A butterfly valve for a conduit defining a passage for a flow of a fluid therethrough in a flow direction. The butterfly valve includes a shaft rotatably mounted to the conduit and defining a longitudinal axis along its length. The butterfly valve includes a valve body coupled to the shaft, such that the valve body is rotatable along with the shaft about the longitudinal axis between a closed position and a fully open position. The valve body includes a first major surface, a second major surface opposite to the first major surface, a perimeter surface, a central plane, a first lobe, a second lobe, and a third lobe.
F16K 1/22 - Soupapes ou clapets, c.-à-d. dispositifs obturateurs dont l'élément de fermeture possède au moins une composante du mouvement d'ouverture ou de fermeture perpendiculaire à la surface d'obturation à éléments de fermeture articulés à pivot comportant disque ou volet pivotant dont l'axe de rotation traverse le corps de soupape, p. ex. régulateurs à papillon
A combined cycle power generation and storage system is shown. A liquid air energy storage system uses excess power to liquefy air and stores it in a liquid state. A combustion engine produces power by combustion of a carbon-based fuel along with an exhaust stream containing carbon dioxide. A heat recovery system exchanges heat from the exhaust stream to air from the liquid air storage system, and thereby produce a cooled exhaust stream and heated air. An air expansion machine recovers power by expansion of heated air from the heat recovery system. A separation system separates carbon dioxide from ambient air prior to liquefaction during operation of the liquid air energy storage system, and separates carbon dioxide from the cooled exhaust stream during operation of the combustion engine prior to emission of the cooled exhaust stream to atmosphere.
F02B 63/04 - Adaptations des moteurs pour entraîner des pompes, des outils tenus à la main ou des génératrices électriquesCombinaisons portatives de moteurs avec des dispositifs entraînés par des moteurs pour génératrices électriques
F01N 3/02 - Silencieux ou dispositifs d'échappement comportant des moyens pour purifier, rendre inoffensifs ou traiter les gaz d'échappement pour refroidir ou pour enlever les constituants solides des gaz d'échappement
F01N 3/08 - Silencieux ou dispositifs d'échappement comportant des moyens pour purifier, rendre inoffensifs ou traiter les gaz d'échappement pour rendre les gaz d'échappement inoffensifs
A power electronics converter connectable, on a DC-side, to a DC electrical network and either, on an AC-side, to an electrical machine coupled to a drive shaft of an engine or propulsor, or, on a second DC-side, to a battery pack. The power electronics converter includes: a power conversion unit including a plurality of semiconductor switching elements and a DC-link; and a gate driver unit, configured to control the semiconductor switching elements so that the power conversion unit: inverts DC power received from the DC electrical network to AC power and provides the AC power to the electrical machine, rectifies AC power received from the electrical machine to DC power and provides the DC power to the DC electrical network, or performs DC-DC conversion between the DC-electrical network and the battery pack; wherein the gate driver unit includes an equipment health monitoring component.
G01R 31/00 - Dispositions pour tester les propriétés électriquesDispositions pour la localisation des pannes électriquesDispositions pour tests électriques caractérisées par ce qui est testé, non prévues ailleurs
H02M 1/00 - Détails d'appareils pour transformation
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
A gas turbine engine configured with an engine core. A fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier on which the planet gears are mounted, the planet carrier having an effective linear torsional stiffness and the gearbox having a gear mesh stiffness between the planet gears and the ring gear. Additionally, the product of the effective linear torsional stiffness of the planet carrier and the gear mesh stiffness between the planet gears and the ring gear is greater than or equal to 5.0×1018 N2m−2.
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 1/28 - Transmissions à engrenages pour transmettre un mouvement rotatif avec engrenages à mouvement orbital
F16H 57/08 - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 2.5×10−2 Nkg−1 m−1 mm−2. A tilt stiffness to moment of inertia ratio of:
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 2.5×10−2 Nkg−1 m−1 mm−2. A tilt stiffness to moment of inertia ratio of:
the
tilt
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
Gas turbine engine for aircraft includes: an engine core including a turbine, compressor, and core shaft connecting the turbine to the compressor; a fan located upstream of the core; a gearbox; and a gearbox support arranged to at least partially support the gearbox. The fan has a mass in a range of 150 kg to 1200 kg. A moment of inertia of the fan is greater than or equal to 7.40×107 kgmm2. A radial bending stiffness to moment of inertia ratio of:
the
radial
bending
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 2.5×10−2 Nkg−1 m−1 mm−2. A tilt stiffness to moment of inertia ratio of:
the
tilt
stiffness
of
at
least
one
of
the
fan
shaft
at
the
output
of
the
gearbox
and
the
gearbox
support
the
moment
of
inertia
of
the
fan
may be greater than or equal to 4.0×10−4 Nmrad−1 kg−1 mm−2.
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 3/107 - 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
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
F16H 57/02 - Boîtes de vitessesMontage de la transmission à l'intérieur
F16H 57/025 - Support des boîtes de vitesses, p. ex. bras de couple, ou attachement à d'autres dispositifs
F16H 57/08 - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
There is disclosed an energy storage system for an electric aircraft, the energy storage system comprising: at least one battery pack configured to be disposed onboard the aircraft, and a thermal management system. The thermal management system comprises a first circulation loop configured to be disposed onboard the aircraft and configured to contain a first working fluid, the first circulation loop including: a variable speed pump configured to pump the first working fluid around the first circulation loop, a battery heat exchanger configured to provide a thermal interface between the at least one battery pack and the first working fluid, and a controller configured to control operation of the variable speed pump to intermittently pump the first working fluid to distribute heat around the thermal management system.
B60L 58/26 - Procédés ou agencements de circuits pour surveiller ou commander des batteries ou des piles à combustible, spécialement adaptés pour des véhicules électriques pour la surveillance et la commande des batteries pour la commande de la température des batteries par refroidissement
An exhaust nozzle for a gas turbine engine comprises an exhaust duct and a first flap. The exhaust duct is configured to receive an exhaust flow of gas from a combustor of the gas turbine engine. The first flap is rotatably coupled to the exhaust duct for rotation about a first axis of rotation. Further, the first flap at least in part defines an exhaust gas passageway configured to convey the exhaust flow of gas to an exterior of the gas turbine engine. Additionally, the first flap comprises a first pin. The exhaust nozzle comprises a first moveable cam having a first moveable slot configured to slidably receive the first pin. The exhaust nozzle is configured such that movement of the first moveable cam causes the first flap to be moved about the first axis of rotation between a first inner position and a first outer position.
A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.
F02C 9/24 - Commande du niveau de pression dans les cycles fermés
F01D 19/00 - Démarrage des "machines" ou machines motricesDispositifs de régulation, de commande ou de sécurité en rapport avec les organes de démarrage
F02C 7/268 - Entraînement du rotor pour le démarrage
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 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:
bypass
exhaust
nozzle
pressure
ratio
core
exhaust
nozzle
pressure
ratio
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of:
bypass
exhaust
nozzle
pressure
ratio
core
exhaust
nozzle
pressure
ratio
is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.
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
F01D 17/16 - Organes de commande terminaux disposés sur des parties du stator faisant varier l'aire effective de la section transversale des injecteurs ou tuyères de guidage en obturant les injecteurs
F02K 1/00 - Ensembles fonctionnels caractérisés par la forme ou la disposition de la tubulure de jet ou de la tuyèreTubulures de jet ou tuyères particulières à cet effet
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
10.
METHOD AND FURNACE FOR TREATING AN IRON-COBALT COMPONENT
A method of treating an iron-cobalt component is disclosed. The method comprises: heat treating the component using a temperature of at least 700° C.; applying a static magnetic field of at least 3000 A/m to the component during the heat treatment. Also disclosed is a furnace for treating the iron-cobalt component, and a method of forming a stator for a transverse flux electric machine.
