Abstract: There is disclosed a gas turbine engine comprising: an electrically-powered motor configured to rotate a driven spool of the gas turbine engine; a primary controller configured to conduct a primary function using the motor; and a bow controller configured to selectively control the motor to perform a rotor bow mitigation operation in which the motor drives the driven spool to rotate to mitigate a non-uniform thermal distribution in a rotor of a spool of the gas turbine engine.

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 fuel management system includes: a fuel tank; a fuel supply line that supply fuel from the fuel tank to a combustor of the gas turbine engine; a reheat fuel supply line that supply fuel from fuel tank to a reheat of the gas turbine engine, the reheat fuel supply line extending from a reheat branching point on the fuel supply line to the reheat; a fuel supply pump disposed along fuel supply line upstream of the reheat branching point; a reheat pump disposed along reheat fuel supply line, the reheat pump that pressurise fuel to a reheat delivery pressure for delivery to the reheat; and a reheat recirculation line that recirculate fuel from the reheat fuel supply line to a location upstream of fuel supply pump, the reheat recirculation line extending from a reheat recirculation branching point on the reheat fuel supply line downstream of the reheat pump.

Abstract: A thermal regulating unit for regulating the temperature of a pouch cell battery is provided. The thermal regulating unit is formed as a container having: one or more internal cooling channels for conveying a liquid coolant through the unit; a flexible outer covering which contains the cooling channels; and inlet and outlet ports which penetrate the covering for respectively providing liquid coolant to and removing the liquid coolant from the cooling channels. The flexible outer covering forms a substantially flat major external surface of the unit corresponding in shape to, and for pressing against, a major external surface of the pouch cell battery such that the unit and the pouch cell battery can be held in face-to-face contact. The cooling channels are arranged in the flexible outer covering such that when the provided liquid coolant is pressurized it causes the unit to expand and press against the major external surface of the pouch cell battery.

Abstract: A method for manufacturing a fuel spray nozzle for a gas turbine engine includes forming a first section of the fuel spray nozzle by additive layer manufacturing. The first section includes a main chamber and internal passageways. The method includes forming a second section of the fuel spray nozzle by additive layer manufacturing on the first section. The second section includes a metering feature disposed in fluid communication with one of the internal passageways of the first section. The method includes modifying the metering feature of the second section by a first subtractive manufacturing process to obtain a desired diameter of the metering feature and/or a desired surface roughness of the metering feature. The method includes forming a third section of the fuel spray nozzle by additive layer manufacturing on the second section. The third section includes a deflector and a spin chamber.

Abstract: A combustor assembly for a gas turbine engine, the combustor assembly including: combustor wall defining combustion chamber and including opening periphery, the opening periphery defining first opening; sealing element including annular body, annular body extending around sealing element axis and extending through first opening, sealing element further including flange extending radially outward from annular body, flange slidingly coupling sealing element and combustor wall and forming seal or first partial seal between sealing element and combustor wall, annular body including outer and inner surface, inner surface defining second opening configured to receive fuel nozzle, wherein first total clearance between outer surface and opening periphery in first direction is greater than second total clearance between outer surface and opening periphery in second direction perpendicular to first direction such that sealing element is able to slide relative to combustor wall by greater extent in first direction than

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines, first thermal bus and heat exchanger. The first thermal bus includes a first heat transfer fluid in a closed loop flow sequence, between the first gas turbine engine, or each first electric machine, and the first heat exchanger. Waste heat energy transfers to the first heat transfer fluid. The first heat exchanger configures to transfer waste heat energy from the first heat transfer fluid to a dissipation medium. During steady-state operation of the first gas turbine engine, the first heat transfer fluid entering the first heat exchanger has a temperature of TFLUID(° C.), and a temperature of an inlet air flow entering the first gas turbine engine has a temperature TAIR(° C.) and a ratio B is in a range of between 5.0-18.0.

Abstract: Reconfigurable permanent magnet electrical machines and aircraft power and propulsion systems including the electrical machines. An aircraft power and propulsion system includes: a gas turbine engine; DC electrical network; permanent magnet electrical machine including a rotor drivingly coupled to a spool of the engine, and a stator including windings controllably switchable between a star and a delta configuration; an AC-DC power electronics converter, an AC side is connected to terminals of the stator windings and a DC side is connected to the DC electrical network; an additional electrical power source connected to and controllable to supply electrical power to the DC electrical network; and a control system configured to control the switching of the stator windings between the configurations and to control the additional electrical power source to supply electrical power to the DC electrical network during a time interval when the stator is being switched between the configurations.

Abstract: A gas turbine engine for an aircraft includes: an engine core; fan; turbomachinery bearings; power gearbox; and a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings and including at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink, wherein the heat management system is configured to provide a first proportion of heat generated by the power gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed in the range of 0.25 to 0.70, and a second proportion of heat generated by the power gearbox and the turbomachinery and dissipated to air at 65% of the core shaft maximum take-off speed in the range of 0.60 to 1.

