Abstract: A heat management system includes a fuel tank storing a fuel; a first heat exchanger thermally coupled to the fuel tank; a hydraulic pump for circulating a hydraulic fluid; a hydraulic circuit including first and second hydraulic lines fluidly coupled to the first heat exchanger and the hydraulic pump, such that the first heat exchanger brings the hydraulic fluid and the fuel into a heat exchange relationship; an oil circuit; and a second heat exchanger thermally coupled to the oil circuit and at least one of the first and second hydraulic lines, such that the second heat exchanger brings the hydraulic fluid and the oil into a heat exchange relationship, thereby allowing heat transfer between the fuel and the oil via the hydraulic fluid.

Abstract: A vane assembly for a gas turbine engine which is a single unitary component that includes an aerofoil. A leading edge passageway is disposed proximal to a leading edge of the aerofoil and configured to receive a flow of a fluid therein. The vane assembly further includes a connecting passageway fluidly communicating the leading edge passageway with a trailing edge distribution passageway that is spaced apart from the leading edge, the leading edge passageway and a trailing edge of the aerofoil. The vane assembly further includes a plurality of trailing edge passageways disposed proximal to a pressure surface of the aerofoil and extending from the trailing edge distribution passageway towards the trailing edge along a chordwise direction. Each trailing edge passageway is configured to discharge the fluid through a corresponding passageway outlet disposed on the pressure surface and in fluid communication with a corresponding trailing edge passageway.

Abstract: A power system for an aircraft comprises a gas turbine engine, arranged to burn a fuel in a combustor so as to provide power to the aircraft; a plurality of fuel tanks arranged to contain fuel to be used to provide power to the aircraft; and a fuel manager. A first fuel tank of the plurality of fuel tanks is arranged to contain a first fuel, and a second tank of the plurality of fuel tanks is arranged to contain a second, different, fuel. The fuel manager is arranged to store information on the fuel contained in each fuel tank and to control fuel supply so as to take fuel from the second tank for engine start-up, before switching to the first fuel tank.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.

Abstract: A component for a hot section of an engine comprises a substrate and an infrared-reflective layer configured to reflect infrared radiation away from the substrate and in some cases the infrared-reflective layer has: a spectral reflectance R_? of no less than about 0.5 when measured by spectrophotometry, according to ISO 15368:2021, using incident electromagnetic radiation having a wavelength ? no less than about 500 nm and no greater than about 1 mm; and a spectral emissivity ?_? of less than about 0.4, for example no greater than about 0.3, when measured using incident electromagnetic radiation having a wavelength ? no less than about 500 nm and no greater than about 1 mm.

Abstract: An aircraft has first and second fuel sources containing fuels with different characteristics, and one or more gas turbine engines powered by the fuels and each having a staged combustion system having pilot and main fuel injectors and being operable in pilot-only and pilot-and-main ranges of operation. The gas turbine engines each have a fuel delivery regulator arranged to control fuel delivery to the pilot and main fuel injectors. The method includes: obtaining a proposed mission description; obtaining nvPM impact parameters for the gas turbine engines, the impact parameters being associated with each operating condition of the proposed mission; calculating an optimised set of one or more fuel characteristics for each flight condition of the proposed flight defined in the flight description based on the nvPM impact parameters; and determining a fuel allocation based on the optimised set of one or more fuel characteristics.

Abstract: A turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly comprises a plurality of fan blades defining a fan diameter (D), and the heat exchanger module comprises a plurality of heat exchanger elements. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, with the heat exchanger module having an axial length along a central axis between an upstream-most face of the heat exchanger elements and a downstream-most face of the heat exchanger elements. The axial length is in the range of 0.1*D to 5.0*D.

Abstract: A computer-implemented method including: receiving a first data set including a plurality of values for a plurality of features; identifying at least a first feature of the first data set that is non-redundant and at least a second feature of the first data set that is redundant; identifying one or more clusters of features in the plurality of features of the first data set, a first cluster of the one or more clusters including at least the first feature and the second feature; and controlling a display to display the first feature and one or more redundant features from the first cluster, the displayed one or more redundant features from the first cluster including the second feature.