C21D 1/04 - Procédés ou dispositifs généraux pour le traitement thermique, p. ex. recuit, durcissement, trempe ou revenu avec application simultanée d'ondes supersoniques, de champs magnétiques ou électriques
C21D 6/00 - Traitement thermique des alliages ferreux
C21D 9/00 - Traitement thermique, p. ex. recuit, durcissement, trempe ou revenu, adapté à des objets particuliersFours à cet effet
C21D 11/00 - Commande ou régulation du processus lors de traitements thermiques
H02K 15/02 - Procédés ou appareils spécialement adaptés à la fabrication, l'assemblage, l'entretien ou la réparation des machines dynamo-électriques des corps statoriques ou rotoriques
AC-DC converter circuits which may include: first and second rectifier circuits each having a plurality of input connections for connection to respective first and second sets of windings of a generator, each input connection connected between a pair of series-connected rectifier diodes connected between first and second output terminals, an output capacitor connected between the first and second output terminals; a first output diode connected between the second output terminals of the first and second rectifier circuits; and a first output switch connected between the second output terminal of the first rectifier circuit and the first output terminal of the second rectifier circuit.
H02M 1/00 - Détails d'appareils pour transformation
H02M 7/06 - 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 sans électrode de commande ou des dispositifs à semi-conducteurs sans éléctrode de commande
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
12.
METHOD OF MANUFACTURING A COMPONENT OF A GAS TURBINE ENGINE
There is provided a method of manufacturing a component of a gas turbine engine, comprising: providing a precursor component having at least one internal cooling passage configured to receive a flow of cooling air therethrough; estimating a predicted temperature profile for the component based on one or more operating parameters of the gas turbine engine, the predicted temperature profile indicating a predicted operating temperature of at least one gas-washed surface of the component; determining a thermal barrier coating (TBC) configuration for the component based on the predicted temperature profile, comprising setting a TBC thickness to be below a threshold thickness in a region of the at least one gas-washed surface of the component based on the predicted operating temperature of the at least one gas-washed surface of the component exceeding a threshold temperature; and applying a TBC to the precursor component according to the TBC configuration.
A fluid control valve is disclosed, comprising: a valve body defining an outer wall for a curved flow path through the fluid control valve; and a valve element defining an opposing inner wall and rotatable relative to the valve body about a pivot point through between a closed position a range of open positions. A separation between the first inner wall and the outer wall along a restrictor portion of the flow path varies between open positions. The first inner wall and the outer wall are cooperatively defined so that for at least some open positions the separation between the first inner wall and the outer wall is constant along the respective restrictor portion of the flow path.
F16K 1/24 - Soupapes ou clapets, c.-à-d. dispositifs obturateurs dont l'élément de fermeture possède au moins une composante du mouvement d'ouverture ou de fermeture perpendiculaire à la surface d'obturation dont le corps de soupape est initialement soulevé de son siège à l'ouverture et tourne ensuite autour d'un axe parallèle au siège
A measurement apparatus for measuring a flow rate of a powder includes a casing, a nozzle configured to dispense the powder, a fixture plate, a weighing scale, and a powder collector. The fixture plate includes a plurality of pinhole members. Each pinhole member includes a tip, a cylindrical hole extending from the tip, and a discharge passage. The cylindrical hole of each pinhole member has a diameter. The diameters of the cylindrical holes of the plurality of pinhole members are different from each other. The nozzle is configured to dispense the powder selectively into the cylindrical hole, and the powder collector is configured to receive at least a portion of the powder from the discharge passage.
A fuel system for a gas turbine engine. The fuel system comprises 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 and at least first and second heat exchangers. The first heat exchanger is configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and the second heat exchanger is provided upstream in hydrogen flow of the first heat exchanger and is configured to transfer heat from a further heat exchange fluid to hydrogen fuel in the main fuel conduit. In an embodiment, the further heat exchange fluid is compressor bleed air bled from a core compressor of the gas turbine engine.
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
A system for inserting a flexible robotic arm into a workspace, the system having: a feed mechanism having a housing containing a drive mechanism, the drive mechanism gripping the flexible robotic arm using at least one drive wheel which is coupled to a drive motor, the feed mechanism is provided with at least one force sensor; and a haptic controller coupled to the feed mechanism, the haptic controller having input controller for instructing the drive motor within the feed mechanism to activate in response to the signal from the input controller, and wherein the input controller is provided with haptic feedback to provide an operator an indication of the state of operation of the feed mechanism based on signals output from signals provided by the force sensors.
A system for inspecting a complex component, the system comprising: a flexible inspection device having a driver for controlling the motion of the flexible inspection device by manipulating the relative positions of a plurality of joints within the device, the flexible device having a sensor on its distal end; a feed mechanism for controlling the insertion or retraction of the flexible device into or from the complex component; a component driver for driving the movement of an aspect of the complex component from at least a first position to a second position; and a computer running a computer program, the computer interfacing with the flexible inspection device, the feed mechanism, and the component driver, so as to link the operation of the flexible inspection device, the feed mechanism and the component driver through a single program.
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
G01M 15/02 - Détails ou accessoires pour appareils de test
A feed mechanism for inserting and/or retracting a flexible robotic arm from an area, the feed mechanism including a housing, a passageway extending about a central longitudinal axis through the length thereof, a drive portion and rotational portion, the drive portion not being fixedly connected within the housing and having at least a pair of drive wheels coupled to a drive motor, the drive wheels are connected to a mounting frame with the mounting frame being connected to pivotable linkages with at least one of the linkages being connected to a spring, such that the expansion or contraction of the spring allows the drive wheels to move relative to the longitudinal axis, and the rotational portion being coupled to a motor, the rotational motor having a gear system that connects to the drive portion and causes a rotation of the drive portion about the central longitudinal axis within the housing.
A method of operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor to combust fuel, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; a bearing supporting the fan shaft; an oil loop system supplying oil to the bearing; and a heat exchange system including: an air-oil heat exchanger through which oil in the oil loop system flows; and a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow to transfer heat between the oil and the fuel; and a bypass pipe to allow a proportion of the oil to flow past one of the air-oil and the fuel-oil heat exchanger; and a bypass valve to allow the proportion of the oil sent through the bypass pipe to be varied.
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 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
F01D 15/12 - Combinaisons avec des transmissions mécaniques
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
An annulus filler for mounting to a rotor disc of a gas turbine engine includes a coupling portion and an outer lid. The outer lid includes a leading edge, a trailing edge, a pair of longitudinal edges, and an outer radial surface. Each longitudinal edge includes a first edge portion and a second edge portion. At least one longitudinal edge includes a recessed portion that extends at least axially between the first edge portion and the second edge portion, such that the recessed portion connects the first edge portion to the second edge portion. The recessed portion further extends circumferentially inwards from each of the first edge portion and the second edge portion towards the other longitudinal edge.
A gas turbine engine component that has a web provided with an array of cooling holes distributed with respect to a first direction, wherein the cooling holes have a cross-sectional shape that varies along the first direction. The component may be configured so that the first direction corresponds to a loading distribution of the component which increases along the first direction from a low loading position to a high loading position.
A propulsion device comprises: a casing structure to surround a fan of the propulsion device; a core structure to support a core of the propulsion device; a thrust reverser unit comprising two thrust reverser halves, each thrust reverser half being pivotable about a respective hinge line between an open position for access to the core structure, and a closed position. Upper and lower support members extend from the casing structure at diametrically opposing sides of a centreline axis. For each thrust reverser half there is at least one locating arrangement comprising an upper locating arrangement with cooperating portions configured to engage each other as the thrust reverser half moves towards the closed position in a closing operation; and/or a lower locating arrangement with cooperation portions configured to engage each other as the thrust reverser half moves towards the closed position in the closing operation.
