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.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
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.
F02C 6/04 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p.ex. sans sortie de puissance mécanique
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
H01M 8/04111 - Dispositions pour la commande des paramètres des réactifs, p.ex. de la pression ou de la concentration des réactifs gazeux utilisant un assemblage turbine compresseur
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.
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.
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.
F01D 19/00 - Démarrage des "machines" ou machines motrices; Dispositifs de régulation, de commande ou de sécurité en rapport avec les organes de démarrage
F01D 21/00 - Arrêt des "machines" ou machines motrices, p.ex. dispositifs d'urgence; Dispositifs de régulation, de commande ou de sécurité non prévus ailleurs
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.
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.
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.
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.
B32B 37/14 - Procédés ou dispositifs pour la stratification, p.ex. par polymérisation ou par liaison à l'aide d'ultrasons caractérisés par les propriétés des couches
B32B 3/12 - Produits stratifiés caractérisés essentiellement par le fait qu'une des couches comporte des discontinuités ou des rugosités externes ou internes, ou bien qu'une des couches est de forme générale non plane; Produits stratifiés caractérisés essentiellement par des particularismes de forme caractérisés par une couche discontinue, c. à d. soit continue et percée de trous, soit réellement constituée d'éléments individuels caractérisés par une couche d'alvéoles disposées régulièrement, soit formant corps unique dans un tout, soit structurées individuellement ou par assemblage de bandes indépendantes, p.ex. structures en nids d'abeilles
B32B 7/12 - Liaison entre couches utilisant des adhésifs interposés ou des matériaux interposés ayant des propriétés adhésives
B32B 37/00 - Procédés ou dispositifs pour la stratification, p.ex. par polymérisation ou par liaison à l'aide d'ultrasons
B32B 37/10 - Procédés ou dispositifs pour la stratification, p.ex. par polymérisation ou par liaison à l'aide d'ultrasons caractérisés par la technique de pressage, p.ex. faisant usage de l'action directe du vide ou d'un fluide sous pression
B32B 37/12 - Procédés ou dispositifs pour la stratification, p.ex. par polymérisation ou par liaison à l'aide d'ultrasons caractérisés par l'usage d'adhésifs
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.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p.ex. à la température, à la pression, à la vitesse du rotor
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.
D03D 11/00 - Tissus doubles ou à couches multiples non prévus ailleurs
B29C 70/22 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins deux directions formant une structure bidimensionnelle
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.
A gas turbine engine for an aircraft includes: an engine core including a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor, wherein the core shaft has a core shaft maximum take-off speed in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm; a fan; turbomachinery bearings; a power gearbox adapted to drive the fan at a lower rotation speed than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the 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.
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F02C 7/06 - Aménagement des paliers; Lubrification
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
A gas turbine engine for an aircraft includes: an engine core including a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor; a fan including a plurality of fan blades and arranged upstream of the engine core; turbomachinery bearings; a power gearbox adapted to drive the fan at a lower rotation speed than the turbine; and a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings.
A thermoelectric generator system comprising a thermoelectric generator, a vortex tube for receiving a compressed gas from a flow input and separating the compressed gas into a hot flow exiting a first output and a cold flow exiting a second output, a sensor system for determining a first parameter, a second parameter, and a third parameter, the third parameter being indicative of a temperature of a third fluid flow, a radiator system comprising a first and second heat exchangers disposed on opposing sides of the thermoelectric generator; a tube system to separately direct the hot, cold, and third fluid flows towards a switch arrangement configured to be moveable between a first configuration, a second configuration, and a third configuration, with a control unit for controlling the switch arrangement based on the first, second, and third parameter.
H10N 10/13 - Dispositifs thermoélectriques comportant une jonction de matériaux différents, c. à d. dispositifs présentant l'effet Seebeck ou l'effet Peltier fonctionnant exclusivement par les effets Peltier ou Seebeck caractérisés par les moyens d'échange de chaleur à la jonction
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p.ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
H10N 10/17 - Dispositifs thermoélectriques comportant une jonction de matériaux différents, c. à d. dispositifs présentant l'effet Seebeck ou l'effet Peltier fonctionnant exclusivement par les effets Peltier ou Seebeck caractérisés par la structure ou la configuration de la cellule ou du thermocouple constituant le dispositif
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.
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.
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
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.
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.
F02C 7/12 - Refroidissement des ensembles fonctionnels
F02C 7/06 - Aménagement des paliers; Lubrification
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
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.
F02C 7/06 - Aménagement des paliers; Lubrification
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
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.
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.
An apparatus for an aircraft includes an electric machine and a gas turbine. The gas turbine includes a compressor, a turbine unit having a high-pressure turbine and a low-pressure turbine, and a combustion chamber. The compressor is connected via a high-pressure shaft to the high-pressure turbine, and the low-pressure turbine is connected via a low-pressure shaft directly to an output shaft of the electric machine. The high-pressure shaft and the low-pressure shaft are rotatable in different directions. The electric machine is operable as a generator and may be operated as a motor to start the gas turbine. The output shaft is connected via a gear box unit and a free-wheel to the high-pressure shaft. The free-wheel separates the connection between the output shaft of the electric machine and the high-pressure shaft in the presence of a torque flow from the high-pressure shaft in a direction of the output shaft.
A liner for use with a gas turbine engine includes a first liner portion including a first upstream surface and a first downstream surface. The liner further includes a second liner portion spaced apart from the first liner portion. The second liner portion includes a second upstream surface and a second downstream surface. The second upstream surface faces the first downstream surface. Each of the first liner portion and the second liner portion at least circumferentially and radially extends with respect to a central axis. Each of the first liner portion and the second liner portion includes a substrate made of a metallic material and a wear resistant coating disposed on at least a portion of the substrate. The wear resistant coating is made of a polymeric material. The wear resistant coating at least forms the first downstream surface and the second upstream surface.
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
B21D 53/84 - Fabrication d'autres objets particuliers d'autres parties de moteurs, p.ex. bielles de liaison
F23R 3/42 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la disposition ou la forme des tubes à flamme ou des chambres de combustion
A novel configuration for an axial flow gas turbine engine for aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to an engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and combustor volume (in litres) which when divided by fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit an electric motor to be located within fan hub diameter.
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
A platform for a stator vane includes a platform body and a plug. The platform body defines: an inner surface, an outer surface offset from the inner surface, an internal fluid passageway disposed between the outer surface and the inner surface, and a core aperture extending from the outer surface to the internal fluid passageway. The plug 700 is configured to be secured to the platform body. The plug is configured to be partially disposed inside the core aperture and extend outside of the core aperture whilst secured to the platform body for improved heat transfer between a fluid conveyed by the internal fluid passageway and a fluid to which the outer surface is exposed.
A platform for a stator vane includes a platform body and a plug. The platform body defines: an inner surface, an outer surface offset from the inner surface, an internal fluid passageway disposed between the outer surface and the inner surface, and a core aperture extending from the outer surface to the internal fluid passageway. The plug is configured to be secured to the platform body. The plug is configured to occupy no less than 50% of the core aperture by volume whilst secured to the platform body to at least partially eliminate recirculation of a fluid conveyed by the internal fluid passageway within the core aperture.