Abstract: Integrated arrangements of electrical machines and power electronics converters are described. One such arrangement comprises: an electrical machine comprising one or more windings; a power electronics converter arranged to supply current to or receive current from the one or more windings of the electrical machine; a magnetocaloric effect (MCE) material in thermal contact with the power electronics converter; and a heat sink for removing heat from the MCE material. The MCE material is arranged in proximity to the one or more windings of the electrical machine whereby, in use, stray magnetic flux from the windings of the electrical machine passes through the MCE material and activates the MCE material. The repeated application and removal of the stray flux during normal operation of the electrical machine creates cycles of magnetic refrigeration, which removes heat from the power electronics converter.

Abstract: A sensor assembly is shown for sensing a crossing of the critical point in a system utilising a working fluid in a transcritical cycle passing through the critical point. A first broadband acoustic sensor is located upstream of a component and a second broadband acoustic sensor is located downstream of the component, each of which are arranged to detect high-frequency and low-frequency sounds caused by the crossing of the critical point. A flow regulation device regulates flow of working fluid through the component in response to the output of one or both of the first broadband acoustic sensor and the second broadband acoustic sensor, thereby adjusting the location of the crossing of the critical point.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines, a first thermal bus, and a first heat exchanger. The first thermal bus includes a first heat transfer fluid in a closed loop flow sequence, between the first gas turbine engine, the or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to first heat transfer fluid. When airspeed of aircraft is less than Mn0.6, the first heat exchanger transfers the waste heat energy from the first heat transfer fluid to a first dissipation medium. When the airspeed of the aircraft is greater than Mn0.6, the first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a second dissipation medium.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger. Waste heat energy generated by at least one first gas turbine engine, and first electric machine, transfers to the first heat transfer fluid. The first heat exchanger directs a first proportion of the first heat transfer fluid through a first heat dissipation portion wherein a first proportion of the waste heat energy transfers to a first dissipation medium dependent on the first dissipation medium temperature and mass flow rate. The first heat exchanger directs a second proportion of the first heat transfer fluid through a second heat dissipation portion wherein the second proportion of waste heat energy transfers to a second dissipation medium dependent on the second dissipation medium temperature and mass flow rate.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, a first thermal bus, a first heat exchanger, and a chiller. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the first heat exchanger, and the chiller. Waste heat energy generated by the first gas turbine engine, is transferred to the first heat transfer fluid. The chiller is configured to lower a temperature of the first heat transfer fluid prior to the first heat transfer fluid being circulated through the gas turbine engine. The first heat is exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger module. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to the first heat transfer fluid. The first heat exchanger module comprises a first flow path and a second flow path.

Abstract: A thermal management system for an aircraft includes a first gas turbine engine, first thermal bus, first heat exchanger, one or more first ancillary systems, vapour compression system, one or more second ancillary systems and second heat exchanger. A waste heat energy generated by a first gas turbine engine, and a first ancillary system, transfers to the first heat transfer fluid. A waste heat energy generated by a second ancillary system transfers to a second heat transfer fluid, and the second heat exchanger transfers the waste heat energy from the second heat transfer fluid to the first heat transfer fluid. The waste heat energy generated by a second ancillary system transfers to the first heat transfer fluid, and the first heat exchanger transfers the waste heat energy to a dissipation medium. The waste heat energy transferred to the second heat transfer fluid ranges from 20 kW to 300 kW.

Abstract: A method of operating an aircraft including a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine. The method includes: determining at least one fuel characteristic of the fuel arranged to be provided to the gas turbine engine; and proposing or initiating a change to a flight profile of the aircraft based on the at least one fuel characteristic. For example, intended route and/or altitude of the aircraft may be changed based on the one or more fuel characteristics.

Abstract: A method of manufacturing a component includes supporting the component on a mounting structure. The method further includes impacting first peening bodies on the component to increase sub-surface compressive residual stresses of the component. The first peening bodies impact the component at a first intensity of from about 0.45 mmA to about 0.61 mmA. A first coverage of the component by the first peening bodies is from about 120% to about 205%. The method further includes impacting second peening bodies on the component to increase surface compressive residual stresses of the component. The second peening bodies impact the component at a second intensity of from about 0.20 mmA to about 0.30 mmA. A second coverage of the component by the second peening bodies is from about 120% to about 205%. Each of the first peening bodies and the second peening bodies includes a ceramic material.

Abstract: A thermal management system for an aircraft includes a first thermal bus including one or more first heat sources, a heat sink, a vapour compression system, and one or more second heat sources. The vapour compression system includes a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor. A first heat flow (Q1) of waste heat energy generated by the first heat sources is transferred via the first heat transfer fluid to the heat sink. A second heat flow (Q2) of waste heat energy generated by the second heat source(s) being transferred via the evaporator to a refrigerant. A third heat flow (Q3) of heat energy in the refrigerant is transferred via the condenser to the first heat transfer fluid. The controller is configured to ensure that: 1.1*Q2<Q3<3.0*Q2.

Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, a first heat exchanger, and one or more first ancillary systems. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the or each first electric machine, the first gas turbine engine, the first heat exchanger, and the or each first ancillary system. Waste heat energy generated by at least one of the first gas turbine engine, the or each first electric machine, and the or each first ancillary system, is transferred to the first heat transfer fluid. The first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.