Abstract: A variable speed pumping system, for an aircraft, comprising: a first electrical machine, the first electrical machine comprising an electrical machine rotor shaft mechanically connected to an engine of the aircraft and electrical machine stator coils; an electric pump motor-generator, the electric pump motor-generator comprising an electric pump rotor shaft mechanically connected to the electrical machine rotor shaft via a one way drive arrangement and electric pump stator coils; wherein the electric pump rotor shaft is mechanically connected to a fluid pumping system.

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 gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.

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; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a bypass duct delimited by a bypass duct inner wall and a bypass duct outer wall and located radially outwardly from the engine core and downstream of the fan; and an outlet guide vane assembly, located within the bypass duct and, comprising a plurality of outlet guide vanes distributed circumferentially within the bypass duct, each outlet guide vane extending radially along a span between the bypass duct inner wall and the bypass duct outer wall, wherein a space-chord ratio of at least one outlet guide vane, at 50% of the span length from the bypass duct inner wall, is less than 0.72.

Abstract: A method of magnetising or demagnetising an annular component for a rotary machine, a flux arrangement for performing a magnetising or demagnetising method, and a flux assembly for such a flux arrangement; wherein, the annular component includes an alternating arrangement of radial elements and angular elements for forming a Halbach array. A magnetizer is caused to induce magnetic flux in a primary set of elements of the annular component including a primary radial element and an adjacent primary angular element. A shield element shields a secondary angular element of the annular component from magnetic flux from the magnetizer.

Abstract: A highly efficient gas turbine engine is provided. 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 method of operating a gas turbine engine; the gas turbine engine includes a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The plurality of fuel spray nozzles includes a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes: providing fuel to the one or more fuel-oil heat exchangers. Also provided is a gas turbine engine for an aircraft.

Abstract: Aircraft with battery thermal regulation systems, including: an aircraft air inlet and an aircraft air outlet; a battery pack including battery cells; and a battery thermal regulation system including a first air channel, a surface of which is in thermal communication with the battery cells so that air flowing through the first air channel exchanges heat with the cells through the surface; and an airflow device for driving an airflow. The battery thermal regulation system has a first and a second mode operation. In the first mode, the first air channel is fluidly connected with the aircraft air inlet and the aircraft air outlet whereby, during flight of the aircraft, air can enter the aircraft through the air inlet, flow through the first air channel and exit the aircraft through the air outlet. In the second mode, the airflow device drives an airflow through the first air channel.

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 turbine shroud segment, e.g. for a gas turbine engine, which is coolable by a supply of cooling air, e.g. from a gas turbine engine compressor. The turbine shroud segment has a segment casing that has a radially outer surface and a radially inner surface. The segment casing houses a main plenum, a first layer of impingement chambers, a second layer of impingement chambers, and a plurality of effusion passages. The first layer of impingement chambers fluidly communicates with the main plenum via transfer passages that are formed in the segment casing. The second layer of impingement chambers fluidly communicates with the first layer of impingement chambers via impingement passages that are formed in the segment casing. The effusion passages run between impingement chamber core surface, which is a radially inner surface of an impingement chamber, and the radially inner surface of the segment casing.

Abstract: An aircraft includes a fuselage and a wing, at least one hydrogen fuel tank and a fuel line configured to delivery fuel from the fuel tank to the engine. The fuel tank is installed externally of the fuselage and wing, and comprises an external aeroshell containing a plurality of pressure vessels configured to contain pressurised fuel therein.

Abstract: A gas turbine engine includes a core engine having a compressor, a combustor and a turbine in sequential air flow series. The engine further includes a fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit and a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit. At least first and second compressor bleed offtakes are at different pressure stages of the compressor, each being configured to bleed a portion of air from the compressor. At least first and second compressor bleed offtake valves are configured to control flow through the first and second bleed offtakes respectively. The burner is configured to receive bleed air from the compressor bleed offtakes.