There is provided a dynamic sealing assembly for a rotary machine, comprising a primary sandwich plate, a secondary sandwich plate and a bristle pack. The primary sandwich plate comprises a plurality of primary vane openings, and the secondary sandwich plate comprises a plurality of secondary vane openings. The bristle pack comprises a plurality of bristles and is disposed between the primary sandwich plate and the secondary sandwich plate. Each of the plurality of primary vane openings overlies and aligns with a respective secondary vane opening to form a vane channel for receiving a vane along a longitudinal axis of the dynamic sealing assembly. The bristle pack is configured to: provide a brush seal between each vane received within the respective vane channels and the dynamic sealing assembly; and allow relative movement between the dynamic sealing assembly and the vane received within each vane channel along the longitudinal axis.
There is disclosed a gas turbine engine comprising: an electrically-powered motor configured to rotate a driven spool of the gas turbine engine; a primary controller configured to conduct a primary function using the motor; and a bow controller configured to selectively control the motor to perform a rotor bow mitigation operation in which the motor drives the driven spool to rotate to mitigate a non-uniform thermal distribution in a rotor of a spool of the gas turbine engine.
F02C 7/268 - Entraînement du rotor pour le démarrage
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gaz; Commande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
30.
SUPER-COOLED ICE IMPACT PROTECTION FOR A GAS TURBINE ENGINE
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.
F02C 7/05 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
A device for generating conversational replies, including a processor with a memory; a speech input module, a user input module; a natural language processing module including one or more encoder-decode modules; the device being configured to: record portions of a conversation through the speech input module, use a speech recognition module to identify words in the conversation, and when one or more words have been recognised: generate one or more responses based on the one or more words using the natural language processing module; selecting a group of the context sensitive responses, prompt the user via the user input module to select a response from the group, output the selected response.
G09B 21/00 - Moyens d'enseignement ou de communication destinés aux aveugles, sourds ou muets
G06F 3/01 - Dispositions d'entrée ou dispositions d'entrée et de sortie combinées pour l'interaction entre l'utilisateur et le calculateur
G10L 13/033 - Procédés d'élaboration de parole synthétique; Synthétiseurs de parole Édition de voix, p.ex. transformation de la voix du synthétiseur
G10L 15/06 - Création de gabarits de référence; Entraînement des systèmes de reconnaissance de la parole, p.ex. adaptation aux caractéristiques de la voix du locuteur
G10L 15/18 - Classement ou recherche de la parole utilisant une modélisation du langage naturel
G10L 15/22 - Procédures utilisées pendant le processus de reconnaissance de la parole, p.ex. dialogue homme-machine
32.
METHOD FOR COATING A TIP OF AN AEROFOIL AND AEROFOIL
A method (400) for coating a tip (106) of an aerofoil (100) is provided. The method (400) includes depositing a layer of nickel-based gamma/gamma prime chemistry (112) on the tip (106) of the aerofoil (100). The method (400) further includes depositing plurality of abrasive particles (114) on the layer of nickel-based gamma/gamma prime chemistry (112) to form a coating matrix (116). The method (400) further includes heating the tip (106) of the aerofoil (100) at a predetermined temperature in order to perform heat treatment of the coating matrix (116) and increase the strength of the coating (110) on the tip (106) of the aerofoil (100).
A combustor assembly for a gas turbine engine, the combustor assembly including: combustor wall defining combustion chamber and including opening periphery, the opening periphery defining first opening; sealing element including annular body, annular body extending around sealing element axis and extending through first opening, sealing element further including flange extending radially outward from annular body, flange slidingly coupling sealing element and combustor wall and forming seal or first partial seal between sealing element and combustor wall, annular body including outer and inner surface, inner surface defining second opening configured to receive fuel nozzle, wherein first total clearance between outer surface and opening periphery in first direction is greater than second total clearance between outer surface and opening periphery in second direction perpendicular to first direction such that sealing element is able to slide relative to combustor wall by greater extent in first direction than in second direction.
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.
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.
B
=
(
First
Heat
Transfer
Fluid
Temperature
=
T
FLUID
)
(
Inlet
Air
Temperature
=
T
AIR
)
A method for manufacturing a fuel spray nozzle for a gas turbine engine includes forming a first section of the fuel spray nozzle by additive layer manufacturing. The first section includes a main chamber and internal passageways. The method includes forming a second section of the fuel spray nozzle by additive layer manufacturing on the first section. The second section includes a metering feature disposed in fluid communication with one of the internal passageways of the first section. The method includes modifying the metering feature of the second section by a first subtractive manufacturing process to obtain a desired diameter of the metering feature and/or a desired surface roughness of the metering feature. The method includes forming a third section of the fuel spray nozzle by additive layer manufacturing on the second section. The third section includes a deflector and a spin chamber.
A gas turbine engine for an aircraft includes: an engine core; fan; turbomachinery bearings; power gearbox; and a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings and including 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 proportion of heat generated by the power gearbox and the turbomachinery and dissipated to air at 85% of a core shaft maximum take-off speed in the range of 0.25 to 0.70, and a second proportion of heat generated by the power gearbox and the turbomachinery and dissipated to air at 65% of the core shaft maximum take-off speed in the range of 0.60 to 1.
Reconfigurable permanent magnet electrical machines and aircraft power and propulsion systems including the electrical machines. An aircraft power and propulsion system includes: a gas turbine engine; DC electrical network; permanent magnet electrical machine including a rotor drivingly coupled to a spool of the engine, and a stator including windings controllably switchable between a star and a delta configuration; an AC-DC power electronics converter, an AC side is connected to terminals of the stator windings and a DC side is connected to the DC electrical network; an additional electrical power source connected to and controllable to supply electrical power to the DC electrical network; and a control system configured to control the switching of the stator windings between the configurations and to control the additional electrical power source to supply electrical power to the DC electrical network during a time interval when the stator is being switched between the configurations.
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.
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p.ex. ensembles fonctionnels de chauffage à turbine à gaz
F25B 27/02 - Machines, installations ou systèmes utilisant des sources d'énergie particulières utilisant la chaleur perdue, p.ex. chaleur dégagée par des moteurs à combustion interne
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. The first flow path is configured to direct a flow of the first heat transfer fluid either to a first heat dissipation portion in which a first proportion QA of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium, or additionally to a second heat dissipation portion in which a second proportion QB of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium, in dependence on a temperature of the first heat transfer fluid entering the first heat exchanger module, a temperature of the first heat dissipation medium, and a temperature of the second heat dissipation medium. The second flow path is configured to direct the flow of the first heat transfer fluid to a second heat dissipation portion in which a second proportion QB of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium, or additionally to the first heat dissipation portion in which a first proportion QA of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium, in dependence on a temperature of the first heat transfer fluid entering the first heat exchanger module, a temperature of the first heat dissipation medium, and a temperature of the second heat dissipation medium.
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p.ex. ensembles fonctionnels de chauffage à turbine à gaz
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, a first heat exchanger, and one or more first ancillary systems. 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 or each first electric machine, the first gas turbine engine, the first heat exchanger, and the or each first ancillary system. Waste heat energy generated by at least one of the first gas turbine engine, the or each first electric machine, and the or each first ancillary system, is transferred to the first heat transfer fluid. The first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.
A thermal management system for an aircraft includes 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 includes 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 includes a first flow path and a second flow path. The first flow path is configured to direct a flow of the heat transfer fluid to a first heat dissipation portion.
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p.ex. ensembles fonctionnels de chauffage à turbine à gaz
A thermal management system for an aircraft includes a thermal bus including one or more first heat sources; a heat sink; a vapour compression system including a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor; and one or more second heat sources. A heat transfer fluid transfers waste heat energy generated by the first heat sources to the heat sink. A second flow of waste heat energy from the second heat source(s) is transferred to a refrigerant. A third flow of heat energy in the refrigerant is transferred to the heat transfer fluid. The compressor includes a supplementary refrigerant reservoir, and the volume of refrigerant in the vapour compression system is increased during operation when a temperature of the heat transfer fluid is at or below a temperature of the second heat source(s).
