Abstract: A noise mitigation system for an aircraft comprises a cyclic pitch mechanism arranged to apply a cyclic pitch schedule to rotor blades of an unducted propulsive rotor (UPR) under control of a processor. The processor receives input data corresponding to the position and attitude of the UPR, the position of one or more ground points stored in a memory and the velocity of the aircraft. If the processor determines that a ground point will enter the plane of the UPR, a control signal is output to the cyclic pitch mechanism as necessary to adjust the phase of the cyclic pitch schedule such that the azimuthal position of the ground point on entry to the plane of the UPR lies within an azimuthal interval over which the blade pitch of the cyclic pitch schedule is below its average value, thus reducing noise in the direction of the ground point.

Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Abstract: A system for supplying lubricant to a component of a gas turbine engine is provided. The system includes a first fluid circuit having a first pump drivably couplable to a fan shaft of the engine for pumping a lubricant from a lubricant tank to the component. The first fluid circuit allows only a unidirectional flow of the lubricant from the lubricant tank to the component via the first pump. The system also includes a second fluid circuit having a second pump drivably couplable to the fan shaft for pumping the lubricant from the component to the lubricant tank. The second fluid circuit allows only a unidirectional flow of the lubricant from the component to the lubricant tank via the second pump. First and second ports of the two pumps operate as input ports or output ports depending on the rotation direction of the fan shaft.

Abstract: An aircraft hybrid propulsion system (5) comprising an inboard gas turbine engine (10a, 10c) and an outboard gas turbine engine (10b, 10d), each comprising a propulsor (12a, 12b) and a respective electric machine (32a, 32b) coupled to one or more engine shaft (24a, 24b). An electrical interconnection (34) is provided between the electric machine (32a) of the inboard gas turbine engine (10a) and the electric machine (32b) of the outboard gas turbine engine (10b). A controller (36) is configured to transfer electrical power between the inboard gas turbine engine electrical machine and the outboard gas turbine engine electrical machine when a thrust setting change is selected.

Abstract: An inspection system for mounting on a user's hand. The inspection system comprising: an imaging unit comprising two sub-units, the first sub-unit being configured to provide images from a first point of view and the second sub-unit being configured to provide images from a second point of view; and a measuring unit configured to provide data relating to a physical property measured at a measurement location on the user's hand. The imaging unit has a separation sensor configured to measure the separation between the two sub-units of the imaging unit. A method of inspecting and/or servicing a machine is also disclosed.

Abstract: A fluid pump is shown, comprising: a chamber comprising an inlet and an outlet, the outlet comprising a non-return valve, the chamber having a cavity comprising a cylinder; a piston slidably disposed within the cylinder; and a Tesla valve in fluid communication with the inlet, wherein the fluid pump is configured to pump fluid from the inlet to the outlet by reciprocation of the piston within the cylinder.

Abstract: A fuel delivery system for a gas turbine engine comprises a cryogenic fuel tank, a first fuel line for connection to the cryogenic fuel tank, a fuel pump connected to receive fuel via the first fuel line, a plurality of fuel lines connecting the fuel pump to a combustor of the gas turbine engine, a controller configured to operate the fuel delivery system, a purge gas tank connected to the first fuel line and configured to store a purge gas for purging the plurality of fuel lines and a fuel gas tank connected to the first fuel line and configured to store a fuel gas for flushing purge gas from the plurality of fuel lines.

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: Fault-tolerant electric drive systems including a machine having a rotor and a stator having coils arranged in pairs. Each coil in each pair separated by 180 degrees, a first phase (?A) having one of the coil pairs and a phase drive circuit connected therewith, a second phase (?B) having a second one of the coil pairs and a second phase drive circuit connected therewith, a third phase (?C) having a third one of the coil pairs and a third phase drive circuit connected therewith, and a fourth phase having a fourth one of the coil pairs and a fourth phase drive circuit connected therewith. Further included is a controller connected with the first, second, third and fourth phase drive circuits to control operation thereof.

Abstract: A centrifugal compressor is shown, comprising: an impeller; a rotatably mounted diffuser surrounding the impeller; and a driving arrangement configured to drive the impeller and diffuser to rotate in opposing directions.

