Abstract: A gas turbine engine for an aircraft comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a maximum electrical power PEM1 (W), and the gas turbine engine is configured to generate a maximum dry thrust T (N); and a ratio S of: S = ( Maximum ? Electrical ? Power ? Generated = P E ? M ? 1 ) ( Maximum ? Dry ? Thrust = T ) is in a range of between 2.0 and 10.0.

Abstract: An article may include a substrate including a ceramic matrix composite (CMC); a composite coating layer including a first coating material that includes a rare-earth disilicate and a second coating material that includes at least one of a rare-earth monosilicate, a CMAS-resistant material, or a high-temperature dislocating material, where the second coating material forms a substantially continuous phase in the composite coating layer.

Abstract: The disclosure relates to an electrical power system with current fault protection provided by current limiting diodes. Example embodiments include an electrical power system comprising: an electrical machine; an AC:DC power electronics converter connected to receive an input AC supply from the electrical machine and provide an output DC supply across first and second output terminals; a DC bus connected to first and second output terminals of the AC:DC power electronics converter; a load connected across the DC bus; a first circuit breaker switch and a first current limiting diode connected in series between the AC:DC power electronics converter and the DC bus; and a second circuit breaker switch and a second current limiting diode connected in series between the DC bus and the load, wherein the first current limiting diode is configured to limit current to a higher level than the second current limiting diode.

Abstract: Described herein is a flow splitter for a gas turbine engine, the flow splitter comprising a leading edge, wherein the radial and/or axial position of the leading edge varies circumferentially around the flow splitter, such that the leading edge is non-axisymmetric. Also described is a method of manufacture of a flow splitter and an engine comprising a flow splitter.

Abstract: The present disclosure provides a method of manufacturing a turbomachinery component for a gas turbine engine, comprising: receiving an ideal design of the turbomachinery component, the turbomachinery component comprising: an aerofoil portion comprising an aerofoil surface and having an ideal aerofoil wall thickness; and a platform portion comprising a platform surface coterminous with the aerofoil surface, the platform portion having an ideal radial platform thickness along a radial direction; determining an intermediate design of the turbomachinery component comprising an intermediate platform portion having an intermediate platform surface and an intermediate radial platform thickness which is greater than the ideal radial platform thickness; forming, by an investment casting process, an intermediate turbomachinery component based on the intermediate design, wherein the intermediate turbomachinery component comprises at least one detectable reference feature formed by an internal core element during the in

Abstract: A fan blade includes an aerofoil portion including a leading edge extending from a root to a tip and defining a blade span therebetween. A leading edge thickness is defined as a thickness of a cross-section at a given radius at a location along a camber line that is 9% of the total length of the camber line from the leading edge. For cross-sections through the aerofoil portion at radii between 50% and 70% of the blade span from a root radius, the leading edge thickness includes a first maximum value. For cross-sections through the aerofoil portion at radii greater than 70% of the blade span from the root radius, the leading edge thickness includes a second maximum value that is between 105% and 125% of the first maximum value.

Abstract: A fuel delivery system (201) is shown for delivering the hydrogen fuel from a cryogenic storage system to a fuel injection system in a gas turbine engine. The fuel delivery system includes a pump (301), a metering device (302), and a fuel heating system (303,304) for heating the hydrogen fuel to an injection temperature for the fuel injection system.

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: 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: 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: 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: A locking system for a continuum arm robot system, the robot system includes: a continuum arm robot having a manipulatable tip, a passive robot section through which controls for the manipulatable tip, and at least one ferromagnetic collar, and at least one external controllable electromagnetic device which can be activated so that the ferromagnetic section on the continuum arm robot is attracted to the electromagnetic device.

Abstract: An aircraft propulsion system comprises first and second thrust producing gas turbine engines. The system comprises a controller configured to determine a required overall propulsion system thrust level, and determine an engine core power level contribution from each aircraft gas turbine engine such that the overall propulsion system produces a minimum overall noise level and meets the required overall propulsion system thrust level. In meeting the minimum overall noise level, at least the first and second gas turbine engines are operated at different engine core power settings.

Abstract: A rodstock barrel drive having a first and second roller gears, each having a toothed section and a gripping surface section, such that the teeth of the roller gears intermesh, causing rotation of one of roller gear to rotate the other roller gear in the opposite direction, so as to draw a rodstock through a space between the gripping surfaces of the pair of roller gears, a ring gear which engages with one of the pair of roller gears, a drive gear which engages with the ring gear, and one or more idler gears which engage with the ring gear, all within a housing, the drive gear being connected to, and rotated by, a motor, which causes the rodstock to be drawn through the roller gears, wherein all the gears are all positioned within, and oriented parallel to, a common plane.

Abstract: A needle apparatus for making a pinning hole in a composite material, the apparatus having a needle with a longitudinal axis, a collar fixed about the needle, a motor having an axis of rotation, and being fixed to the needle such that the needle can be rotated by the motor, the motor and the collar being situated within a motor housing, the motor housing having a first bush, and a second bush, wherein the collar is positioned between the first bush and second bush so as to prevent the needle from moving away from or towards the motor, but allow the needle to rotate freely around its longitudinal axis, wherein each needle apparatus further comprises a linear actuator connected to the motor such that the motor can be moved along the axis of rotation of the motor by the linear actuator.

Abstract: Disclosed is an aerofoil for a gas turbine engine comprising: a first conduit formed in the aerofoil; a second conduit formed in the aerofoil; and a dividing wall separating the first and second conduits, the dividing wall comprising a transfer port configured to permit fluid flow between the first and second conduits; wherein the dividing wall further comprises a reinforcing boss at least partially encircling the transfer port. Also disclosed is a gas turbine engine comprising the aerofoil and an aircraft comprising the gas turbine engine.

Abstract: A method of determining at least one fuel characteristic of a fuel provided to a gas turbine engine of an aircraft includes making an operational change, the operational change being effected by a controllable component of a propulsion system of which the gas turbine engine forms a part, and being arranged to affect operation of the gas turbine engine, sensing a response to the operational change; and determining the at least one fuel characteristic based on the response to the operational change.

Abstract: A tool for curing and bonding a honeycomb stack with void filler. The tool has a first tool piece with a cavity for receiving layers of unfilled honeycomb material and a second tool piece with a cavity for receiving at least one layer of void-filled honeycomb. The tool also has a compression seal arrangement that creates gas-tight seals between the first tool piece, second tool piece, and a barrier layer positioned between them, with a further seal covering the cavity of the first tool piece to form a gas-tight sealed volume including the first cavity, a pressure plate to cover the cavity of the second tool piece to apply pressure to the void-filled honeycomb, with a further seal to cover the pressure plate to form a gas-tight sealed volume including the second cavity.

Abstract: A woven structure formed by warp and weft tows A, E of fiber reinforcement material includes: a plurality of multi-warp stacks SA, each including a plurality of warp tows A which are in superposition along a longitudinal extent of the warp stack SA, a plurality of multi-weft stacks SE, each including a plurality of weft tows E which are in superposition along the stack SE, wherein one or more multi-warp stacks SA and/or one or more multi-weft stacks SE has an embedded taper structure, including: an embedded tow A1, A2, E1, E2 which has a terminal portion M1, M2 disposed between two locally outermost tows A0, E0 of the respective stack SA SE, the terminal portion M1, M2 terminating at a taper position D1 D2 along the respective path of the stack SA SE; and a method for manufacturing a composite component.