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p.ex. ensembles fonctionnels de chauffage à turbine à gaz
A thermal management system for an aircraft comprises a first gas turbine engine, a first thermal bus, a first heat exchanger, and a chiller. 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 first heat exchanger, and the chiller. Waste heat energy generated by the first gas turbine engine, is transferred to the first heat transfer fluid. The chiller is configured to lower a temperature of the first heat transfer fluid prior to the first heat transfer fluid being circulated through the gas turbine engine. The first heat is exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p.ex. ensembles fonctionnels de chauffage à turbine à gaz
Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines. A gearbox for an aircraft gas turbine engine includes: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 7/06 - Aménagement des paliers; Lubrification
A method of operating an aircraft including a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine. The method includes: determining at least one fuel characteristic of the fuel arranged to be provided to the gas turbine engine; and proposing or initiating a change to a flight profile of the aircraft based on the at least one fuel characteristic. For example, intended route and/or altitude of the aircraft may be changed based on the one or more fuel characteristics.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p.ex. à la température, à la pression, à la vitesse du rotor
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 37/04 - Disposition de ceux-ci à l'intérieur ou sur les aéronefs
A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines, a first thermal bus, and a first heat exchanger. The first thermal bus includes a first heat transfer fluid 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 first heat transfer fluid. When airspeed of aircraft is less than Mn0.6, the first heat exchanger transfers the waste heat energy from the first heat transfer fluid to a first dissipation medium. When the airspeed of the aircraft is greater than Mn0.6, the first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a second dissipation medium.
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
F02C 7/12 - Refroidissement des ensembles fonctionnels
A thermal management system for an aircraft includes a first thermal bus including one or more first heat sources, a heat sink, a vapour compression system, and one or more second heat sources. The vapour compression system includes a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor. A first heat flow (Q1) of waste heat energy generated by the first heat sources is transferred via the first heat transfer fluid to the heat sink. A second heat flow (Q2) of waste heat energy generated by the second heat source(s) being transferred via the evaporator to a refrigerant. A third heat flow (Q3) of heat energy in the refrigerant is transferred via the condenser to the first heat transfer fluid. The controller is configured to ensure that:
A thermal management system for an aircraft includes a first thermal bus including one or more first heat sources, a heat sink, a vapour compression system, and one or more second heat sources. The vapour compression system includes a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor. A first heat flow (Q1) of waste heat energy generated by the first heat sources is transferred via the first heat transfer fluid to the heat sink. A second heat flow (Q2) of waste heat energy generated by the second heat source(s) being transferred via the evaporator to a refrigerant. A third heat flow (Q3) of heat energy in the refrigerant is transferred via the condenser to the first heat transfer fluid. The controller is configured to ensure that:
1.1*Q2
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p.ex. ensembles fonctionnels de chauffage à turbine à gaz
A thermal management system for an aircraft includes 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. Waste heat energy generated by at least one first gas turbine engine, and first electric machine, transfers to the first heat transfer fluid. The first heat exchanger directs a first proportion of the first heat transfer fluid through a first heat dissipation portion wherein a first proportion of the waste heat energy transfers to a first dissipation medium dependent on the first dissipation medium temperature and mass flow rate. The first heat exchanger directs a second proportion of the first heat transfer fluid through a second heat dissipation portion wherein the second proportion of waste heat energy transfers to a second dissipation medium dependent on the second dissipation medium temperature and mass flow rate.
A method of manufacturing a component includes supporting the component on a mounting structure. The method further includes impacting first peening bodies on the component to increase sub-surface compressive residual stresses of the component. The first peening bodies impact the component at a first intensity of from about 0.45 mmA to about 0.61 mmA. A first coverage of the component by the first peening bodies is from about 120% to about 205%. The method further includes impacting second peening bodies on the component to increase surface compressive residual stresses of the component. The second peening bodies impact the component at a second intensity of from about 0.20 mmA to about 0.30 mmA. A second coverage of the component by the second peening bodies is from about 120% to about 205%. Each of the first peening bodies and the second peening bodies includes a ceramic material.
B24C 1/10 - Méthodes d'utilisation de jet abrasif en vue d'effectuer un travail déterminé; Utilisation d'équipements auxiliaires liés à ces méthodes pour augmenter la compacité des surfaces, p.ex. par grenaillage
Monitoring circuitry connectable to power conversion circuitry enclosed within packaging, the power conversion circuitry including terminals providing external connections to nodes of the power conversion circuitry enclosed within the packaging, the monitoring circuitry including: a comparator including a first input terminals connectable to first and second terminals of the power conversion circuitry and a second input terminal connected to a reference voltage, wherein the first and second terminals provide external connections to first and second nodes connected along respective current paths each connected to switching circuitry of the power conversion circuitry enclosed within the packaging, wherein the comparator detects a switch state of the switching circuitry based on a voltage generated at the first input terminal and the reference voltage; processing circuitry connected to the comparator and configured to monitor the power conversion circuitry based on the detected switch state and a signal of the power conversion circuitry.
G01R 31/26 - Test de dispositifs individuels à semi-conducteurs
H02M 1/00 - APPAREILS POUR LA TRANSFORMATION DE COURANT ALTERNATIF EN COURANT ALTERNATIF, DE COURANT ALTERNATIF EN COURANT CONTINU OU VICE VERSA OU DE COURANT CONTINU EN COURANT CONTINU ET EMPLOYÉS AVEC LES RÉSEAUX DE DISTRIBUTION D'ÉNERGIE OU DES SYSTÈMES D'ALI; TRANSFORMATION D'UNE PUISSANCE D'ENTRÉE EN COURANT CONTINU OU COURANT ALTERNATIF EN UNE PUISSANCE DE SORTIE DE CHOC; LEUR COMMANDE OU RÉGULATION - Détails d'appareils pour transformation
H02M 1/08 - Circuits spécialement adaptés à la production d'une tension de commande pour les dispositifs à semi-conducteurs incorporés dans des convertisseurs statiques
H02M 5/293 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant alternatif, p.ex. pour changement de la tension, pour changement de la fréquence, pour changement du nombre de phases sans transformation intermédiaire en courant continu par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs
52.
SYSTEM AND METHOD FOR FINISHING A SURFACE OF A WORKPIECE
A system for finishing a surface of a workpiece. The system includes a laser unit configured to emit a laser radiation. The system further includes an attenuator disposed within the laser unit and configured to adjust optical parameters of the laser radiation. The system further includes a cavitation chamber storing a liquid medium. The workpiece is mounted within the cavitation chamber and is in contact with the liquid medium. The system further includes at least one lens configured to focus at least a portion of the laser radiation and transmit at least one laser beam towards the workpiece.