A gas turbine engine includes a staged combustion system having pilot fuel injectors and main fuel injectors. A fuel delivery regulator controls delivery of fuel to the pilot and main fuel injectors, receives fuel from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic. In a transition range of operation between the pilot-only and the pilot-and-main ranges of operation, fuel is delivered to both the pilot and main fuel injectors at a transition staging ratio different from the pilot-and-main staging ratio. The fuel delivery regulator delivers fuel to one or both the pilot and main fuel injectors during the transition range of operation having a different fuel characteristic from fuel delivered to one or both the pilot and main fuel injectors during at least part of the pilot-and-main range of operation.
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
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
F23R 3/34 - Alimentation de différentes zones de combustion
25.
METHOD OF MANUFACTURING A COMPOSITE ARTICLE PRECURSOR
A method of manufacturing a composite article precursor, e.g. from which to make an aerofoil. The method of manufacturing involves: laying down a first grouping of at least one layer of composite material upon a substantially upwardly-facing first surface of a preform of the composite article precursor; disposing a first deformable support over the first grouping of at least one layer of composite material; securing the first deformable support to the preform via a first securing fixing; rotating the preform about an axis such the first surface is no longer substantially upwardly-facing and a second surface of the preform is substantially upwardly-facing; laying down a second grouping of at least one layer of composite material upon the second surface of the preform; and vacuum debulking an assembly comprising the first grouping composite material, the second grouping of composite material, the first deformable support and the preform.
B29C 70/54 - Parties constitutives, détails ou accessoiresOpérations auxiliaires
B29C 70/44 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.-à-d. d'objets distincts utilisant une pression isostatique, p. ex. moulage par différence de pression, avec un sac à vide, dans un autoclave ou avec un caoutchouc expansible
B29L 31/08 - Pales pour rotors, stators, ventilateurs, turbines ou dispositifs analogues, p. ex. hélices
A decoking cleaning system for cleaning at least one component in-situ within a combustion engine system, the decoking system including at least one ultrasonic transducer with a means of connecting the at least one ultrasonic transducer to a component to be cleaned within the combustion engine system, the at least one ultrasonic transducer being connected to a power and voltage supply.
F02C 7/30 - Prévention de la corrosion dans les espaces balayés par les gaz
B08B 3/00 - Nettoyage par des procédés impliquant l'utilisation ou la présence d'un liquide ou de vapeur d'eau
B08B 7/02 - Nettoyage par des procédés non prévus dans une seule autre sous-classe ou un seul groupe de la présente sous-classe par distorsion, battage ou vibration de la surface à nettoyer
B08B 13/00 - Accessoires ou parties constitutives, d'utilisation générale, des machines ou appareils de nettoyage
27.
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 highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.
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
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
F01D 5/30 - Fixation des aubes au rotorPieds de pales
F02C 3/073 - 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 les étages de la turbine et du compresseur étant concentriques
29.
HIGH POWER EPICYCLIC GEARBOX AND OPERATION THEREOF
A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft.
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
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
An electrical system includes: a battery; a fuel-cell pack; a load electrically coupled to the fuel-cell pack; a switching arrangement electrically coupled to the battery, the fuel-cell pack and the load; a DC-DC converter; and a control system. The switching arrangement configures the electrical system in at least one of a battery-charge mode and a combined-drive mode. In the battery-charge mode the battery is coupled in series to the fuel-cell pack and the load via the DC-DC converter for simultaneous charging of the battery and driving of the load by the fuel-cell pack. In the combined-drive mode, the battery is coupled in series to the fuel-cell pack and the load via the DC-DC converter for driving of the load by both the battery and the fuel-cell pack. The control system is configured to: monitor a parameter of an electrical power provided to the load; and control the DC-DC converter.
An electrical system includes: a battery; a fuel-cell pack; a load; a switching arrangement; and a control system. The switching arrangement selectively configures the electrical system in a battery-isolation mode and at least one of a battery-charge mode and a combined-drive mode. In the battery-isolation mode, the battery is decoupled from the fuel-cell pack and the load, and the fuel-cell pack is coupled to the load for driving of the load by the fuel-cell pack. In the battery-charge mode, the battery is coupled in series to the fuel-cell pack and the load for simultaneous charging of the battery and driving of the load by the fuel-cell pack. In the combined-drive mode, the battery is coupled in series to the fuel-cell pack and the load for driving of the load by both the battery and the fuel-cell pack.
The disclosure relates to fault protection in a DC-DC electric power converter. The converter comprises first, second, third and fourth switches connected either side of an inductor. Fifth and sixth switches are connected between respective input and output terminals and a common line, the fifth and sixth switches providing protection in a fault event.
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
H02M 1/00 - Détails d'appareils pour transformation
H02M 3/158 - 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 avec commande automatique de la tension ou du courant de sortie, p. ex. régulateurs à commutation comprenant plusieurs dispositifs à semi-conducteurs comme dispositifs de commande finale pour une charge unique
This invention concerns a fuel delivery system for an aircraft engine, comprising a fuel delivery regulator configured to receive fuel from a plurality of fuel sources for supply to the engine. An engine operating condition sensor reading is received by a control unit configured to control operation of the regulator. The control unit is configured to actuate the regulator based on a received signal from the engine operating condition sensor in order to vary the volume of fuels from the plurality of fuel sources supplied to the engine during at least a portion of a cruise phase relative to a further portion of an aircraft flight. The different fuels may comprise kerosene and an alternative fuel, such as a biofuel.
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 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
An apparatus for determining the health of a battery including a plurality N of series-connected cells. The apparatus includes an AC signal source operable to cause an AC current to flow through one or more cells of the plurality of series-connected cells, a current measurement unit operable to measure the AC current, a voltage measurement unit operable to measure an AC voltage induced across the one or more cells by the AC current, and a health determination unit operable to determine a measurement of the health of the battery based on the measured AC current and the measured AC voltage.
G01R 31/392 - Détermination du vieillissement ou de la dégradation de la batterie, p. ex. état de santé
G01R 19/02 - Mesure des valeurs efficaces, c.-à-d. des valeurs moyennes quadratiques
G01R 31/36 - Dispositions pour le test, la mesure ou la surveillance de l’état électrique d’accumulateurs ou de batteries, p. ex. de la capacité ou de l’état de charge
G01R 31/385 - Dispositions pour mesurer des variables des batteries ou des accumulateurs
G01R 31/389 - Mesure de l’impédance interne, de la conductance interne ou des variables similaires
There is provided a method of operating a gas turbine engine including a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the chamber. The nozzles include a first and second subset of fuel spray nozzles. The combustor is operable in a condition wherein the first subset is supplied with more fuel than the second. A ratio of the number of nozzles in the first subset to the second is 1:2 to 1:5. The method includes operating the engine so a reduction of 20-80% in an average of particles/kg of nvPM in the exhaust when the engine is operating at 85% available thrust for given operating conditions and when the engine is operating at 30% is obtained when fuel provided to the nozzles is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. Also provided is a gas turbine engine for an aircraft.
F23R 3/34 - Alimentation de différentes zones de combustion
F02C 3/20 - 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
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 7/228 - Division du fluide entre plusieurs brûleurs
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
36.
RECESSION RESISTANT INTERMEDIATE LAYER FOR CMC VANE
A ceramic matrix composite (CMC) component is provided that includes: a CMC body in which an environmental protection layer is completely embedded within a CMC material of the CMC body, the environmental protection layer comprising a ceramic that has a higher impact and/or environmental resistance than the CMC material. Methods for manufacturing the CMC component are also provided.
