Abstract: Methods for modifying an Engine Control and Monitoring System (ECaMS) for an engine. A method comprises identifying a provisioning deficit in the ECaMS, wherein the ECaMS comprises a first processor node and a second processor node. The first processor node comprises a first acquisition integrated circuit (IC), a first output IC, and a first processor. The second processor node comprises a second acquisition IC, a second output IC, and a second processor. The first acquisition IC is connected directly to the second acquisition IC. The method comprises connecting an expansion unit to the ECaMS. The expansion unit comprises one or more expansion unit ICs, said one or more expansion unit ICs being connected to one or more ICs of the ECaMS.

Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.

Abstract: A gas turbine engine includes an engine core including a compressor, a combustor, and a turbine, the compressor being connected to the turbines through a respective shaft; and a cabin blower system comprising: an electric variator comprising a first electrical machine connected to a first shaft arranged along a first axis, a second electrical machine connected to a second shaft arranged along a second axis, and a power management system; a cabin blower comprising a compressor driven by a third shaft arranged along a third axis, the compressor comprising an air inlet and an air outlet; and a differential gearbox. The gas turbine engine further includes an accessory gearbox arranged within an accessory gearbox casing and adapted to drive the cabin blower system.

Abstract: An apparatus for manufacturing an annular or semi-annular composite component having a circumferentially-extending base and a flange, the apparatus including: a loom for weaving a woven preform of fibre reinforcement material for the composite component; a rotatable mandrel configured to receive and draw the woven preform through the loom for weaving; and a guide disposed between the loom and the rotatable mandrel, configured so that a woven preform drawn by the rotatable mandrel along a guide path under tension engages the guide to transition to a flanged profile at a lip of the guide before being received on the mandrel.

Abstract: An aircraft comprises a machine body which encloses a turbofan gas turbine engine. The turbofan gas turbine engine includes a heat exchanger module, fan assembly, compressor module, turbine module, and exhaust module. The heat exchanger module communicates with the fan assembly by an inlet duct. The heat exchanger module includes first heat transfer elements that transfer heat energy from a first fluid within the transfer elements to an airflow passing over a surface of the transfer elements before entry of the airflow into a fan assembly inlet. The first fluid contained within transfer elements has a temperature, and the airflow passing over the transfer element surface has a temperature. The turbofan gas turbine engine further includes at least one second heat transfer element, with the or each second heat transfer element transfers heat energy from the first fluid to a second fluid.

Abstract: The present disclosure relates to an electric drive apparatus 201, 202 comprising an inverter system 220, an electric motor 230 and a controller 290. The electric motor 230 includes a rotor 234 and a plurality of phase windings 221, 222, 223, 224. The inverter system 220 is configured to provide a respective current waveform 321, 322, 323, 324 to each of the phase windings 221, 222, 223, 224 of the electric motor 230. Each current waveform 321, 322, 323, 324 is defined by a group of electrical characteristics including: a frequency composition, a phase shift relative to a reference waveform 349, and an amplitude. For one or more of the current waveforms 321, 322, 323, 324, the controller 290 is configured to control one or more of the group of electrical characteristics to cause a periodically varying torque to be applied to the rotor 234.

Abstract: A shaft for a gas turbine engine includes a composite tube including a plurality of grooves extending along a longitudinal axis of the shaft. The shaft has a load fuse mechanism that has at least one metallic coupling that has a plurality of splines extending along the longitudinal axis of the shaft. Each of the plurality of splines is received within and engages with a corresponding groove of the composite tube to form a preloaded interference fit between the load fuse mechanism and the composite tube. The at least one metallic coupling includes a first portion defining a first diameter and a second portion extending from the first portion along the longitudinal axis. The second portion defines a second diameter that is greater than the first diameter. The second portion has a smooth annular outer surface devoid of any splines.

Abstract: A continuum arm robot comprising: a tool, a tip section comprising a number of sections a manipulatable robotic section having multiple degrees of freedom, a stiffening section comprising a passive core with an inflatable section surrounding the passive core and a valve for allowing a fluid into the inflatable outer; and a passive section comprising a length of flexible conduit, wherein the core of the passive section and the stiffening section contain the cables for manipulating the tip section and the fluid conduit for supplying the fluid to the inflatable outer.

Abstract: A spline cleaning device for cleaning splines formed within a component of a gas turbine engine. The spline cleaning device comprises: a central support; a central support sleeve that surrounds and is movable with respect to the central support; a scraper having protrusions that substantially correspond to the splines, the scraper being attachable to the central support sleeve, and configured to remove surface contaminants from the surface of the splines; and a collector sump that is attachable to the central support and configured to collect the surface contaminants that have been removed from the surface of the splines.

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: An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 300, 400, 500, 600, 700 configured to introduce a pressurised fluid into a core airflow A at a region downstream of the turbine 19, wherein a rearward movement of the turbine 19 or a shaft 26 in a shaft break event directly actuates the pressure equalisation apparatus 300, 400, 500, 600, 700 to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.