B23K 26/352 - Travail par rayon laser, p.ex. soudage, découpage ou perçage pour le traitement de surface
B23K 26/06 - Mise en forme du faisceau laser, p.ex. à l’aide de masques ou de foyers multiples
B23K 26/0622 - Mise en forme du faisceau laser, p.ex. à l’aide de masques ou de foyers multiples par commande directe du faisceau laser par impulsions de mise en forme
B23K 26/57 - Travail par transmission du faisceau laser à travers ou dans la pièce à travailler le faisceau laser entrant dans une face de la pièce à travailler d’où il est transmis à travers le matériau de la pièce à travailler pour opérer sur une face différente de la pièce à travailler, p.ex. pour effectuer un enlèvement de matière, pour rac
A titanium alloy comprising Al 4.78 to 6.44 wt. %; V 3.65 to 5.15 wt. %; Mo 1.32 to 3.58 wt. %; Cr 0.75 to 2.28 wt. %; Fe 0.00 to 0.42 wt. %; C 0.00 wt. % to 0.10 wt. %; S 0.00 wt. % to 0.10 wt. %; N up to 500 ppm; O up to 2000 ppm and H up to 150 ppm; the balance being Ti and incidental elements and unavoidable impurities. Such a titanium alloy is useful for manufacturing gas turbine engine components including turbine blades and stators.
A method of mitigating contrails produced by an aircraft having a set of gas turbine engines, comprises the steps of (i) for each engine in a first subset of the engines, reducing the operating efficiency of the engine to produce a reduction in thrust provided by that engine and (ii) for each engine in a second subset, increasing the fuel flow to the engine to increase the thrust provided by that engine, the set of at least two gas turbine engines consisting of the first and second subsets. The method provides for contrail mitigation action by means of engine operating efficiency reduction to be directed to a first subset of engines for which contrail mitigation per unit engine operating efficiency reduction is greatest, the resulting reduction in thrust provided by such engines being at least partially compensated by increasing fuel flow to engines of the second subset.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p.ex. à la température, à la pression, à la vitesse du rotor
F02C 9/42 - Commande de l'alimentation en combustible spécialement adaptée à la commande simultanée d'au moins deux ensembles fonctionnels
F02C 9/44 - Commande de l'alimentation en combustible sensible à la vitesse de l'aéronef, p.ex. commande du nombre de Mach, optimisation de la consommation en combustible
An electrical machine with a rotor coupled to a drive shaft of an engine; a DC electrical network; a power electronics converter connected, on a DC-side, to the DC electrical network, and, on an AC-side, to the electrical machine, the power electronics converter comprising a plurality of transistors and associated diodes connected in anti-parallel with the transistors; and a controller configured to control switching of the transistors so that the converter either inverts DC power to AC power or rectifies AC power to DC power. The controller is further configured, responsive to a determination that there is a fault in the DC electrical network, to operate in a fault mode in which the controller controls the switching of the transistors so that a fault current is shared between the transistors and the associated diodes.
H02M 7/219 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continu sans possibilité de réversibilité par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs dans une configuration en pont
H02J 3/38 - Dispositions pour l’alimentation en parallèle d’un seul réseau, par plusieurs générateurs, convertisseurs ou transformateurs
H02M 7/217 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continu sans possibilité de réversibilité par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs
An inlet assembly includes a web extending between a first end and a second end along a first axis. The web includes first and second major surfaces. The inlet assembly includes a load transfer flange at least partially disposed around the web and configured to be fixedly coupled to a gas turbine engine. The inlet assembly includes first vanes spaced apart from each other at least along the first axis. Each first vane extends between the first major surface of the web and the load transfer flange. The inlet assembly includes second vanes spaced apart from each other at least along the first axis. Each second vane extends between the second major surface of the web and the load transfer flange.
B64D 33/02 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des entrées d'air de combustion
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p.ex. l'air
57.
ELECTRIC MACHINE HAVING ROTOR HUB WITH SHAPED PASSAGES FOR COOLING
An electric machine adapted for use in a gas turbine engine includes a shaft extending along a central axis, a magnetic rotor drum, and a non-magnetic rotor hub rotatably coupled with the shaft and the magnetic rotor drum. The magnetic rotor drum includes a rotor and a plurality of magnets arranged circumferentially about the central axis.
An electric machine for use in an aircraft propulsion system includes a shaft, a magnetic rotor drum, and a rotor hub. The shaft extends along an axis and is configured to rotate about the axis. The magnetic rotor drum includes a rotor that extends circumferentially around the axis and a plurality of magnets that are arranged circumferentially around the rotor. The rotor hub extends radially between and interconnects the shaft and the rotor. The rotor hub is configured to move air along the electric machine to cool components thereof in response to rotation of the shaft about the axis.
H02K 9/06 - Dispositions de refroidissement ou de ventilation par l'air ambiant s'écoulant à travers la machine comportant des moyens pour établir la circulation d'un agent de refroidissement avec des ventilateurs ou des dispositifs d'entraînement mûs par l'arbre de la machine
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur, l'électricité ou l'énergie de ressorts
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
H02K 1/30 - Moyens de montage ou de fixation des parties magnétiques tournantes sur ou aux structures constituant le rotor utilisant des pièces intermédiaires, p.ex. des croisillons
H02K 5/20 - Enveloppes ou enceintes caractérisées par leur configuration, leur forme ou leur construction avec des canaux ou des conduits pour la circulation d'un agent de refroidissement
H02K 7/00 - Dispositions pour la mise en œuvre d'énergie mécanique associées structurellement aux machines dynamo-électriques, p.ex. association structurelle avec des moteurs mécaniques d'entraînement ou des machines dynamo-électriques auxiliaires
An electrical-discharge machining tool comprising a ground electrode, a moveable and activatable assistive electrode and a cutting electrode, a first conduit for supplying a dielectric and a second conduit for flushing dielectric material, wherein when used on a dielectric layer on a substrate the assistive electrode is activated and positioned substantially between the ground electrode and the cutting electrode.
A sensing system for a continuum arm robot, the sensing system comprising at least one camera ring system mounted to the continuum robot, the camera ring system having at least two forward-facing cameras that face along the axis of the robot towards the tip.
There is disclosed a method of manufacturing a component. The method comprises melting and fusing powder in layers from a build plate to form the component and a disposable support structure, heat treating the component; and removing the support structure. Each layer of the support structure and the component are formed such that there is at least a 10-micron gap between them, to ensure that the component and the support structure are not fused together.
B22F 10/47 - Structures destinées à soutenir des pièces ou des articles pendant la fabrication et retirées par la suite caractérisées par des caractéristiques structurelles
B22F 10/62 - Traitement de pièces ou d'articles après leur formation par des moyens chimiques
A fuel precursor is provided for producing hydrogen fuel by reacting aluminium and water. The fuel precursor has aluminium particles suspended in a hydrophobic liquid such that when the fuel precursor is introduced to water, the suspended aluminium particles migrate to the water and react therewith to produce hydrogen. The suspended aluminium particles are non-spherical, angular particles. The surfaces of the suspended aluminium particles have substantially no oxide layer thereon
C01B 3/08 - Production d'hydrogène ou de mélanges gazeux contenant de l'hydrogène par réaction de composés inorganiques comportant un hydrogène lié électropositivement, p.ex. de l'eau, des acides, des bases, de l'ammoniac, avec des agents réducteurs inorganiques avec des métaux
Disclosed is a method of manufacturing a component, the method comprising: determining a predicted direction and extent of distortion of a portion of the component; designing a support structure based on the predicted direction and extent of distortion of the portion of the component; additive manufacturing, by melting and fusing powder from a build plate, the component and the support structure. The support structure comprises, a hook structure comprising an elongate arm and a hook protruding from the elongate arm. The hook is deposited to abut a distortion surface of the component to restrain distortion of the component. The distortion surface is the furthest in the direction of distortion of the component. The elongate arm is deposited to extend from the hook in a direction having a vector component opposing the direction of distortion such that shrinkage of the elongate arm provides a force to minimise distortion of the component.