B32B 3/08 - Caractérisés par des caractéristiques de forme en des endroits déterminés, p. ex. au voisinage des bords caractérisés par des éléments ajoutés à des endroits déterminés
B32B 5/06 - Produits stratifiés caractérisés par l'hétérogénéité ou la structure physique d'une des couches caractérisés par les caractéristiques de structure d'une couche comprenant des fibres ou des filaments caractérisés par une couche fibreuse imbriquée ou cousue avec une autre couche, p. ex. de fibres, de papier
B32B 18/00 - Produits stratifiés composés essentiellement de céramiques, p. ex. de produits réfractaires
C04B 41/50 - Revêtement ou imprégnation avec des substances inorganiques
A fixture for removably securing a plurality of components includes a hollow body including a plurality of channels and a plurality of sets of holes. The fixture further includes a plurality of hollow wedges and a plurality of sets of balls. Upon an axial movement of an elongate member in a first direction, the elongate member axially moves a first hollow wedge, thereby causing one or more elastic members to be sequentially compressed and move the subsequent hollow wedges. Upon the axial movement, a frustoconical surface portion of each hollow wedge moves the corresponding set of balls radially outwards within the corresponding set of holes, such that each ball extends partially into the corresponding channel, thereby moving the corresponding component into locking engagement with the corresponding channel.
F16B 21/16 - Dispositifs sans filetage pour empêcher le mouvement relatif selon l'axe d'une broche, d'un ergot, d'un arbre ou d'une pièce analogue par rapport à l'organe qui l'entoureFixations à ergots et douilles largables sans filetage à parties séparées par gorges ou encoches pratiquées dans l'axe ou l'arbre
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
A power routing circuit having first and second input nodes VIN1, VIN2 for connection to first and second power supplies and output node VOUT1 for load connection; and includes a first main switching circuit connected between the first input node VIN1 and output node VOUT1 and conductively connects the first input node VIN1 to the output node VOUT1 if a voltage at the first input node VIN1 is higher than a voltage at the output node VOUT1, and a first auxiliary switching circuit connected between the second input node VIN2 and output node VOUT1 and conductively connects the second input node VIN2 to output node VOUT1 if a voltage at the second input node VIN2 is higher than the voltage at the output node VOUT1 and the voltage at the first input node VIN1 is at least a threshold amount lower than the voltage at the second input node VIN2.
The disclosure relates to a heat exchanger for a cooling system, the heat exchanger comprising: a first part having a first surface for connecting to a component to be cooled and a second opposing surface, the first part having a fluid flow channel extending from a fluid inlet through the first part between the first and second surfaces; a second part extending from the second surface to an external third surface, the second part having an open-cell porous structure in fluid communication with the fluid flow channel such that fluid flowing through the fluid flow channel passes into the second part and exits the heat exchanger at the external third surface.
The disclosure relates to power electronics converters having circuit breaker protection for use in case of faults. Example embodiments include a power electronics converter for converting an input AC supply having a plurality of phases to an output DC supply, the converter comprising: a plurality of AC input terminals connectable to the plurality of phases of the AC supply; first and second DC output terminals; a capacitor connected between the first and second DC output terminals; a first MOSFET connected between each of the plurality of AC input terminals and the first DC output terminal; a second MOSFET connected between each of the plurality of AC input terminals and the second DC output terminal; and a third MOSFET reverse connected in series with the first MOSFET between the first DC output terminal and each of the plurality of AC input terminals.
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
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
A combustor assembly for a gas turbine engine includes a combustor shell extending along a shell axis. The combustor shell includes a shell wall extending around the shell axis and a meter panel connected to an upstream end of the shell wall. The shell wall and the meter panel together define a combustion space therebetween. The combustor assembly further includes: a tile disposed within the combustor shell and including a tile flange portion extending radially with respect to the shell axis; a heatshield connected to the meter panel and extending at least radially towards the tile; a fastener connecting the tile flange portion to the meter panel; and a compartment thermally shielded from the combustion space. Either the heatshield and the tile together form the compartment or the tile alone forms the compartment. The fastener is spaced apart from the combustion space and at least partially disposed within the compartment.
A resin composition that is useful in the preparation of protective coatings. The resin composition comprises a polyfunctional cyanate ester, a phenol-end-modified PDMS oligomer, and an imidazolium dicyanamide catalyst. The method of preparing a resin blend by mixing a polyfunctional cyanate ester, a phenol-end-modified PDMS oligomer and an imidazolium dicyanamide catalyst to form a resin composition and curing the resin composition to form the resin blend.
C08L 65/00 - Compositions contenant des composés macromoléculaires obtenus par des réactions créant une liaison carbone-carbone dans la chaîne principaleCompositions contenant des dérivés de tels polymères
C09D 165/00 - Compositions de revêtement à base de composés macromoléculaires obtenus par des réactions créant une liaison carbone-carbone dans la chaîne principaleCompositions de revêtement à base de dérivés de tels polymères
A method of manufacturing a component includes forming a mould assembly including an initial mould unit, providing a seed crystal including a primary growth direction, determining an optimal angular orientation of the unit, rotating the unit to dispose the unit's optimal angular orientation, encasing the unit in a refractory material, and forming a refractory mould unit having a component mould including a mould wall defining a mould cavity, and a seed holder. In the optimal angular orientation, the seed crystal's primary growth direction is angled away from the wall, thereby forming a converging disposition with the wall in a of the wall's first region facing the central sprue and a diverging disposition with the wall in the wall's second region facing a mould heater. The method includes receiving the seed crystal within the seed holder and filling the mould cavity with molten castable material to form the component.
A method of operating a gas turbine engine and a gas turbine engine includes a fuel delivery system arranged to provide fuel, a combustor arranged to combust at least a proportion of the fuel, a primary fuel-oil heat exchanger arranged to have up to 100% of the fuel provided by the fuel delivery system flow therethrough, and a secondary fuel-oil heat exchanger arranged to have a proportion of the fuel from the primary fuel-oil heat exchanger flow therethrough. Fuel is arranged to flow from the primary fuel-oil heat exchanger to the secondary fuel-oil heat exchanger whereas oil is arranged to flow from the secondary fuel-oil heat exchanger to the primary fuel-oil heat exchanger. A fuel viscosity is adjusted to a maximum of 0.58 mm2/s on entry to the combustor at cruise conditions.
A combustor assembly for a gas turbine engine comprises a combustor casing. The combustor casing has a diametral height D and an axial length L. An Aspect Ratio parameter S is defined as:
A combustor assembly for a gas turbine engine comprises a combustor casing. The combustor casing has a diametral height D and an axial length L. An Aspect Ratio parameter S is defined as:
S
=
D
L
A combustor assembly for a gas turbine engine comprises a combustor casing. The combustor casing has a diametral height D and an axial length L. An Aspect Ratio parameter S is defined as:
S
=
D
L
and the S parameter lies in the range of 0.30 to 1.00.
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
F02C 3/14 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail caractérisés par l'aménagement de la chambre de combustion dans l'ensemble
An integral guide vane combustor assembly for a gas turbine engine comprises an annular array of radially extending guide vane combustors. The gas turbine engine comprises, in axial flow sequence, a compressor assembly, the integral guide vane combustor assembly, a turbine assembly, and an exhaust assembly. Each of the guide vane combustors comprises, in axial flow sequence, a guide vane portion, and a combustor body portion. The guide vane portion is formed integrally with the combustor body portion, and the guide vane portion is configured to direct a gas flow exiting the compressor assembly into the combustor body portion.
F23R 3/14 - Aménagements de l'entrée d'air pour l'air primaire créant un tourbillon au moyen d'ailettes de tourbillonnement
B33Y 80/00 - Produits obtenus par fabrication additive
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
A kinetic energy absorptive composite article comprising a plurality of plies, and a method of manufacturing the kinetic energy absorptive composite article. Each ply comprises a plurality of substantially parallel fibers encapsulated within a resin. A plurality of localised weaknesses are comprised within the fibers of the kinetic energy absorptive composite article, the locations of the plurality of localised weaknesses forming a pre-defined pattern. The pre-defined pattern comprises a first sub-pattern and a second sub-pattern superposed upon the first sub-pattern. The first sub-pattern is configured so that a crack formed by an impact event propagates substantially along a predetermined fracture path to separate the kinetic energy absorptive composite article into at least a first portion and a second portion and the second sub-pattern is configured to set a threshold below which the crack does not propagate.