Abstract: An annulus filler of a gas turbine engine includes an outer lid including a leading edge, a trailing edge, a first longitudinal edge, a second longitudinal edge, and an outer radial surface. The outer radial surface includes a first nominal surface portion and a second nominal surface portion. The outer radial surface further includes a protruding surface portion extending radially outwardly from each of the first nominal surface portion and the second nominal surface portion and a recessed surface portion extending radially inwardly from each of the first nominal surface portion and the second nominal surface portion. Each of the protruding surface portion and the recessed surface portion is configured to redirect an air drawn through the gas turbine engine.

Abstract: There is provided an electrical system, comprising: a DC source configured to provide a DC electrical power; a bus electrically coupled to the DC source, the bus comprising a load, a DC-DC converter having an input side and an output side; a switching arrangement electrically coupled to the DC-DC converter, the DC source and the bus; and a control system. The input side of the DC-DC converter is configured to receive a fraction of the DC electrical power provided by the DC source. The fraction is less than unity. The control system is configured to operate the switching arrangement to selectively configure the electrical system, in at least one of a voltage increase mode and a voltage decrease mode.

Abstract: An apparatus for calculating an angular speed of a rotary component includes a processor circuitry configured to: receive an in-phase signal and a quadrature signal from a sensor arrangement, each signal relating to a position of the rotary component; arithmetically sum the in-phase signal and the quadrature signal generated by the sensor arrangement to produce a combined time-domain signal; and calculate an angular speed of the rotary component based on the combined time-domain signal. Methods of calculating an angular speed of a rotary component are also provided.

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: The disclosure relates to a ducted fan aircraft propulsion system and to an aircraft incorporating such a propulsion system. Example embodiments include a ducted fan aircraft propulsion system (300), comprising: a duct (301); a central body portion (302) having first and second ends (303, 304) and extending through the duct (301); a payload portion (305) extending from the first end (303) of the central body portion (302); and a rotor (306) extending across an internal volume (307) of the duct (301) from the central body portion (302), wherein the duct (301) comprises an outwardly flared inlet end (308) such that an inlet air flow passage (309) between the central body portion (302) and an inner surface (310) of the duct (301) has a sectional area that decreases from the inlet end (308) of the duct (301) to the rotor (306).

Abstract: A ducting arrangement for a gas turbine engine and a method of assembling such a ducting arrangement, wherein in the gas turbine engine includes an outer annular duct structure with an outer mount and an inner annular duct structure with an inner mount. The inner annular duct structure is configured to be received within the outer annular duct structure such that the inner mount and the outer mount are angularly aligned. In one aspect, the inner mount and the outer mount are coupled together by a tension tie to attach the outer annular duct structure to the inner annular duct structure, the tension tie being a cable or a wire. In another aspect, the inner mount and the outer mount are configured to be coupled together by a tension tie formed by a cable or a wire to attach the outer annular duct structure to the inner annular duct structure.

Abstract: A power electronics cooling assembly with a housing for power electronics. The housing is mountable to a bulkhead for separating a fire zone from a non-fire zone of a gas turbine engine. In the non-fire zone, one or more ducts extend from the bulkhead to the housing. A fluid supply and return pipes extend through the one or more ducts from the bulkhead to the housing. The fluid supply and return pipes are for carrying an ignitable fluid coolant. A heat exchange structure arranged within the housing is configured to receive the ignitable fluid coolant from the fluid supply pipe, facilitate transfer of heat from the power electronics within the housing to the ignitable fluid coolant, and output the ignitable fluid coolant to the fluid return pipe. A system, gas turbine engine assembly and an aircraft comprising the power electronics cooling assembly are also provided.

Abstract: A continuum arm robot segment that has a proximal end section and a distal end section. The continuum arm robot segment has at least one vertebra between the proximal end section and the distal end section, the vertebra being linked to the proximal end section and to the distal end section by linkages, wherein the proximal end section, the distal end section and the at least one vertebra are configured to have hollow cores to provide a passage within the continuum arm robot segment. The continuum arm robot segment also has at least two tendons that are connected to the distal end section and pass through the at least one vertebra and the proximal end section. The at least two tendons are routed helically so that they undergo substantially at least a 360° rotation between the proximal end section and the distal end section.

Abstract: A method includes forming a moulded component that is axisymmetric about a component axis; segmenting the moulded component into a plurality of segments, each pair of adjacent segments including a pair of surfaces that is formed during segmentation of the moulded component; providing, via computerised numerical control machining, complementary finger joint profiles on the pair of surfaces of the each pair of adjacent segments; providing a plurality of slots on at least one of the pair of surfaces of the each pair of adjacent segments; positioning a plurality of blades partially within the plurality of slots; mating the complementary finger joint profiles provided on the pair of surfaces of the each pair of adjacent segments; and joining the each pair of adjacent segments to each other, such that the plurality of blades is retained within the plurality of slots.