B22F 10/47 - Structures destinées à soutenir des pièces ou des articles pendant la fabrication et retirées par la suite caractérisées par des caractéristiques structurelles
B22F 10/85 - Acquisition ou traitement des données pour la commande ou la régulation de procédés de fabrication additive
Rolls-Royce North American Technologies, Inc. (USA)
Rolls-Royce Plc (Royaume‑Uni)
Inventeur(s)
Schenk, Peter
Tanner, Mark Angelo
Abrégé
An example system includes an electrical machine electrically configured to generate electrical energy used by one or more components of a gas-turbine engine; an energy storage system; and a controller electrically connected to the energy storage system and configured to receive electrical energy from the energy storage system, wherein, in response to the gas-turbine engine being shut off, the controller is configured to cause the electrical machine to rotate a rotor of the gas-turbine engine using the energy received from the energy storage system.
H02K 11/33 - Circuits d’entraînement, p.ex. circuits électroniques de puissance
H02J 7/00 - Circuits pour la charge ou la dépolarisation des batteries ou pour alimenter des charges par des batteries
H02K 7/00 - Dispositions pour la mise en œuvre d'énergie mécanique associées structurellement aux machines dynamo-électriques, p.ex. association structurelle avec des moteurs mécaniques d'entraînement ou des machines dynamo-électriques auxiliaires
H02K 7/18 - Association structurelle de génératrices électriques à des moteurs mécaniques d'entraînement, p.ex. à des turbines
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur, l'électricité ou l'énergie de ressorts
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 31/00 - Commande des groupes moteurs; Leur disposition
B60L 50/52 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant de la puissance de propulsion fournie par des batteries ou des piles à combustible caractérisée par des moteurs à courant continu
An electric machine for use in an aircraft propulsion system includes a shaft, a magnetic rotor drum, and a rotor hub. The shaft extends along an axis and is configured to rotate about the axis. The magnetic rotor drum includes a rotor that extends circumferentially around the axis and a plurality of magnets that are arranged circumferentially around the rotor. The rotor hub extends radially between and interconnects the shaft and the rotor. The rotor hub is configured to move air along the electric machine to cool components thereof in response to rotation of the shaft about the axis.
H02K 9/06 - Dispositions de refroidissement ou de ventilation par l'air ambiant s'écoulant à travers la machine comportant des moyens pour établir la circulation d'un agent de refroidissement avec des ventilateurs ou des dispositifs d'entraînement mûs par l'arbre de la machine
H02K 5/20 - Enveloppes ou enceintes caractérisées par leur configuration, leur forme ou leur construction avec des canaux ou des conduits pour la circulation d'un agent de refroidissement
A combustor assembly for gas turbine engine, combustor assembly including: combustor wall including interior surface and combustor wall opening, wherein interior surface partially defines combustion chamber, wherein combustor wall opening extends between combustion chamber and its exterior; sealing element disposed partially within combustor wall opening and includes air inlet, air outlet and air passageway, wherein air outlet exits into combustion chamber and delivers flow of air received from exterior of combustion chamber via air inlet and air passageway to combustion chamber; and fuel nozzle coupled to sealing element and configured to deliver fuel into combustion chamber, wherein combustor wall further includes first bearing surface and sealing element further includes second bearing surface, wherein first bearing surface is concave and forms part of first spherical surface such that first and second bearing surfaces are configured to move relative to each other about central point of first spherical surface.
A method for manufacturing a composite component, including: weaving a multi-layer woven preform from warp and weft tows of fibre-reinforcement material, the woven preform includes: multi-layer weave including: plurality of weft tow layers; plurality of laterally-adjacent stacks extending along the longitudinal direction, primary portion having longitudinal extent along woven preform, the primary portion having one or more edge regions each defining respective lateral side of primary portion; wherein for the or each edge region: the multi-layer weave defines at least the edge region; plurality of stacks in the edge region are binding stacks in which one or more warp tows are interlaced to bind a respective plurality of weft tow layers; weave property differs between binding stacks in the edge region to reduce a thickness of the edge region towards respective lateral side. Also disclosed herein is a woven structure, formed by warp and weft tows of fibre reinforcement material.
D03D 11/00 - Tissus doubles ou à couches multiples non prévus ailleurs
B29C 70/24 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins trois directions formant une structure tridimensionnelle
68.
COMPOSITE COMPONENT, METHOD OF MANUFACTURING A PREFORM FOR THE COMPONENT
A woven composite component for an aerospace structure or machine, comprising: a compound member extending from a junction with a feeder member, a noodle element extending in a weft direction through the junction. The feeder member comprises first and second feeder portions either side of the junction, each feeder portion comprising warp tows extending towards the junction. There is a compound set of warp tows extending from the first and second feeder portions, each turning at the junction to define warp tows for the compound member. There is a crossing set of warp tows belonging to the compound set, the crossing set comprising warp tows from the first feeder portion and warp tows from the second portion which cross each other at the junction to pass around the noodle element. There is also disclosed a method of manufacturing a preform for a woven composite component.
B29C 70/22 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins deux directions formant une structure bidimensionnelle
Disclosed is an apparatus for the additive manufacture of a component comprising: a bed for supporting the component; a powder depositor; an energy beam source for emitting a beam of energy towards a beam target location; and control means for adjusting the beam target location; wherein the apparatus is configured to provide rotation of the beam target location and/or the powder depositor relative to the bed, the relative rotation being about a rotational axis. Also disclosed are an additive manufacture apparatus comprising a substantially annular bed, a method for additive manufacture, a gas turbine engine, and an aircraft.
B22F 12/00 - Appareils ou dispositifs spécialement adaptés à la fabrication additive; Moyens auxiliaires pour la fabrication additive; Combinaisons d’appareils ou de dispositifs pour la fabrication additive avec d’autres appareils ou dispositifs de traitement ou de fabrication
B33Y 30/00 - Appareils pour la fabrication additive; Leurs parties constitutives ou accessoires à cet effet
B33Y 50/02 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
B29C 64/241 - Moyens d’entraînement pour mouvement rotatif
B29C 64/153 - Procédés de fabrication additive n’utilisant que des matériaux solides utilisant des couches de poudre avec jonction sélective, p.ex. par frittage ou fusion laser sélectif
B29C 64/393 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
B29C 64/268 - Agencements pour irradiation par faisceaux d’électrons [FE]
B29C 64/232 - Moyens d’entraînement pour un mouvement le long de l'axe orthogonal au plan d’une couche
B22F 10/28 - Fusion sur lit de poudre, p.ex. fusion sélective par laser [FSL] ou fusion par faisceau d’électrons [EBM]
B22F 10/36 - Commande ou régulation des opérations des paramètres du faisceau d’énergie
B22F 12/44 - Moyens de rayonnement caractérisés par la configuration des moyens de rayonnement
B28B 1/00 - Fabrication d'objets façonnés à partir du matériau
Disclosed is an apparatus for the additive manufacture of a component comprising a substantially annular bed for supporting the component, a powder depositor, an energy beam source for emitting a beam of energy towards a beam target location, and control means for adjusting the beam target location. The apparatus is configured to provide rotation of the beam target location and/or the powder depositor relative to the bed, the relative rotation being about a rotational axis. Also disclosed are a method for additive manufacture, a gas turbine engine, and an aircraft.
