Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, compressor, and core shaft connecting turbine to the compressor; a fan located upstream of engine core, the fan containing a plurality of fan blades mounted for rotation about an engine axis, each blade of the plurality of fan blades having a leading edge and trailing edge extending across a span of an airflow duct from blade root to tip; and a gearbox that receives an input from core shaft and outputs drive to fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the trailing edge of each blade is characterized by a metric M defined as a rate of change of an angle of trailing edge between 0.4 and 0.8 of the span divided by an area averaged trailing edge angle, M being not less than around 4.

Abstract: A tool for machining an object including a first part including a rotatable member that is rotatable to cause rotation of a machine tool, a second part, a joint coupling the first part and the second part to enable relative movement between the first part and the second part, and a sensor to sense an object to be machined.

Abstract: A computer-implemented method. The method comprising: (a) obtaining a candidate characterising code, the candidate characterising code being indicative of character of an example of a physical system; (b) accessing a database of extant characterising codes, the extant characterising codes being indicative of character of other examples of the physical system; (c) determining a degree of similarity between the candidate characterising code, and at least one of the extant characterising code; and (e) providing, as an output, the results of the comparison.

Abstract: A computer-implemented method. The method comprising: (a) obtaining a plurality of unique identifiers, each unique identifier being associated with a respective characterising code, each characterising code being indicative of character of an example of a same physical system; (b) accessing a database of characterising codes, and retrieving each respective characterising code; (c) determining a degree of similarity between each retrieved characterising code; and (d) providing, as an output, the determined degrees of similarity.

Abstract: A gas turbine engine for an aircraft has an engine core comprising turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine, and having turbine blades, and the compressor being the lowest pressure compressor of the engine; fan located upstream of the engine core; and gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, the three bearings having two rearward bearings located downstream of the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade, and/or, when the turbine comprises four sets of turbine blades, downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

Abstract: A gas turbine engine for an aircraft includes an engine core including an engine core, a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage. A fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.

Abstract: A combustion staging system has: a splitter; pilot and mains fuel manifolds; mains flow control valves; and fuel servo line. Each mains flow control valve has a chamber containing a piston, the chamber to a piston mains side fed by the mains fuel manifold, and the chamber to a piston servo side fed by the servo line. The piston has an open pilot-and-mains position allowing flow from the chamber mains side to the respective injector mains discharge orifice. The piston prevents flow from the chamber mains side. The piston is movable under a pressure change in the servo line relative to the mains fuel manifold. The system has a servo pump and a hydraulic motor driving it. The servo pump changes fuel pressure. Motive power for the hydraulic motor is fuel diverted from a fuel pump high pressure output, the motor returning the diverted fuel to a low pressure input.

Abstract: An aircraft comprising a wing having a spanwise lift distribution extending from a root to a tip, the lift distribution defining an inboard region defining a positive lift contribution, an outboard region defining a negative lift contribution, and an intermediate region defining a neutral lift contribution, the neutral region being spaced from the tip and from the root. A propulsion system is provided, comprising a wing mounted propulsor. The wing mounted propulsor has a rotational axis (x) positioned substantially at a span of the wing where a value of ?Lift/?Span is at a maximum for the span of the wing, and may be located at the intermediate region along the span of the wing.

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, includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft, wherein the fan assembly has fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub having slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the fan blades is within the range of around 0.45 to around 0.7.

Abstract: The present disclosure concerns a heat exchanger, which may for example be utilised in a gas turbine engine or in other applications. Example embodiments include a heat exchanger comprising: an external surface for exchanging heat with an external fluid flow passing over the external surface; a first fluid passage extending through the heat exchanger from a first fluid inlet to a first fluid outlet, a first portion of the first fluid passage extending along the heat exchanger adjacent to the external surface for a first cooling fluid passing through the first fluid passage to exchange heat with the external fluid flow; and a second fluid passage extending through the heat exchanger from a second fluid inlet to a second fluid outlet located at the external surface for a second cooling fluid to pass from the second fluid inlet into the external fluid flow.

Abstract: A gas turbine engine for an aircraft includes a fan system having a reverse travelling wave first flap mode, Fan RTW, and including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode comprising a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first reverse whirl rotor dynamic mode, Rotor RW, and a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A backward whirl frequency margin of: the ? ? lowest ? ? frequency ? ? of ? ? either ? ? mode ? ? Fan ? ? RTW ? ? or ? ? Rotor ? ? RW ? ? at the ? ? MTO ? ? speed the ? ? MTO ? ? speed may be in the range from 15 to 50%.