A fuel swirl nozzle for a gas turbine engine comprises, in axial fuel flow sequence, a fuel inlet portion, a fuel stem portion, and a fuel/air swirler portion. The fuel inlet portion is configured to receive a flow of a fuel, with an inlet to the fuel stem portion being fluidly sealed against the fuel inlet portion, and an outlet from the fuel stem portion being fluidly sealed against the fuel/air swirler portion. The fuel/air swirler portion comprises an outer air swirler and an inner air swirler, with the inner air swirler being positioned concentrically within the outer air swirler. The outer air swirler comprises a plurality of first vanes that are arranged in an annular array within an outer air swirler body and is configured to impose a first rotational flow component on an air flow entering the outer air swirler. The inner air swirler has a plurality of second vanes that are arranged in an annular array within an inner air swirler body and are configured to impose a second rotational flow component on an air flow entering the inner air swirler. The fuel stem portion passes radially inwardly through the outer air swirler to a region radially between the outer air swirler and the inner air swirler. The inner air swirler also comprises an annular fuel gallery that is formed on a radially outwardly facing surface such that the flow of fuel enters the fuel gallery before being atomised into the air flow passing through the inner air swirler, and then further atomised into the air flow passing through the outer air swirler.
A combustor assembly for a gas turbine engine comprises a combustor casing, and an integer quantity N of fuel swirl nozzles. The combustor casing defines a total internal volume of the combustor casing V (cm3) together with a Fuel Swirl Nozzle Density ratio (m−3) that is defined as:
A combustor assembly for a gas turbine engine comprises a combustor casing, and an integer quantity N of fuel swirl nozzles. The combustor casing defines a total internal volume of the combustor casing V (cm3) together with a Fuel Swirl Nozzle Density ratio (m−3) that is defined as:
D
F
S
N
=
N
V
A combustor assembly for a gas turbine engine comprises a combustor casing, and an integer quantity N of fuel swirl nozzles. The combustor casing defines a total internal volume of the combustor casing V (cm3) together with a Fuel Swirl Nozzle Density ratio (m−3) that is defined as:
D
F
S
N
=
N
V
and the DFSN parameter is a value in the range of 200 to 1,500.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
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
F23D 11/10 - Brûleurs à pulvérisation directe de gouttelettes de liquide ou de liquide vaporisé dans l'enceinte de combustion la pulvérisation étant réalisée par un milieu gazeux, p. ex. de la vapeur d'eau
A fuel swirl nozzle for a gas turbine engine comprises, in axial fuel flow sequence, a fuel inlet portion, a fuel stem portion, and a fuel/air swirler portion. An inlet to the fuel stem portion is fluidly sealed against the fuel inlet portion, and an outlet from the fuel stem portion is fluidly sealed against the fuel/air swirler portion. The fuel/air swirler portion has a cross-sectional profile that is generally dolioform shaped along an axis parallel to the engine's longitudinal axis, and has a first end and a second end. The first end of the fuel/air swirler portion has an air inlet face, and the second end of the fuel/air swirler portion has a fuel/air mixture outlet face. The first end comprises a rim portion enclosing the air inlet face in which the rim portion has a rim portion radius (rrim) normal to an outer circumference of the inlet face.
A single-piece annular vane array for a gas turbine engine. The single-piece annular vane array comprises at least one mounting feature, a ring of stator vanes, and at least one baffle. The at least one baffle is attached at a first radial location to the ring of stator vanes, and is attached at a second radial location to the at least one mounting feature. The at least one baffle is deformable.
A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
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
The present disclosure provides an accessory power module for a gas turbine engine having a high-pressure spool and a low-pressure spool, the accessory power module comprising: a first electrical machine configured to couple to the low-pressure spool of the gas turbine engine; a casing housing the first electrical machine, the casing comprising at least one support element configured to support the first electrical machine within the casing, wherein the casing is configured for airflow therethrough; an accessory gearbox comprising an input coupling configured to couple to the high-pressure spool of the gas turbine engine, the accessory gearbox mounted to the casing; and a second electrical machine configured to couple to an output coupling of the accessory gearbox; wherein the casing is configured to be mounted along an inlet air flow path into the gas turbine engine.
A fuel cell system includes a fuel pre-heater, a fuel cell stack and a cooling circuit which is arranged to implement a Rankine cycle and includes a condenser. The fuel pre-heater is arranged to heat a flow of liquid hydrogen provided to an input thereof to provide a flow of gaseous hydrogen. The system further includes a conveying apparatus arranged to convey the gaseous hydrogen to a fuel input of the fuel cell stack such that the gaseous hydrogen is in thermal contact with coolant fluid in the condenser. The size and mass of the condenser may thereby be reduced. The pre-heater is arranged to heat coolant fluid within the cooling circuit, thereby increasing the efficiency of the Rankine cycle.
H01M 8/04082 - Dispositions pour la commande des paramètres des réactifs, p. ex. de la pression ou de la concentration
B64D 27/355 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles à combustible
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 8/04014 - Échange de chaleur par des fluides gazeuxÉchange de chaleur par combustion des réactifs
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 8/10 - Éléments à combustible avec électrolytes solides
A method of operating a gas turbine engine, the gas turbine engine including an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a main gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system arranged to supply oil to lubricate the main gearbox; and a heat exchange system arranged to transfer heat between the oil and the fuel, the oil having an average temperature of at least 180° C. on entry to the heat exchange system at cruise conditions. The method includes controlling the heat exchange system so as to raise the fuel temperature to at least 135° C. on entry to the combustor at cruise conditions.
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
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 (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of:
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of:
gearbox location/engine length
is in a range from 0.19 to 0.45.
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
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
F16H 1/28 - Transmissions à engrenages pour transmettre un mouvement rotatif avec engrenages à mouvement orbital
An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 300, 400, 500, 600, 700 configured to introduce a pressurised fluid into a core airflow A at a region downstream of the turbine 19, wherein a rearward movement of the turbine 19 or a shaft 26 in a shaft break event directly actuates the pressure equalisation apparatus 300, 400, 500, 600, 700 to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.
F01D 21/14 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à d'autres conditions spécifiques
An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 200, 300, 400, 500, 600, 700 configured to introduce a pressurised fluid into a core airflow A at a region directly downstream of the turbine 19 in the event of a shaft break to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
An apparatus and method for reducing a pressure differential across a turbine 19 of a gas turbine engine 10 during a shaft break event, comprises a pressure equalization apparatus 200 configured to introduce a pressurised fluid into a core airflow A at a region downstream of the turbine 19 in the event of a shaft break to directly increase a local pressure at the downstream region 29 of the turbine 19 and thereby reduce the pressure differential across the turbine 19. The pressure equalization apparatus comprises a sensor 216 configured to directly detect a shaft break event. The reduction in the pressure differential may result in a reduction in the acceleration of the turbine 19.
F01D 21/14 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à d'autres conditions spécifiques
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
A gas turbine engine configured to combust gaseous hydrogen fuel comprises a combustor comprising an annular combustion chamber outer casing surrounding an inner combustor case and a fuel manifold configured to provide gaseous fuel to a plurality of fuel injectors. The fuel manifold is formed integrally with the combustion chamber outer casing.
An annulus filler for mounting to a rotor disc of a gas turbine engine includes a coupling portion and an outer lid coupled to the coupling portion. The outer lid includes a leading edge, a trailing edge, and a main portion including an outer radial surface. The outer lid further includes a protruding portion connected to the main portion and extending radially outwardly from the main portion. The protruding portion is axially disposed between and spaced apart from the leading edge and the trailing edge. The protruding portion includes a protruding surface contiguous with and extending radially outwardly from the outer radial surface, such that the protruding surface is configured to redirect air drawn through the gas turbine engine.