B29C 64/153 - Procédés de fabrication additive n’utilisant que des matériaux solides utilisant des couches de poudre avec jonction sélective, p.ex. par frittage ou fusion laser sélectif
B22F 12/00 - Appareils ou dispositifs spécialement adaptés à la fabrication additive; Moyens auxiliaires pour la fabrication additive; Combinaisons d’appareils ou de dispositifs pour la fabrication additive avec d’autres appareils ou dispositifs de traitement ou de fabrication
A manufacturing method includes: in a loom, weaving a woven reinforcing fibre fabric including a plurality of reinforcing fibre tows and polymeric material; and heating the woven reinforcing fibre fabric as it exits the loom to cause the polymeric material to melt and/or cure.
D03D 15/283 - Tissus caractérisés par la matière, la structure ou les propriétés des fibres, des filaments, des filés, des fils ou des autres éléments utilisés en chaîne ou en trame caractérisés par la matière des fibres ou des filaments formant les filés ou les fils à base de polymères synthétiques, p.ex. fibres polyamides ou fibres polyesters
B29C 70/22 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins deux directions formant une structure bidimensionnelle
D03D 49/04 - Contrôle de la tension de la chaîne ou du tissu
D03D 15/587 - Tissus caractérisés par la matière, la structure ou les propriétés des fibres, des filaments, des filés, des fils ou des autres éléments utilisés en chaîne ou en trame caractérisés par les propriétés des filés ou des fils fusibles
B29C 70/24 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins trois directions formant une structure tridimensionnelle
A method of training a neural network to generate conversational replies, the method including: providing a first dataset of stored phrases linked to form a plurality of conversational sequences; training the neural network to generate responses to input phrases using the first dataset; and using the trained neural network to generate a list of conversational replies in response to conversational inputs.
G10L 15/16 - Classement ou recherche de la parole utilisant des réseaux neuronaux artificiels
G10L 15/06 - Création de gabarits de référence; Entraînement des systèmes de reconnaissance de la parole, p.ex. adaptation aux caractéristiques de la voix du locuteur
A manufacturing method comprises: providing a woven fabric comprising a plurality of reinforcing fibre tows and a plurality of thermoplastic polymer yarns woven together; and moulding the woven fabric in a heated mould to form a preform for a composite component.
B29C 70/22 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins deux directions formant une structure bidimensionnelle
74.
WOVEN STRUCTURE, METHOD AND APPARATUS FOR A FLANGED COMPOSITE COMPONENT
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.
B29C 53/56 - Enroulement et assemblage, p.ex. enroulement en spirale
B29C 53/80 - Façonnage par cintrage, pliage, torsion, redressage ou aplatissage; Appareils à cet effet - Eléments constitutifs, détails ou accessoires; Opérations auxiliaires
An oil system for an aircraft propulsion system. The aircraft propulsion system comprises a motor configured to drive a propulsor via a reduction gearbox. The oil system comprises a first feed line configured to supply cooling oil to the reduction gearbox, a second feed line configured to supply cooling oil to a further oil cooled component, and an air-cooled oil cooler. The air-cooled oil cooler is provided upstream of the first feed line to provide cooled oil to the reduction gearbox; and the second feed line is configured to bypass the air-cooled oil cooler to cool the further oil cooled component.
A novel configuration for axial flow gas turbine engine for aircraft has an engine core having a core length and including first turbine, axial compressor, and drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of the engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to an engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and a combustor volume (in litres) which when divided by the fan tip radius (in mm) is 0.015 to 0.083. The fan may be a multistage fan configured to permit an electric motor to be located within the fan hub diameter.
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
77.
High-Power Epicyclic Gearbox and Operation Thereof
An engine for an aircraft includes 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. The gearbox is an epicyclic gearbox and comprises a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted. The radial bending stiffness of the planet carrier is equal to or greater than 1.20×109 N/m, and/or the tilt stiffness of the planet carrier is greater than or equal to 6.00×108 Nm/rad. A method of operation of such an engine is also disclosed.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F16H 57/08 - TRANSMISSIONS - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
A novel configuration for an axial flow gas turbine engine for an aircraft has an engine core having a core length and including a first turbine, an axial compressor, and a drive connecting the first turbine to the axial compressor, the engine core further including a second turbine, and a fan shaft connecting the second turbine to a fan located upstream of engine core, the fan including a plurality of fan blades extending from a fan hub, the fan having a tip radius from 90 mm to 225 mm and wherein the ratio of fan tip radius to engine length is 0.15 to 0.25. The engine may have 3, 4 or 5 compressor stages and a combustor volume (in litres) which when divided by the fan tip radius (in mm) is 0.015 to 0.083. The fan may be multistage fan configured to permit electric motor to be located within fan hub diameter.
F02C 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
79.
Combustor casing component for a gas turbine engine
A combustor casing component for a gas turbine engine which is manufactured as a single-piece. The combustor casing component has a combustor outer casing portion, an outlet guide vane outer case portion, a pre-diffuser portion, a plurality of outlet guide vanes, and an outlet guide vane inner case portion.
The disclosure relates to an electrical power system for connecting an electrical machine to first and second DC networks operating at different voltages. In an embodiment, an electrical power system comprises: an electrical machine having first, second, third and fourth windings; first, second, third and fourth AC:DC power electronics converters each connected to receive an input AC supply from the respective first, second, third and fourth windings, each AC:DC power electronics converter having first and second DC output terminals connecting the AC:DC power electronics converters to first, second and third DC supply output terminals, wherein the first and second windings are arranged to provide an AC supply to the respective first and second AC:DC power electronics converters in quadrature to each other and the third and fourth windings are arranged to provide an AC supply to the respective third and fourth AC:DC power electronics converters in quadrature to each other.
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur, l'électricité ou l'énergie de ressorts
H02M 7/219 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continu sans possibilité de réversibilité par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs dans une configuration en pont
H02P 9/02 - Dispositions pour la commande de génératrices électriques de façon à obtenir les caractéristiques désirées à la sortie - Détails
H02J 1/00 - Circuits pour réseaux principaux ou de distribution, à courant continu
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
The disclosure relates to an electrical power system for connecting an electrical machine to first and second DC networks operating at different voltages. Example embodiments include an electrical power system comprising: an electrical machine having first and second pluralities of windings; a first AC:DC power electronics converter connected to the first plurality of windings; a second AC:DC power electronics converter connected to the second plurality of windings; a switching controller configured to provide switching signals to the first and second AC:DC power electronics converters to provide a first DC supply at a DC side of the first AC:DC power electronics converter and a second DC supply at a DC side of the second AC:DC power electronics converter.
H02P 25/22 - Enroulements multiples; Enroulements pour plus de trois phases
H02P 27/06 - Dispositions ou procédés pour la commande de moteurs à courant alternatif caractérisés par le type de tension d'alimentation utilisant une tension d’alimentation à fréquence variable, p.ex. tension d’alimentation d’onduleurs ou de convertisseurs utilisant des convertisseurs de courant continu en courant alternatif ou des onduleurs
H02K 3/28 - Schémas d'enroulements ou de connexions entre enroulements
82.
CIRCUITRY AND METHODS FOR DETECTING AN ELECTRICAL FAULT IN POWER CONVERSION CIRCUITRY
A power conversion circuitry configured to detect an electrical fault, and methods of detecting an electrical fault in power conversion circuitry, the power conversion circuitry including: switching circuitry configured to control a current flowing along a current path between first and second nodes of the power conversion circuitry; a current sensor connected to the current path, the current sensor configured to measure a rate of change of the current flowing along the current path; and control circuitry connected to the current sensor and the switching circuitry, the control circuitry configured to: detect an electrical fault if a rate of change of the current measured by the current sensor exceeds a fault threshold, and configure the switching circuitry to adjust current flow in the power conversion circuitry when an electrical fault is detected.