Abstract: A gas turbine engine for an aircraft includes a fan system including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode including a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A forward whirl frequency margin of: the ? ? frequency ? ? difference ? ? between ? ? the ? ? intersection ? ? of ? ? 1 ? ? FW ? ? and the ? ? first ? ? engine ? ? order ? ? line ? ? and ? ? the ? ? first ? ? engine ? ? order at ? ? the ? ? MTO ? ? speed the ? ? MTO ? ? speed is in the range from 10 to 100%.

Abstract: A nacelle for a gas turbine engine includes a leading edge at an upstream side of the nacelle. The nacelle further includes a trailing edge at a downstream side of the nacelle. The nacelle further includes an outer surface disposed between the leading edge and the trailing edge. The nacelle further includes a concave section continuous with the outer surface and disposed proximal to the trailing edge. The concave section includes an upstream end and a downstream end spaced apart from the upstream end. The concave section is curved radially inwards relative to the outer surface between the upstream end and the downstream end.

Abstract: An aircraft gas turbine engine includes a fan system having a reverse traveling wave first flap mode, Fan RTW, and including a fan upstream of the engine core; a fan shaft; and a front engine structure to support the shaft and having a front engine structure nodding mode FSN including two modes at similar, but unequal, natural frequencies in orthogonal directions; and a gearbox. The engine includes a gearbox, and a gearbox output shaft to couple output of the gearbox to the fan shaft. An LP rotor system including the fan system and the gearbox output shaft has a first reverse whirl rotor dynamic mode, Rotor RW.

Abstract: A method of estimating a parameter of a fluid flowing in a passage includes: having a plurality of instruments operable to measure one or more fluid properties flowing in the passage, the plurality of instruments being disposed in the passage and arranged within a common measurement plane; assigning a stream tube to each instrument, each stream tube represents a region of space in the common measurement plane within the passage and each stream tube surrounds one of the plurality of instruments, the stream tubes together correspond to the cross-sectional shape and area of the passage in the common measurement plane; measuring the one or more fluid properties using the instruments to obtain one or more measured values for each stream tube; using the measured value(s) for each stream tube to calculate a derived value for each stream tube; and summing the derived values across all of the stream tubes.

Abstract: A method of controlling two or more sources of electrical power to supply electrical power to a shared bus. Each source of electrical power has a rated power Pr, a maximum overload power Pm, and is operable along a multi-voltage droop slope, the multi-voltage droop slope including a share region and a catch region, the share region being defined by an upper voltage limit, Vhigh and a base voltage Vbase, and the catch region being defined by the base voltage Vbase and a lower voltage limit Vlow.

Abstract: An aircraft gas turbine engine has an engine core having a turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine and having a turbine length being the distance between the roots of the most upstream turbine blade at its leading edge and of the most downstream turbine blade trailing edge, and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine; and a gearbox receiving an input from the core shaft and outputting drive to the fan. The engine core has three bearings to support the core shaft, the three bearings having a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings, and wherein further a minor span to turbine length ratio of: minor ? ? span tu ? ? rbine ? ? length is equal to or less than 1.05.

Abstract: A tail bearing housing of a turbomachine comprises an inner casing and an outer casing concentrically arranged along a longitudinal axis, and a plurality of vanes having a leading edge, a trailing edge, a first wall extending from the leading edge to the trailing edge and a second wall extending from the leading edge to the trailing edge, the first wall and the second wall defining an aerofoil, the vanes being connected at a first end to the outer casing and at a second end to the inner casing. At least one vane of the plurality of vanes is a vent vane comprising a first aperture arranged in the first end and a second aperture arranged in the aerofoil, the second aperture being configured to be in fluid communication with the first aperture through an inner cavity to vent bleed air.

Abstract: A rotor of an electric machine includes a main rotor body having a longitudinal axis, the main rotor body comprising: a hollow cylinder body with a first end and a second end; an integral end flange at the first end; a separate end flange disposed at the second end of the cylinder body; and laminations disposed on an outer surface of the cylinder body, the laminations comprising magnetic material.