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 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 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: 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 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: 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 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.

Abstract: There is described a fuel injector for a gas turbine engine. The fuel injector comprises an air passageway having an inlet region and an outlet region in fluid communication with the inlet region at an air passageway interface, the inlet region being configured to receive a flow of air from a compressor of the gas turbine engine at an air inlet, the outlet region being configured to receive the flow of air from the inlet region via the air passageway interface and discharge the flow of air to a combustor head of the gas turbine engine. The fuel injector also comprises a fuel passageway having a fuel outlet configured to discharge a flow of fuel into the combustor head. A width of the inlet region in a direction perpendicular to a centreline of the air passageway decreases continuously from the air inlet to the air passageway interface.

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: An example article includes a substrate and a barrier coating on the substrate. The barrier coating includes a matrix including a rare-earth disilicate extending from an inner interface facing the substrate to an outer surface opposite the inner interface. The barrier coating includes a graded volumetric distribution of rare-earth oxide rich (REO-rich) phase regions in the matrix along a direction from the inner interface to the outer surface. The graded volumetric distribution defines a first volumetric density of the REO-rich phase regions at a first region of the matrix adjacent the outer surface. The graded volumetric distribution defines a second volumetric density of the REO-rich phase regions at a second region of the matrix adjacent the inner surface. The second volumetric density is different from the first volumetric density. An example technique includes forming the barrier coating on the substrate of a component.

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 gas turbine engine includes a core engine casing and a bypass duct defined between a nacelle and the core engine casing. The gas turbine engine further includes a plurality of diverter fences pivotally coupled to the core engine casing. Each diverter fence is pivotable relative to the core engine casing about a pivot axis, which is circumferentially and obliquely inclined with respect to a principal rotational axis. Each diverter fence is configured to move between a first position in which an outboard edge is disposed adjacent to a casing outer surface, and a second position in which the outboard edge is radially spaced apart from the casing outer surface, such that each diverter fence radially extends outwards from the casing outer surface into the bypass duct.

Abstract: A method of designing a modified aerofoil shape, the method comprising the steps of: determining a plurality of resonant frequencies of a baseline aerofoil shape; determining that one of the resonant frequencies falls within a predetermined range of frequencies; providing a plurality of aerofoil sections defining the baseline aerofoil shape; and defining a modified aerofoil shape by displacing at least one of the plurality of aerofoil sections relative to its baseline position in the baseline aerofoil shape until one of the plurality of resonant frequencies previously falling within the range of frequencies is shifted outside the range of frequencies.

Abstract: A method of operating a gas turbine engine for an aircraft including: a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor; a fan; turbomachinery bearings; a power gearbox; and a heat management system configured to provide lubrication and cooling to the gearbox and turbomachinery bearings. The method includes operating the heat management system to provide a first amount of heat and a second amount of heat such that a first proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed is in the range of from 0.25 to 0.70; and operating the fan at cruise condition to provide a fan pressure ratio in the range of from 1.35 to 1.43.

Abstract: A gas turbine engine for an aircraft including: a compressor stage; a bleed air line diverted from the compressor stage; and a thermoelectric generator system, the thermoelectric generator system including a thermoelectric generator; a vortex tube and a radiator system, the vortex tube including a flow input fluidically connected to the bleed air line to receive an input of compressed gas from the bleed air line, wherein the vortex tube is configured to separate the compressed gas into a hot flow discharged from a first output of the vortex tube, and a cold flow discharged from a second output of the vortex tube; the radiator system including a first heat exchanger and a second heat exchanger disposed on opposing sides of the thermoelectric generator; and at least one of the hot flow is directed to the first heat exchanger, and the cold flow is directed to the second heat exchanger.

Abstract: Disclosed is a rotary disc balancing simulation mass for balancing a rotary disc configured to support a set of rotary components, wherein the rotary disc balancing simulation mass is configured to be attached to the rotary disc and configured to simulate the mass of two or more of the rotary components. Also disclosed are balancing simulation apparatus, and methods of balancing a rotary disc.

Abstract: A geared gas turbine engine includes a heat management system configured to provide lubrication and cooling to a power gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, an air-lubricant heat exchanger to dissipate a first amount of heat, and a fuel-lubricant heat exchanger to dissipate a second amount of heat wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that at cruise conditions a proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is in the range of from 0.35 to 0.80.

Abstract: An aircraft gas turbine engine includes: an engine core including a compressor, combustor, turbine, and a core shaft connecting the turbine to the compressor; a fan including fan blades and arranged upstream of the engine core; turbomachinery bearings; a power gearbox to drive the fan at a lower rotation speed than the turbine; and a heat management system providing lubrication and cooling to the power gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, 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 first heat sink is air and the second heat sink is fuel.

Abstract: A gas turbine engine for an aircraft includes: an engine core; a fan; turbomachinery bearings; a power gearbox; and a heat management system for providing lubrication and cooling to the power gearbox and turbomachinery bearings, 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 amount of heat and a second amount of heat such that a ratio of a first proportion of heat generated by the gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed at an environment temperature of ISA +40° C. to the first proportion at an environment temperature of ISA ?69° C. is in the range of from 1.5 to 4.5.