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
H02M 3/155 - Transformation d'une puissance d'entrée en courant continu en une puissance de sortie en courant continu sans transformation intermédiaire en courant alternatif par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs
83.
CIRCUITRY AND METHOD FOR TRANSITIONING BETWEEN MODES OF OPERATION DURING AN ELECTRICAL FAULT
A power conversion circuitry configured to transition between modes of operation during an electrical fault, including: switching circuitry configured to control currents flowing along a plurality of current paths between respective pairs of nodes, wherein the paths include first and second transistor-diode pairs; and control circuitry connected to the switching circuitry configured to: configure the switching circuitry to transition, when an electrical fault is detected, to a first mode of operation by configuring the transistors to switch to a non-conducting state so currents flow through the diodes, determine when current begins to flow through a diode, and configure the switching circuitry to transition from the first mode to a fault mode of operation by configuring the transistors to sequentially switch to a conducting state so currents flow through the configured transistors. Also, a method of controlling power conversion circuitry to transition between modes of operation during an electrical fault.
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
H02M 7/219 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continu sans possibilité de réversibilité par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs dans une configuration en pont
H02M 1/00 - APPAREILS POUR LA TRANSFORMATION DE COURANT ALTERNATIF EN COURANT ALTERNATIF, DE COURANT ALTERNATIF EN COURANT CONTINU OU VICE VERSA OU DE COURANT CONTINU EN COURANT CONTINU ET EMPLOYÉS AVEC LES RÉSEAUX DE DISTRIBUTION D'ÉNERGIE OU DES SYSTÈMES D'ALI; TRANSFORMATION D'UNE PUISSANCE D'ENTRÉE EN COURANT CONTINU OU COURANT ALTERNATIF EN UNE PUISSANCE DE SORTIE DE CHOC; LEUR COMMANDE OU RÉGULATION - Détails d'appareils pour transformation
A gas turbine engine system includes a hydrogen-burning gas turbine engine and a fuel system including a fuel line arranged to receive gaseous hydrogen at an input thereof and provide the gaseous hydrogen to combustion apparatus of the hydrogen-burning gas turbine engine and a vent line including a vent valve and having a first end coupled to the fuel line and a second end disposed remotely from the hydrogen-burning gas turbine engine. A controller is arranged to switch the vent valve from a closed state to an open state upon detection of an engine shaft-break or similar condition, thus providing rapid evacuation of gaseous hydrogen from the fuel line and hence rapid shut-down of the engine. The engine may be shut down more rapidly than is possible by means of a shut-off valve within the fuel line.
F02C 9/46 - Commande de secours de l'alimentation en combustible
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
A gas turbine engine includes a core engine casing, a bypass duct and a core engine duct. The gas turbine engine further includes a plurality of flaps pivotally coupled to the core engine casing and arranged circumferentially around a principal rotational axis. Each flap extends from a first casing end and is configured to pivotally rotate relative to core engine casing about a pivot axis between a first position and a second position. In the first position, each flap is disposed in a circumferential direction and is radially disposed between a plurality of outlet guide vanes and a plurality of stator vanes. In the second position, each flap is inclined to the first position and extends at least partially into the bypass duct and the core engine duct.
F02K 1/70 - Inversion du flux de la soufflante utilisant des volets inverseurs de poussée ou des portes montées sur le carter de la soufflante
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux commande du rapport des débits des différents flux
A laser ablation tool comprising a laser source which produces a beam heaving a beam path and an ablation head, the ablation head comprising a housing which on each side of the beam path is separated into a first and second portion, and at least a first Risley prism and a second Risley prism connected to the first and second portions of the housing the first and second Risley prisms being connected to a rotation mechanism, so that the two Risley prisms can be moved relative to each other and the housing so as to deflect the beam.
B23K 26/361 - Enlèvement de matière pour l'ébarbage ou l'ébavurage mécanique
B23K 26/06 - Mise en forme du faisceau laser, p.ex. à l’aide de masques ou de foyers multiples
B23K 26/082 - Systèmes de balayage, c. à d. des dispositifs comportant un mouvement relatif entre le faisceau laser et la tête du laser
B23K 26/0622 - Mise en forme du faisceau laser, p.ex. à l’aide de masques ou de foyers multiples par commande directe du faisceau laser par impulsions de mise en forme
Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines, including a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F16C 17/03 - Paliers à contact lisse pour mouvement de rotation exclusivement pour charges radiales uniquement avec segments supportés obliquement, p.ex. paliers Michell
F16H 1/28 - Transmissions à engrenages pour transmettre un mouvement rotatif avec engrenages à mouvement orbital
F16H 57/02 - Boîtes de vitesses; Montage de la transmission à l'intérieur
F16H 57/08 - TRANSMISSIONS - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
A laser ablation tool including, a pulsed laser source, a prism that rotates in a plane perpendicular to light from the laser source, and a mirror that is moveable in at least one direction and which is set at an angle relative to the light from the laser source.
G02B 7/18 - Montures, moyens de réglage ou raccords étanches à la lumière pour éléments optiques pour miroirs
G02B 26/08 - Dispositifs ou dispositions optiques pour la commande de la lumière utilisant des éléments optiques mobiles ou déformables pour commander la direction de la lumière
A repair system for in-situ repair of a machine. The repair system includes a laser unit configured to emit a laser beam, a galvanometer optically coupled to the laser unit and configured to adjust the laser beam received from the laser unit, a source of a flexible metallic mesh, a delivery unit configured to receive the flexible metallic mesh from the source of the flexible metallic mesh, and a controller communicably coupled to the laser unit, the galvanometer, and the delivery unit. The controller is configured to control the delivery unit to deposit the flexible metallic mesh on a surface of the machine. The controller is further configured to control the laser unit to emit the laser beam and to control the galvanometer to adjust the laser beam in order to weld the flexible metallic mesh to the surface of the machine, thereby forming a repair patch.
A method for facilitating maintenance management of a propulsion system, such as an engine, for a vehicle is disclosed. The method comprises obtaining, for each of a plurality of propulsion systems, records of maintenance events experienced by the propulsion system, and records of traversals of transport routes by the propulsion system during a period of propulsion system operation. The method further comprises using a Machine Learning model to classify the recorded maintenance events into a plurality of maintenance categories. The method then comprises identifying, from the classified recorded maintenance events and the records of traversals of transport routes, a correlation between a given maintenance category and the transport routes traversed by propulsion systems during operational periods preceding maintenance events classified into the maintenance category. The correlation may be used in maintenance and/or route planning for propulsion systems.
A computer implemented method of modelling the degradation of at least a component within an abrasive environment, the method comprising: obtaining a 3D scan of a surface of a component, processing the surface images of the component, determining the surface properties, inputting the data into a trained prediction model, predicting the performance of the component, and simulating the surface evolution over time or use of the component.
A bearing carrier support for connecting a rotary shaft of a gas turbine engine to a static portion of the gas turbine engine. The bearing carrier support has: a bearing portion for supporting the rotary shaft; a flange portion for attaching the bearing carrier support to the static portion of the gas turbine engine; and a support portion that connects the flange portion and the bearing portion. The support portion has a plurality of circumferentially distributed struts that are axially aligned with respect to the rotary shaft. The number of struts is at least 500 and the diameter of each strut is no more than 3 mm.