Abstract: A method for testing of a population of monocrystalline components is provided. The method includes obtaining a plurality of component parameters including a crystal angle of each monocrystalline component with respect to a coordinate axis, a three-dimensional geometry, and a material. The method further includes determining a statistical parameter of the crystal angle, and generating a simulation model of the monocrystalline component based on the statistical parameter, the three-dimensional geometry, and the material. The method further includes determining at least one probe parameter based on the simulation model and a predetermined region of interest. The method further includes determining anisotropic delay laws based on the statistical parameter and the probe parameter, and controlling at least one probe based on the anisotropic delay laws to emit ultrasonic waves towards the region of interest in order to test the monocrystalline component for one or more abnormalities.

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A gas turbine engine comprising a variable geometry combustor having pilot fuel injectors and main fuel injectors; a fuel metering system configured to control fuel flow to the pilot fuel injectors and the main fuel injectors; a variable geometry airflow arrangement for the variable geometry combustor, which is configured to vary the airflow through the pilot fuel injectors and/or the main fuel injectors; a control system configured to control the variable geometry airflow arrangement in dependence upon airflow delivered to the combustor, the fuel flow to the pilot fuel injectors and the main fuel injectors, and a target index of soot emissions, thereby controlling airflow through the pilot fuel injectors and/or the main fuel injectors and hence the quantity of soot produced by combustion.

Abstract: A lean burn combustor includes a plurality of lean burn fuel injectors, each including a fuel feed arm and a lean burn fuel injector head with a lean burn fuel injector head tip, wherein the lean burn fuel injector head tip has a lean burn fuel injector head tip diameter, the lean burn fuel injector head including a pilot fuel injector and a main fuel injector, the main fuel injector being arranged coaxially and radially outwards of the pilot fuel injector; and a combustor chamber extending along an axial direction for a length and including a radially inner annular wall, a radially outer annular wall, and a meter panel defining the size and shape of the combustor chamber, wherein the combustor chamber includes primary and secondary combustion zones. A ratio of the combustor chamber length to the lean burn fuel injector head tip diameter is less than 5.

Abstract: Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines. A gearbox for an aircraft gas turbine engine includes: 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: A gas turbine engine comprising a variable geometry combustor having fuel injectors, a rich-burn zone, a quick-quench zone, and a lean-burn zone, and further comprising quench ports for admitting quench air to the quick-quench zone; a variable geometry airflow arrangement for the variable geometry combustor, which is configured to vary an airflow through the fuel injectors and/or the quench ports; and a control system configured to control the variable geometry airflow arrangement in dependence upon an airflow delivered to the combustor, a fuel flow to the fuel injectors, and a target index of soot emissions to control the quantity of soot produced by combustion.

Abstract: A combustion chamber (15) comprising a wall at least partially defining a combustion zone and having a first surface (41) facing away from the combustion zone and a second surface (43) facing the combustion zone, the wall having at least one effusion cooling aperture (69, 73) extending there-through from the first surface to the second surface, the effusion cooling aperture having an inlet in the first surface and an outlet in the second surface, the first surface having a particle separator (84) at least partially located upstream of the inlet of the effusion cooling aperture, the particle separator projecting away from the first surface and away from the combustion zone.

Abstract: A method of manufacturing a planet carrier of a gearbox comprises manufacturing a plurality of preforms. Each preform comprises a base portion, a first end portion connected to a first end of the base portion, a second end portion connected to a second end of the base portion, a first side portion connected to a first side of the base portion, a second side portion connected to a second side of the base portion and a top portion. The first and second end portions and the first and second side portions are folded relative to the base portion and the top portion is folded and the adjacent edges of the portions of each preform are secured together to form a support structure. A first ring and a second ring are manufactured. The first and second end portions of each support structure are secured to the first and second rings respectively.

Abstract: A radial-flux rotary electric machine having a first end at one axial extreme and a second end at another axial extreme, comprising: a fluid input for receiving coolant at the first end of the machine; a rotor comprising a radial impeller at the first end of the machine in fluid communication with the fluid input, and a plurality of permanent magnets with a helical fluid conduit thereunder in fluid communication with the impeller; a stator having a plurality of slots with a fill factor of between zero and unity to define axial fluid channels in the slots, the fluid channels being in fluid communication with the helical fluid conduit at the second end of the machine; a fluid output in fluid communication with the fluid channels.

Abstract: A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.