Abstract: A gas turbine engine for an aircraft configured with an engine core that has a turbine, a compressor, and a core shaft connecting the turbine to the compressor. A fan located upstream of the engine core, that has a plurality of fan blades. A gearbox arranged to receive an input from the core shaft and to output to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox being an epicyclic gearbox having a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The gearbox having an overall gear mesh stiffness, and wherein the overall gear mesh stiffness of the gearbox is greater than or equal to 1.05×109 N/m and less than or equal to 8.0×109 N/m.

Abstract: A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.

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 mounting assembly includes an engine mount link, an engine mount block, a fusible component, and a catcher component. The engine mount block is connected to a pylon and the engine mount link. The fusible component and the catcher component have respective first and second shear strengths. The catcher component is disposed within a catcher aperture. When an applied shear stress on the fusible component is less than the first shear strength, the engine mount link is stationary with respect to the engine mount block. When the applied shear stress on the fusible component exceeds the first shear strength, the fusible component mechanically fails and causes the catcher component to move relative to the catcher aperture further causing a corresponding movement of the engine mount link.

Abstract: There is provided a current limiting diode 200 comprising a gate 206, a source 202, a drain 204 electrically connected to the source by an n-channel or p-channel 220 and a thermoelectric generator 208; wherein the source and the gate are electrically connected by a fill structure 210, and wherein the thermoelectric generator is configured to: generate a thermoelectric voltage by absorbing heat from the n-channel or p-channel and rejecting heat to a heat sink in a current limiting condition; and apply the thermoelectric voltage between the gate and the source.

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: A tool for immobilising a first shaft of a gas turbine engine in relation to a second shaft of a gas turbine engine, the tool comprising a cylindrical body having a first end and a second end opposite the first end, the cylindrical body having one or more interior protrusions on the inner surface of the cylindrical body for engaging with the first shaft, and the outer surface of the cylindrical body having one or more exterior protrusions for engaging with the second shaft, the tool being configured such that at least part of the tool can be inserted within and held fixed relative to a section of the second shaft, and a section of the first shaft can be inserted within and held fixed relative to the interior of the cylindrical body.

Abstract: A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

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: The present disclosure relates to a hydraulic fuse comprising: a fuse body defining a plenum, the plenum comprising: an inlet configured to couple to a first fluid location; and an outlet configured to couple to a second fluid location.

Abstract: An aircraft propulsion system fuel system comprises a fuel line configured to receive liquid hydrogen fuel from a fuel tank, a vaporizer configured to vaporize liquid hydrogen fuel from the fuel line to generate a supercritical or gaseous fuel, a main fuel pump configured to receive and to pump the gaseous or supercritical fuel from the vaporizer during operation of the propulsion system, and a heater provided downstream in fuel flow to the main fuel pump, and configured to raise the temperature of the gaseous or supercritical fuel to a propulsion system delivery temperature.

Abstract: There is provided a method of starting a gas turbine engine, comprising: selecting a baseline starting procedure of the gas turbine engine based on environmental data; receiving condition information for the gas turbine engine and/or for a starting system of the gas turbine engine, the condition information comprising at least one of: health information for the gas turbine engine and/or the starting system; maintenance information for the gas turbine engine and/or the starting system; and operation information relating to an expected future usage of the gas turbine engine. The method further comprises determining a predicted degradation profile for the gas turbine engine and/or the starting system based on the condition information; determining a starting procedure of the gas turbine engine by adapting the baseline starting procedure based on the predicted degradation profile; and controlling the gas turbine engine and/or the starting system according to the determined starting procedure.

Abstract: A gas turbine engine fuel system includes a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and a power recovery device configured to extract power from exhaust gasses of the burner downstream of the heat-exchanger.

Abstract: A combined gas turbine engine and hydrogen fuel cell system includes a hydrogen fuelled gas turbine engine, a cryogenic liquid hydrogen fuel tank, a first 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, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, a second fuel offtake arranged to divert a portion of hydrogen fuel from the main fuel conduit downstream of the heat exchanger, and a hydrogen fuel cell configured and arranged to produce electric power using hydrogen fuel diverted from the second fuel offtake.

Abstract: A thermal management system for cooling at least one electronic component, the thermal management system including: a heat exchanger connectable to at least one electronic component, the heat exchanger configured to dissipate heat generated by the at least one electronic component; a plurality of cooling channels each connected to a respective pump and the heat exchanger, wherein: each cooling channel is configured to carry fluid between the heat exchanger and the respective pump, and each pump is configured to pump fluid along the respective cooling channel such that fluid carried by each cooling channel flows through the heat exchanger.

Abstract: An aircraft including a battery pack having a plurality of battery cells arranged to vent into a venting region of the battery pack. A path is defined between the venting region and an exterior of the aircraft, whereby thermal products vented by the cells can exit the aircraft via the path. In some embodiments, the venting region is or is in fluid communication with an air channel through which air flows from an aircraft air inlet to an aircraft air outlet. In other embodiments, the path joins the venting region and an opening in an external surface of the aircraft.

Abstract: There is provided a method for scanning of an object in a scanning apparatus. The method comprises placing the object into a jacket which surrounds the object or a part of the object to be scanned. The volume of the jacket surrounding the object is solid or filled with a filling material. The jacket is disposed on a support of the scanning apparatus, so that the jacket is positioned between an imaging beam emitting element and an imaging beam receiving element oppositely disposed to either side of the support. The support is rotatable relative to the emitting and receiving elements about an axis of rotation to allow creation of an image from projections each taken at a different relative angle of rotation. The method further comprises operating the scanning apparatus at the multiple relative angles of rotation to produce an image of the object.

Abstract: A data logging and transmission device is provided for an electric aircraft having a charging port for on-ground charging of an energy storage system of an electric propulsion system of the aircraft. The device includes: a control unit, data storage unit, data transmitter, and sensor configured to detect connection of charging port of the electric aircraft to ground-side charger and then issue a trigger signal to the control unit. The control unit is configured wherein a state of flight of the electric aircraft, the control unit collects inflight data from the electric aircraft and saves the inflight data to the data storage unit. The control unit is also configured such that in a state of on-ground charging of the electric aircraft, the control unit, on receipt of the trigger signal, commands the data transmitter to transmit the inflight data saved on the data storage unit to a remote receiver.

Abstract: There is provided a method for scanning of objects in a scanning apparatus. The method comprises disposing the objects on a support of the scanning apparatus, so that the objects are positioned between an imaging beam emitting element and an imaging beam receiving element oppositely disposed to either side of the support. The support is rotatable relative to the emitting and receiving elements about an axis of rotation to allow creation of an image from projections each taken at a different relative angle of rotation. The objects are positioned adjacent each other on the support in a configuration that reduces the variation in material thickness penetrated at the multiple relative angles of rotation. The method further comprises operating the scanning apparatus at the multiple relative angles of rotation to produce an image of the objects.

Abstract: A data logging device for an electric aircraft having charging port for on-ground charging of energy storage system of electric propulsion system of aircraft. Device includes an aircraft-side data storage unit, data interface, trigger signal interface, and control unit. Control unit is configured to detect connection of charging port to ground-side charger based on trigger signal received by control unit from ground-side charger via trigger signal interface. Control unit is configured wherein a state of flight of electric aircraft, control unit collects inflight data from electric aircraft via data interface and saves inflight data to aircraft-side data storage unit. Control unit is also configured wherein a state of on-ground charging of electric aircraft, control unit, on receipt of trigger signal, sends inflight data saved on aircraft-side data storage unit via data interface to charger for forwarding to at least one of a data transmitter and a ground-side data storage unit.