A method for scanning at least one component includes placing the at least one component inside a storage vessel; filling the storage vessel with a cryogenic material; cooling the at least one component; placing the storage vessel between an x-ray source and a detector of a CT scanner; generating, via the x-ray source, an x-ray cone beam after cooling of the component that passes through the storage vessel while the at least one component is disposed within the storage vessel; receiving the x-ray cone beam at the detector; and generating an x-ray image of the at least one component.
G01N 1/42 - Traitement à basse température des échantillons, p. ex. cryofixation
G01N 23/046 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p. ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et formant des images des matériaux en utilisant la tomographie, p. ex. la tomographie informatisée
63.
COMPUTER-IMPLEMENTED METHODS, APPARATUS, COMPUTER PROGRAMS, AND NON-TRANSITORY COMPUTER-READABLE STORAGE MEDIUMS
A computer-implemented method including: receiving a first data set including a plurality of values for a plurality of features; removing one or more features and associated values from the first data set to generate a second data set; determining feature importance of at least a subset of the features of the second data set using multiple-evaluation criteria; and performing an action using at least one feature of the second data set and the determined feature importance.
G06F 16/215 - Amélioration de la qualité des donnéesNettoyage des données, p. ex. déduplication, suppression des entrées non valides ou correction des erreurs typographiques
G06F 16/28 - Bases de données caractérisées par leurs modèles, p. ex. des modèles relationnels ou objet
A component for a hot section of an engine comprises a substrate and an infrared-reflective layer configured to reflect infrared radiation away from the substrate and in some cases the infrared-reflective layer has: a spectral reflectance R_λ of no less than about 0.5 when measured by spectrophotometry, according to ISO 15368:2021, using incident electromagnetic radiation having a wavelength λ no less than about 500 nm and no greater than about 1 mm; and a spectral emissivity ε_λ of less than about 0.4, for example no greater than about 0.3, when measured using incident electromagnetic radiation having a wavelength λ no less than about 500 nm and no greater than about 1 mm.
A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.
F01D 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
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
F02K 3/04 - 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
A power system for an aircraft comprises a gas turbine engine, arranged to burn a fuel in a combustor so as to provide power to the aircraft; a plurality of fuel tanks arranged to contain fuel to be used to provide power to the aircraft; and a fuel manager. A first fuel tank of the plurality of fuel tanks is arranged to contain a first fuel, and a second tank of the plurality of fuel tanks is arranged to contain a second, different, fuel. The fuel manager is arranged to store information on the fuel contained in each fuel tank and to control fuel supply so as to take fuel from the second tank for engine start-up, before switching to the first fuel tank.
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
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
67.
COMPUTER-IMPLEMENTED METHODS, APPARATUS, COMPUTER PROGRAMS, AND NON-TRANSITORY COMPUTER-READABLE STORAGE MEDIUMS
A computer-implemented method including: receiving a first data set including a plurality of values for a plurality of features; identifying at least a first feature of the first data set that is non-redundant and at least a second feature of the first data set that is redundant; identifying one or more clusters of features in the plurality of features of the first data set, a first cluster of the one or more clusters including at least the first feature and the second feature; and controlling a display to display the first feature and one or more redundant features from the first cluster, the displayed one or more redundant features from the first cluster including the second feature.
A method of magnetising or demagnetising an annular component for a rotary machine, a flux arrangement for performing a magnetising or demagnetising method, and a flux assembly for such a flux arrangement; wherein, the annular component includes an alternating arrangement of radial elements and angular elements for forming a Halbach array. A magnetizer is caused to induce magnetic flux in a primary set of elements of the annular component including a primary radial element and an adjacent primary angular element. A shield element shields a secondary angular element of the annular component from magnetic flux from the magnetizer.
H02K 15/00 - Procédés ou appareils spécialement adaptés à la fabrication, l'assemblage, l'entretien ou la réparation des machines dynamo-électriques
H01F 13/00 - Appareils ou procédés pour l'aimantation ou pour la désaimantation
H02K 11/01 - Association structurelle de machines dynamo-électriques à des organes électriques ou à des dispositifs de blindage, de surveillance ou de protection pour le blindage contre les champs électromagnétiques
H02K 1/2783 - Aimants montés en surfaceAimants sertis le noyau étant muni d’aimants disposés en réseaux de Halbach
A gas turbine engine for an aircraft comprises: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a bypass duct delimited by a bypass duct inner wall and a bypass duct outer wall and located radially outwardly from the engine core and downstream of the fan; and an outlet guide vane assembly, located within the bypass duct and, comprising a plurality of outlet guide vanes distributed circumferentially within the bypass duct, each outlet guide vane extending radially along a span between the bypass duct inner wall and the bypass duct outer wall, wherein a space-chord ratio of at least one outlet guide vane, at 50% of the span length from the bypass duct inner wall, is less than 0.72.
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
An aircraft includes a fuselage and a wing, at least one hydrogen fuel tank and a fuel line configured to delivery fuel from the fuel tank to the engine. The fuel tank is installed externally of the fuselage and wing, and comprises an external aeroshell containing a plurality of pressure vessels configured to contain pressurised fuel therein.
A hydrogen fuelled aircraft includes a fuselage, a wing, at least one hydrogen fuelled gas turbine engine, at least one hydrogen fuel tank configured to store liquid or supercritical hydrogen and a hydrogen fuel line configured to delivery hydrogen fuel from the fuel tank to the engine. The gas turbine engine and the hydrogen fuel tank are each installed externally on the wing, and the hydrogen fuel line does not pass through the fuselage.
A propulsive aircraft gas turbine engine comprises a turbine disposed in a gas turbine engine core flow and a recuperator heat exchanger disposed downstream of the turbine in gas turbine engine core flow, the recuperator heat exchanger being configured to transfer heat from the gas turbine engine core flow to gas turbine engine fuel. The recuperator heat exchanger is configured to accommodate a portion of the gas turbine engine core flow therethrough, the remainder being bypassed around the recuperator heat exchanger.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
B64D 37/30 - Circuits de carburant pour carburants particuliers
B64D 37/34 - Conditionnement du carburant, p. ex. réchauffage
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 method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
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
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
74.
SYSTEM AND METHOD FOR DETERMINING THE CHARACTERISTICS OF A MACHINING PROCESS
A method for determining one or more characteristics of a hole includes obtaining depth data, tool velocity data, process current data, and a plurality of gates. The method further includes determining a plurality of first, second, and third key points. The plurality of first, second, and third key points together form a plurality of key points. The method further includes obtaining a plurality of rules. Each rule includes one of: a relationship between a respective key point and a value associated with the respective key point; and a relationship between two or more respective key points. The method further includes determining the one or more characteristics of the hole based on the plurality of rules and the plurality of key points.
A compressor of an aircraft air pressurization system comprises: a rotor configured to be mechanically coupled to a spool of a gas turbine engine; and a housing, wherein the rotor is supported for rotation within the housing about a rotor axis, wherein the rotor and housing define a primary air channel extending between an inlet of the blower compressor for receiving the inlet flow of air and an outlet from which the blower compressor is configured to output pressurized air to a delivery line. The housing further comprises a bleed opening in a wall of the housing between the inlet and the outlet, the bleed opening extending circumferentially about the rotor axis, wherein the bleed opening is for contaminants within the inlet flow of air, which have been driven outward due to the rotation of the rotor, to be bled out of the blower compressor.
B64D 13/02 - Aménagements ou adaptations des appareils de conditionnement d'air pour équipages d'aéronefs, passagers ou pour emplacements réservés au fret l'air étant pressurisé
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
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A compressor of an aircraft air pressurization system comprises: a rotor configured to be mechanically coupled to a spool of a gas turbine engine; an oil-lubricated bearing for supporting rotation of the rotor, wherein the oil-lubricated bearing is disposed within a bearing chamber; and a seal assembly for restricting oil from the bearing chamber from reaching an inlet of the blower compressor, wherein the seal assembly comprises: a seal disposed between the bearing chamber and an air buffer chamber; an air buffer inlet configured to receive a flow of pressurized air into the air buffer chamber; and an air buffer outlet configured to allow the flow of pressurized air to be exhausted from the air buffer chamber together with oil from the bearing chamber that has passed through the seal and become entrained within the flow of pressurized air.
