device (CLD) limits the current flowing from power channels to a load connectable to the electrical power system via the respective load channel.

Abstract: Electrical power systems for distributing electrical power in aircraft are described. One such electrical power system comprises: an electrical power source configured to output a number R?2 of dc power channels, each dc power channel having a respective index r=(1, . . . , R); and a group of N dc load channels connected to the R dc power channels by a switching arrangement, wherein each dc load channel has a respective index n=(1, . . . , N) and R<N?2R. The switching arrangement is operable to connect the r-th power channel to the n-th load channel according to the relationship: r ? ( n ) = { 1 , n = 1 n 2 ? and ? n 2 + 1 , 1 < n < N ? and ? n = 0 ? ? ( mod ? ? 2 ) n - 1 2 ? and ? n + 1 2 ? , 1 < n < N ? and ? n = 1 ? ? ( mod ? ? 2 ) R , n = N . FIG .

Abstract: Electrical power systems for distributing electrical power in aircraft are described. One such electrical power system comprises: an electrical power source configured to output a number R?2 of dc power channels, each dc power channel having a respective index r=(1, . . . , R); and a group of N dc load channels connected to the R dc power channels by a switching arrangement, wherein N>R and each dc load channel has a respective index n=(1, . . . , N). The system further comprises, for each respective one of a plurality of the N load channels, a current limiting device (CLD) operable to limit an amount of current flowing from the power channels to a load connectable to the electrical power system via the respective load channel. The electrical power source may comprise one or more batteries.

Abstract: A planetary gearbox having a sun gear, a plurality of planet gears, a ring gear, and a plurality of journal bearing pins. At its axially forward face side and/or at its axially rearward face side, each planet gear forms a recess that extends inside the planet gear starting from the face side. The planet gear forms a protrusion at its axially forward face side and/or at its axially rearward face side. The protrusion has an axial extent and a radial extent, wherein the radial extent is limited at the respective face side to the area between the planet gear inner bore and the radial inner limit of the respective recess in the planet gear, and wherein the protrusion is configured to provide for a flow path for oil emerging from the journal bearing that guides the oil away from the journal bearing.

Abstract: Electrical power systems for distributing electrical power in arrangements including one or more gas turbine engines; one such system includes: one or more rotary electric machines coupled to a gas turbine engine spool; a set of converter circuits connected to the rotary electric machines converting between alternating current (ac) and direct current (dc), wherein the rotary electric machines and the converter circuits output a number R?2 of dc power channels, each channel having an index r=(1, . . . , R); and a group of N dc load channels connected to the R dc power channels by a switching arrangement, wherein N>R and each dc load channel has an index n=(1, . . . , N). For each respective N load channels, a current limiting device (CLD) limits the current flowing from power channels to a load connectable to the electrical power system via the respective load channel.

Abstract: Performing multivariate spectral analysis to determine relationships between chemical species in a sample, includes: providing intensity measurement data as plural measured spectra from the sample, each spectrum having unique identifier and intensity values for bins of a binned spectral range; expressing the intensity measurement data as an m×n matrix V, m and n respectively represents number of bins of the spectral range and number of unique identifiers; performing non-negative factorisation of the matrix V to obtain an m×p derived spectra matrix W and a p×n spectral weightings matrix H minimizing an error function |V?WH|, p representing a number of derived spectra and is selected such that the non-negative factorisation over-fits WH to V; identifying correlations between the p derived spectra of the derived spectra matrix W which explain variance in the intensity measurement data; and determining chemical relationships between chemical species in the sample.

Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

Abstract: A gas turbine engine for an aircraft. The engine including: an engine core with a turbine, a compressor, a fan located upstream of the compressor and including a plurality of fan blades, and a core shaft connecting the turbine to the compressor; a gearbox which receives an input from the core shaft and outputs drive, via a driveshaft, to the fan so as to drive the fan at a lower rotational speed than the turbine, the drive shaft and core shaft forming a shaft system.

Abstract: An abradable sealing element comprises a substrate and a sealing structure. The sealing structure comprises one or more wall structures extending from the substrate and defining at least one open cell which is filled with abradable material. The one or more wall structures are formed by additive-layer, powder-fed, laser-weld deposition onto the substrate. The one or more wall structures are formed from nickel-based superalloy and constitute from about 10% to about 50% of the total volume of the sealing structure.

Abstract: There is provided a lay-up head for applying elongate fibre reinforcement material to an application surface. The lay-up head comprises a support head having a transverse axis, a dispensing mechanism carried by the support head for dispensing elongate fibre reinforcement material, and a roller for pressing elongate fibre composite material against the application surface. The roller is tiltable with respect to the support head such that the roller axis can be angled with respect to the transverse axis of support head. The lay-up head also comprises a holding mechanism operable to hold the roller in a central position in which the roller axis is parallel to the transverse axis of the support head.

Abstract: A fuel spray nozzle arrangement for a combustor, the fuel spray nozzle arrangement comprising a fuel spray nozzle connected to a feed arm, wherein the feed arm comprises an aerofoil.

Abstract: A method is provided of controlling a gas turbine having a shaft connecting a compressor to a turbine, as well as having a reheat system, and a gas turbine. The method includes the steps of: operating the engine using the reheat system to provide a mass flow rate of reheat fuel into a gas flow of the gas turbine engine downstream of an exit of the turbine; detecting a shaft break event in the shaft; and in response to this detection, maintaining the mass flow rate of the reheat fuel being provided into the gas flow downstream of the turbine exit, whereby the maintained mass flow rate of reheat fuel raises a back pressure downstream of the turbine and thereby reduces a rotational speed of the turbine.

Abstract: A highly efficient gas turbine engine is a system wherein the fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

Abstract: A gas turbine engine (100) for an aircraft comprises a pylon attachment (112) and a shaft (108) defining an engine centreline (110). The engine centreline lies in an engine central plane (120) which intersects the pylon attachment. The gas turbine engine comprises an intake (104) having a non-axisymmetric geometry and a medial plane (130) defining left and right halves of the intake. The left and right halves are configured for at least one of optimum cross wind performance, optimum incidence performance and optimum cruise performance when the medial plane is aligned with a vertical plane. The intake is installed so that the medial plane is angularly offset with respect to the engine central plane. The engine may be installed on a wing of an aircraft with the medial plane closer to its optimal orientation than is the case for a conventional 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 pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

Abstract: An aircraft gas turbine engine comprises at least a first main engine shaft an electric machine coupled to the main engine shaft, and a lubrication pump configured to provide oil to at least a part of the electric machine. The lubrication pump is controlled by a controller. The controller is configured to operate the lubrication pump in a main gas turbine engine running mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a first flow rate relative to the rotational rate of the main engine shaft, and a fault mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a second, higher flow rate relative to the rotational rate of the main engine shaft.

Abstract: There is disclosed a turbine comprising a stator vane ring 122 downstream of a rotor ring 116 of shrouded rotor blades 118. In vanes of the stator vane ring, an upstream portion of a mean camber line 136 diverges from a profile of the mean camber line at a half-span location, particularly to receive a tip leakage flow over the shrouded rotor blades 118.

Abstract: An aircraft propulsion system comprises a vertical lift propulsor mounted within a wing of the aircraft. The vertical propulsor comprises an electric motor comprising a rotor mounted to a plurality of propulsor blades, the motor further comprising a stator. The rotor is provided radially outwardly of the stator.

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.

Abstract: A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The diffuser further includes a diffuser angle (?diff), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (?diff) is from about 0 degrees to about 12 degrees.