A hydrogen-fuelled aircraft power system incorporates a gas turbine engine and a fuel cell, in which air supply and cooling of the fuel cell is integrated with the gas turbine engine to improve overall efficiency of the power system. The system may be part of a turbofan, turboprop or electric propulsion system for an aircraft.
B64D 37/30 - Circuits de carburant pour carburants particuliers
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p.ex. l'air
H01M 8/04111 - Dispositions pour la commande des paramètres des réactifs, p.ex. de la pression ou de la concentration des réactifs gazeux utilisant un assemblage turbine compresseur
H01M 8/04007 - Dispositions auxiliaires, p.ex. pour la commande de la pression ou pour la circulation des fluides relatives à l’échange de chaleur
H01M 8/04082 - Dispositions pour la commande des paramètres des réactifs, p.ex. de la pression ou de la concentration
94.
PLANETARY GEARBOX DEVICE AND GAS TURBINE ENGINE WITH A PLANETARY GEARBOX DEVICE
The invention relates to a planetary gearbox device with a static ring gear, a rotatable planet carrier, a plurality of planet gears, each planet gear being connected to a bearing device, wherein at least one of the bearing devices is coupled to at least one lubricant scooping device and/or at least one lubricant reservoir device for collecting lubricant, in particular oil coming from the at least one bearing device, and at least one lubricant channel is for allowing a flow of the collected lubricant towards a lubrication location and/or a lubricant supply for the planetary gearbox device.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 7/06 - Aménagement des paliers; Lubrification
B64D 35/02 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le type de groupe moteur
A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 3/113 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur avec plusieurs rotors raccordés par transmission de puissance aves des transmissions de puissance variables entre les rotors
F01D 15/12 - Combinaisons avec des transmissions mécaniques
96.
METHOD OF FORMING PROTECTIVE COATING AND COATED ARTICLE COMPRISING PROTECTIVE COATING
A method of forming a protective coating includes providing a substrate including a major surface. The method further includes providing a top coat layer adjacent to the major surface of the substrate. The top coat layer includes a top coat surface distal to the substrate. The method further includes forming a plurality of slots in the top coat layer. The method further includes at least partially heating the top coat surface above a threshold temperature, such that a first portion of the top coat layer extending from the top coat surface partially melts to form a glazed layer, and a second portion of the top coat layer disposed between the glazed layer and the major surface of the substrate is not melted and includes at least a portion of each of the plurality of slots.
A hydrogen-fuelled gas turbine engine has a fuel input path from a fuel input of the engine to a combustor, the engine further comprising a heat-exchanger located in the fuel input path and a fuel turbine located in the fuel input path between the heat-exchanger and the combustor. The heat-exchanger is arranged to receive waste heat from the engine core in order to heat hydrogen fuel introduced at the fuel input prior to input thereof to the combustor. The engine further comprises an air compressor arranged to be driven by the fuel turbine. The engine provides for most of the waste heat of combustion to be recovered and used without the need for heavy electrical or mechanical apparatus and is therefore particularly advantageous in aeronautical applications.
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
B64D 15/04 - Dégivrage ou antigivre des surfaces externes des aéronefs par gaz chaud ou liquide amené par conduit par amenée de gaz chaud
B64D 37/30 - Circuits de carburant pour carburants particuliers
B64D 37/34 - Conditionnement du carburant, p.ex. réchauffage
B64D 13/08 - Aménagements ou adaptations des appareils de conditionnement d'air pour équipages d'aéronefs, passagers ou pour emplacement réservé au fret l'air étant climatisé l'air étant réchauffé ou refroidi
98.
DETERMINING SOLUTIONS TO A NUMBER OF LINEAR MATRIX EQUATIONS
A method, performed on at least one computing device, of determining solutions to a number of linear matrix equations satisfying A{right arrow over (x)}={right arrow over (b)}, where A is a n×n matrix, {right arrow over (x)} is a column vector with n entries, and {right arrow over (b)} is a column vector with n entries, is disclosed. The method comprises determining a linear combination of unitary matrices that is equivalent to the matrix A; based on the linear combination of unitary matrices, determining a column vector {right arrow over (x)} that satisfies the linear matrix equation; forming an updated matrix A based on the obtained column vector {right arrow over (x)}; forming an updated column vector {right arrow over (b)} based on the obtained column vector {right arrow over (x)}; updating the coefficients of the linear combination of unitary matrices based on the updated column vector {right arrow over (x)}; and based on the updated linear combination of unitary matrices, determining an updated column vector {right arrow over (x)} that satisfies the updated linear matrix equation.
A gas delivery system intended for use in delivering hydrogen gas from a cryogenic liquid hydrogen storage tank is provided. The gas delivery system has: a cryogenic liquid storage tank; an evaporator; a pressurisable gas reservoir; a valve sub-system including: a multi-outlet valve arrangement having a first valve inlet, a first valve outlet and a second valve outlet; and a multi-inlet valve arrangement having a second valve inlet, a third valve inlet and a third valve outlet; a delivery line connected to the third valve outlet; an evaporator feed line connecting the storage tank and the evaporator inlet; an evaporator dispensing line connecting the evaporator and the first valve inlet; a reservoir feed line connecting the first valve outlet and the reservoir; a reservoir dispensing line connecting the reservoir and the second valve inlet; and a reservoir bypass line connecting the second valve outlet and the third valve inlet. The multi-outlet valve arrangement has: a first state in which the first valve outlet is open, and the second valve outlet is closed; a second state in which the first valve outlet and the second valve outlet are closed; and a third state in which the first valve outlet is closed, and the second valve outlet is open. The multi-inlet valve arrangement has: a fourth state in which the second valve inlet is open, and the third valve inlet is closed; and a fifth state in which the second valve inlet is closed, and the third valve inlet is open. The gas delivery system further comprises a computer-based controller configured to control the valve sub-system to provide: a first operating condition of the gas delivery system in which the multi-outlet valve arrangement is in the first state and the multi-inlet valve arrangement is in the fourth state, the gas reservoir delivers gas to the delivery line via the reservoir dispensing line, and the gas reservoir is concurrently pressurised by gas fed from the evaporator via the reservoir feed line, the reservoir bypass line being closed; a second operating condition of the gas delivery system in which the multi-outlet valve arrangement is in the second state and the multi-inlet valve arrangement is in the fourth state, the gas reservoir delivers gas to the delivery line via the reservoir dispensing line, the reservoir feed line and the reservoir bypass line being closed; and a third operating condition of the gas delivery system in which the multi-outlet valve arrangement is in the third state and the multi-inlet valve arrangement in the fifth state, the evaporator delivers gas to the delivery line via the reservoir bypass line, the reservoir feed line and the reservoir dispensing line being closed.
F17C 7/00 - Procédés ou appareils pour vider les gaz liquéfiés, solidifiés ou comprimés contenus dans des récipients sous pression, non couverts par une autre sous-classe
Disclosed is an aircraft accessory drive system comprising: a drive shaft configured to provide a drive input; a critical accessory directly coupled to the drive shaft so as to be directly driven by the drive shaft; a parasitic drive system configured to transmit drive input from the drive shaft to one or more less critical accessories, wherein the parasitic drive system comprises a torque limiter configured to decouple the parasitic drive system from the drive shaft and the critical accessory in response to a jam affecting the parasitic drive system or one or more of the less critical accessories. Also disclosed are a gas turbine engine comprising an aircraft accessory drive system, and an aircraft comprising an aircraft accessory drive system.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
B64D 35/00 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur, l'électricité ou l'énergie de ressorts
F02C 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
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