F01D 11/04 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages par obturation non contact, p. ex. du type labyrinthe utilisant un fluide d'obturation, p. ex. de la vapeur
A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. 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.
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
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
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
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
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 3/107 - 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
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
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 3/068 - 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 caractérisé par une longueur axiale courte par rapport au diamètre
A method of analysing a non-contact excitation system for a spin rig includes, calculating a magnetic field per test piece element, integrating the magnetic field over the test piece element surface to get a force per test piece element, performing a Fourier analysis of the forces acting on the test piece element as it moves past the periodic array of exciter elements, calculating the damping added to the test piece element due to magnetic induction, calculating a vibration response of the test piece element as it moves past the stationary array of exciter elements, calculating a transient response of the test piece element where the stationary array of exciter elements is rapidly moved away from the test piece elements, and using the calculated transient response of the test piece element to calculate a required electromagnetic holding force for the array of exciter elements.
An electric machine includes a stator having a phase arrangement, and a rotor. The phase arrangement includes legs connected in parallel at first and second primary junctions. Each leg includes a plurality of coils connected in series through a respective intermediate junction. Each leg is connected to at least one other leg by a branch at the respective intermediate junctions such that the phase arrangement is a bridge circuit. The phase arrangement conducts a motor current through the respective coils between the first and second primary junctions. The phase arrangement permits an alignment current to flow between the first and second primary junctions along an alignment current path which passes through at least one coil of two different legs via a respective branch. The motor currents cause a torque to be applied to the rotor, and the alignment current causes a translational force to be applied to the rotor.
An electric machine includes a stator having a phase arrangement, and a rotor. The phase arrangement includes first and second legs connected in parallel at first and second primary junctions. The legs each include first and second coils connected in series through a first intermediate junction. The intermediate junctions are connected by a branch such that the phase arrangement is a bridge circuit. The phase arrangement permits an alignment current to flow between the primary junctions via the branch, the alignment current being conducted through an alignment current path passing through one coil of each leg. The phase arrangement includes a negative impedance converter to add a negative electrical impedance to the alignment current path by introducing additional electrical energy into the respective alignment current path. The alignment current causes a translational force to be applied to the rotor for maintaining alignment of the rotor with respect to the stator.
A temperature measurement system for a gas turbine engine, the gas turbine engine including, in axial flow sequence, a compressor section, a combustor section having plural fuel spray nozzles, and a turbine section. The temperature measurement system includes one or more optical thermometers, each optical thermometer configured to measure the temperature of a component washed by the working gas of the engine, the or each component being in the combustor section or the turbine section at a first position along the axis of the engine.
G01K 1/02 - Moyens d’indication ou d’enregistrement spécialement adaptés aux thermomètres
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F02C 7/22 - Systèmes d'alimentation en combustible
G01K 7/02 - Mesure de la température basée sur l'utilisation d'éléments électriques ou magnétiques directement sensibles à la chaleur utilisant des éléments thermo-électriques, p. ex. des thermocouples
G01K 11/32 - Mesure de la température basée sur les variations physiques ou chimiques, n'entrant pas dans les groupes , , ou utilisant des changements dans la transmittance, la diffusion ou la luminescence dans les fibres optiques
A method of operating a gas turbine engine including: a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor, wherein the fuel management system includes two fuel-oil heat exchangers through which oil and fuel flow, which are arranged to transfer heat between the oil and fuel and include primary and secondary fuel-oil heat exchangers; a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the heat exchangers; and a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel which has passed through the primary heat exchanger to be returned to the inlet. The method includes selecting one or more fuels such that the calorific value of the fuel provided to the gas turbine engine is at least 43.5 MJ/kg.
F02C 9/38 - Commande de l'alimentation en combustible caractérisée par un étranglement de l'admission du combustible et un retour du combustible au réservoir
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
A gas turbine engine includes a core engine having a compressor, a combustor and a turbine in sequential air flow series. The engine further includes a 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 and a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit. At least first and second compressor bleed offtakes are at different pressure stages of the compressor, each being configured to bleed a portion of air from the compressor. At least first and second compressor bleed offtake valves are configured to control flow through the first and second bleed offtakes respectively. The burner is configured to receive bleed air from the compressor bleed offtakes.
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
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 gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.
F02C 7/10 - Chauffage de l'air d'alimentation avant la combustion, p. ex. par les gaz d'échappement au moyen d'échangeurs de récupération de chaleur
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 gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.
F02C 7/10 - Chauffage de l'air d'alimentation avant la combustion, p. ex. par les gaz d'échappement au moyen d'échangeurs de récupération de chaleur
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
86.
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 57/02 - Boîtes de vitessesMontage de la transmission à l'intérieur
F16H 57/08 - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
A non-contact excitation system for a spin rig includes: a plurality of test piece elements for attachment to a test piece for rotation about an axis of rotation of the spin rig; and a plurality of exciter elements arranged around the axis of rotation, wherein the test piece elements and the exciter elements are magnetically attractive; wherein the test piece elements and/or the exciter elements include magnets; each exciter element being moveable between an active position, in which a magnetic force is provided between each exciter element and at least one of the test piece elements, and an inactive position, in which the magnetic force between the exciter magnets and the at least one test piece element is reduced.
There is provided an air pressurisation system for an aircraft. The air pressurisation system comprises a blower compressor, a delivery line and a catalyst material. The blower compressor is configured to be mechanically coupled to a spool of a gas turbine engine and configured to receive an inlet flow of gases from a bypass duct of the gas turbine engine. The delivery line is configured to convey gases received from the blower compressor to an airframe port for supply to an airframe system. The catalyst material is disposed along the delivery line and configured to catalyse a reaction of volatile organic compounds within gases conveyed by the delivery line.
B64D 13/06 - Aménagements ou adaptations des appareils de conditionnement d'air pour équipages d'aéronefs, passagers ou pour emplacements réservés au fret l'air étant climatisé
B64D 13/08 - Aménagements ou adaptations des appareils de conditionnement d'air pour équipages d'aéronefs, passagers ou pour emplacements réservés au fret l'air étant climatisé l'air étant réchauffé ou refroidi
A battery cooling block, including: a housing defining a flow chamber and including a first and second port configured to allow flow of coolant through flow chamber along a first axis; the housing further including a plurality of compartments disposed in the flow chamber, each compartment extending along a respective compartment axis perpendicular to first axis and each compartment configured to receive respective cylindrical battery cell; wherein each compartment is defined by partially open compartment wall within flow chamber, compartment wall including a plurality of wall segments extending parallel to compartment axis and which are circumferentially spaced apart around compartment axis wherein respective battery cell is partially exposed to flow of coolant. A battery assembly, including: plurality of cylindrical battery cells; and battery cooling block as described above. Each of the plurality of cylindrical battery cells is received in respective one of the plurality of compartments of battery cooling block.
H01M 50/291 - MonturesBoîtiers secondaires ou cadresBâtis, modules ou blocsDispositifs de suspensionAmortisseursDispositifs de transport ou de manutentionSupports caractérisés par des éléments d’espacement ou des moyens de positionnement dans les racks, les cadres ou les blocs caractérisés par leur forme
H01M 10/613 - Refroidissement ou maintien du froid
H01M 50/213 - Bâtis, modules ou blocs de multiples batteries ou de multiples cellules caractérisés par leur forme adaptés aux cellules ayant une section transversale courbée, p. ex. ronde ou elliptique
Described herein is a non-contact excitation system for a spin rig comprising: a plurality of test piece elements for attachment to a test piece for rotation about an axis of rotation of the spin rig; and a plurality of exciter elements arranged around the axis of rotation, wherein the test piece elements and the exciter elements are magnetically attractive; wherein the test piece elements and/or the exciter elements comprise magnets; and wherein, in use, the exciter elements are arranged around the test piece; wherein each exciter element is axially and radially offset from the plurality of test piece elements. Also described is a method of testing a test piece using the system.
A method for scanning of an object in a scanning apparatus includes disposing the object and a complementary object on a support of the apparatus, so that the objects are positioned between an imaging beam emitting element and an imaging beam receiving element oppositely disposed to either side of the support. The support is rotatable relative to the emitting and receiving elements about an axis of rotation to allow creation of an image from projections each taken at a different relative angle. A volume of the complementary object is solid or filled with a filling material and the complementary object is configured to reduce the variation in imaging beam attenuation across the objects or the part to be scanned of the object at the multiple relative angles of rotation. The method includes operating the scanning apparatus at the multiple relative angles of rotation to produce an image of the object.
G01N 23/046 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p. ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et formant des images des matériaux en utilisant la tomographie, p. ex. la tomographie informatisée
G01N 23/083 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p. ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et mesurant l'absorption le rayonnement consistant en rayons X
G01N 23/18 - Recherche de la présence de défauts ou de matériaux étrangers
A computer-based control system for a gas turbine engine includes: a controller including control logic, the controller obtains a set of inputs formed by measurements of one or more process variables and values of one or more engine operation set points, and the control logic determines one or more process command values gas turbine engine operation in response to the set of inputs; and an event detection unit obtaining further measurements of one or more process variables and determine whether an abnormal event has occurred based on the further measurements of one or more process variables. The controller also updates the tuning variables using an event accommodation data array, a given tuning variable being changed in response to detection of a given abnormal event when the respective element in the event accommodation data array is assigned an intervention value.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gazCommande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
A coating system includes a bond layer including CoNiCrAlY; a thermal barrier layer at least partially disposed on the bond layer and including yttria stabilised zirconia; a mesh layer connecting the thermal barrier layer to the bond layer and including CoNiCrAlY; and an abradable layer disposed on the thermal barrier layer and including magnesium spinel. The mesh layer includes a plurality of cells connected to each other and is at least partially embedded in the thermal barrier layer. Each cell defines a cell opening therethrough and the cell opening of each of the plurality of cells at least partially receives the thermal barrier layer therein. The abradable layer and the thermal barrier layer are at least partially removable from the mesh layer and the bond layer without damaging them.
The present disclosure provides a podded engine adaptor configured to secure a podded engine to an aircraft, comprising: an airframe mounting interface configured for attachment to an airframe of an aircraft; an engine mounting interface configured for attachment to an engine; wherein the podded engine adaptor is configured to be attached to an engine at the engine mounting interface and to an airframe of the aircraft at the airframe mounting interface, so as to indirectly couple an engine to the aircraft via the podded engine adaptor.
B64D 27/40 - Aménagements pour le montage de groupes moteurs sur aéronefs
B64D 29/04 - Nacelles, carénages ou capotages des groupes moteurs montés dans le fuselage
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
A method of operating a gas turbine engine having an engine core having a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a heat exchange system having at least one fuel-oil heat exchanger arranged to transfer heat to the fuel; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located downstream of the at least one fuel-oil heat exchanger. The method comprises controlling the heat exchange system so as to raise the fuel temperature to at least 135° C. on entry to the combustor at cruise conditions.
An apparatus for metallurgical heat-treatment includes a solid-oxide electrolyser, a furnace and a heat-exchanger. The electrolyser is arranged to electrolyse water and provide resulting hydrogen to the furnace. A first portion of the hydrogen from the electrolyser is combusted in a combustor to heat the furnace. A second portion provides a treatment atmosphere including hydrogen for the heat-treatment of a metal or metal alloy object. Water vapour output by the combustor is provided to the heat-exchanger which transfers heat within the water vapour to the solid-oxide electrolyser to improve or maintain its efficiency. In contrast to apparatus of the prior art, the apparatus does not produce carbon dioxide at the point of use. By applying waste heat, carried by the water vapour output from the combustor, to the electrolyser, the power consumption of the electrolyser is reduced for a given rate of electrolysis.
The disclosure relates to estimation of junction temperatures of transistors in a power electronics converter. Example embodiments include a method of estimating a junction temperature of a transistor in a power electronics converter configured to convert between first and second supply voltages, the method comprising: providing a gate switching signal to the transistor; measuring a rate of change of current through the converter during a switching period of the transistor; and outputting an estimated junction temperature of the transistor based on the measured rate of change of current, wherein the rate of change of current is measured while a gate voltage of the transistor is above a gate threshold voltage and a drain-source voltage across the transistor is above a predetermined fraction of the first or second supply voltage whereby the rate of change of current is measured in a linear region.
G01K 7/01 - Mesure de la température basée sur l'utilisation d'éléments électriques ou magnétiques directement sensibles à la chaleur utilisant des éléments semi-conducteurs à jonctions PN
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
98.
PREVENTING PROPAGATION OF CELL FAILURES IN BATTERY PACKS
Battery packs, which may be used in vehicles, include: a plurality of battery cells; one or more flow control devices; and sensing and control circuitry configured to: measure one or more parameters indicative of an onset of thermal runaway of one or more of the plurality of cells; determine, based on a change in one or more of the measured parameters, that one or more of the plurality of battery cells has begun or is at risk of beginning thermal runaway; and activate one or more of the flow control devices in response to the determination. Each of the one more flow control devices is configured to, when activated, direct a flow of fluid past one or more of the plurality of cells to cool the cells and/or carry media vented by one or more of the cells following a thermal runaway event away from the one or more cells.
A method of thrust control for an airbreathing jet engine includes obtaining a demanded thrust setting, a value of static pressure at a first axial location of a combustion chamber of the airbreathing jet engine and a value of static pressure at a second axial location of the combustion chamber of the airbreathing jet engine. The second axial location is downstream of the first axial location. The method also includes obtaining a ratio of the value of static pressure at the first axial location of the combustion chamber to the value of static pressure at the second axial location of the combustion chamber; and controlling a fuel flow rate of the airbreathing engine based at least in part on the demanded thrust setting and the ratio of the value of the static pressure at the first axial location to the value of the static pressure at the second axial location.
F02B 3/06 - Moteurs caractérisés par la compression d'air et l'addition subséquente de combustible avec allumage par compression
F02B 23/06 - Autres moteurs caractérisés par des chambres de combustion d'une forme ou d'une structure particulières pour améliorer le fonctionnement avec allumage par compression l'espace de combustion étant disposé dans le piston moteur
An estimation of junction temperatures of transistors in power electronics converter. Example embodiments include method of calibrating electrical power system, electrical power system including: power electronics converter configured to convert between first and second voltage supplies, converter having plurality of semiconductor switches; controller configured to provide switching signals to each semiconductor switch; current sensor arranged to measure current through converter to one of first and second voltage supplies; temperature sensor arranged to measure temperature of one or more semiconductor switches; and junction temperature measurement module configured to receive current signal and temperature signal, method including: controller providing gate switching signal to transistor; junction temperature measurement module measuring threshold voltage of transistor from current signal and temperature of transistor from temperature sensor during switching period of transistor; and junction temperature measurement module updating stored calibration defining relationship between rate of change of current and estimated junction temperature of transistor.
G01K 7/00 - Mesure de la température basée sur l'utilisation d'éléments électriques ou magnétiques directement sensibles à la chaleur
G01R 15/18 - Adaptations fournissant une isolation en tension ou en courant, p. ex. adaptations pour les réseaux à haute tension ou à courant fort utilisant des dispositifs inductifs, p. ex. des transformateurs
