A fuel system for a gas turbine engine comprises 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, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, and an outlet baffle positioned between the burner and the heat exchanger. The outlet baffle is configured to introduce turbulence to combustion gases entering the heat exchanger.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset is 1:3 to 1:6. A thrust nvPM emissions index ratio is EImaxTO/FmaxTO/EIidle/Fidle. EIidle is system loss corrected nvPM emissions index in mg/kg of the engine operating at around 7% available thrust. EImaxTO is system loss corrected nvPM emissions index in mg/kg of the engine at around 100% available thrust. FmaxTO is thrust of the engine at around 100% available thrust. Fidle is thrust at around 7% available thrust. The thrust nvPM emissions index ratio is between 0.0009 and 0.02.
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
EI
idle
×
W
f
,
idle
EI
maxTO
×
W
f
,
maxTO
.
A gas turbine engine for an aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include a first and second subset. The combustor is operable so each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
EI
idle
×
W
f
,
idle
EI
maxTO
×
W
f
,
maxTO
.
EIidle is the system loss corrected nvPM emissions index in mg/kg of the engine at around 7% available thrust. EImaxTO is the index at around 100% available thrust. Wf,idle is the rate of fuel to the nozzles in kg/s at around 7% available thrust. Wf,maxTO is the rate of fuel to the nozzles in kg/s at around 100% available thrust. The index ratio is between 0.357 and 8.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
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
A gas turbine engine comprises an engine core having first and second engine core exhaust paths arranged to pass first and second portions respectively of the mass flow of the engine’s core exhaust mass. A heat-exchange system comprises a recuperator system disposed within the first engine core exhaust path and arranged to transfer heat from said first portion to a buffer fluid, and a heat exchanger arranged to transfer heat from the buffer fluid to fuel within a fuel path arranged to convey fuel to the engine’s combustor. The engine provides for heat to be recovered from the engine’s core exhaust flow to the engine’s fuel supply, thus improving thermal efficiency, but without significantly impeding the engine core exhaust flow or presenting the significant fire or explosion risk associated with a recuperator arranged to heat fuel directly.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
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 power system for an aircraft includes at least one gas turbine engine arranged to burn a fuel so as to provide power to the aircraft; at least one first fuel tank arranged to be used to power ground-based operation of the aircraft; at least one secondary fuel tank arranged to contain a fuel to be used to power the aircraft in flight; and a fuel manager arranged to control fuel supply so as to take fuel from only the at least one first fuel tank to power at least the majority of ground-based operations.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p. ex. à la température, à la pression, à la vitesse du rotor
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 37/04 - Disposition de ceux-ci à l'intérieur ou sur les aéronefs
G08G 5/32 - Gestion des plans de vol pour la préparation des plans de vol
There is provided a reheat assembly 300, 300A for a gas turbine engine 10. The reheat assembly 300, 300A comprises a support duct 340 and a flameholder 370. The flameholder 370 comprises a flange portion 32 defining an inlet aperture 374 to an interior of the flameholder 370 and a boss 36 extending away from the flange portion 32 into the interior of the flameholder 370. The flameholder 370 is mounted to the support duct 340 by a fastener 38 extending through the support duct 340 into a hole 37 defined by the boss 36. The flameholder 370 is configured to receive a flow of air via the inlet aperture 374 to cool the boss 36.
F02K 3/11 - Chauffage du flux dérivé à l'aide de brûleurs ou de chambres de combustion
F02K 3/10 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant avec réchauffage supplémentaire du fluide de travailLeur commande par postcombustion
There is provided a flameholder 370 for a reheat assembly 300, 300A of a gas turbine engine 10. The flameholder comprises an internal flow passageway 376 extending from an inlet aperture 374 and defining a flow direction F for flow through the flameholder 370. The internal flow passageway 376 is defined by an internal surface 34 of the flameholder 370. A grid of recesses 31 is formed in the internal surface 34 of the flameholder 370.
F23R 3/20 - Moyens de stabilisation de la flamme, p. ex. accroche-flamme de postcombustion d'ensembles fonctionnels à propulsion par réaction comprenant des moyens d'injection du combustible
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.
A fuel system for a hydrogen fuelled gas turbine engine includes a main hydrogen fuel storage unit, a hydrogen fuel pump configured to be supplied with hydrogen from the hydrogen storage unit, a hydrogen fuel preheater configured to be supplied with high pressure hydrogen from the hydrogen fuel pump, and configured to supply heated gaseous hydrogen to a combustor of the gas turbine engine, and a hydrogen priming tank configured to store compressed gaseous hydrogen and to deliver gaseous hydrogen to at least the hydrogen fuel pump and preheater. The fuel system comprises a high-pressure gaseous hydrogen fuel offtake downstream in hydrogen fuel flow of the fuel pump in fluid communication with the hydrogen priming tank, and configured to fill the hydrogen priming tank with high pressure gaseous hydrogen.
A method of starting a liquid hydrogen fuelled gas turbine engine of an aircraft propulsion system, wherein the aircraft propulsion system includes a hydrogen storage tank configured to store liquid hydrogen, a liquid hydrogen pump configured to pump hydrogen in at least a liquid state, a core combustor configured to receive hydrogen fuel from the hydrogen fuel pump, and a hydrogen fuel vent provided downstream of the liquid hydrogen pump, and configured to selectively vent hydrogen fuel. The method includes, in a liquid priming step, flowing hydrogen from the hydrogen storage tank through the liquid hydrogen pump and venting hydrogen through the hydrogen fuel vent until the hydrogen pump is primed with liquid hydrogen, then, in a liquid pumping step, operating the liquid hydrogen pump to pump liquid hydrogen to the core combustor at a required flow rate and pressure for engine ignition in an engine ignition step.
An apparatus for inspecting a gas turbine engine component having a plurality of cooling apertures includes a main body including a plurality of body apertures configured to at least partially align with the plurality of cooling apertures. Each body aperture from the plurality of body apertures extends through the main body. The apparatus further includes a seal connected or connectable to the main body and configured to engage with the gas turbine engine component. The seal includes a plurality of seal apertures corresponding to the plurality of body apertures. Each seal aperture from the plurality of seal apertures extends through the seal. The plurality of seal apertures is at least partially aligned with the plurality of body apertures of the main body, such that the plurality of seal apertures is disposed or disposable in fluid communication with the plurality of body apertures.
A method of determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft. The method includes: exposing the surface of a piezoelectric crystal to the fuel; measuring a vibration parameter of the piezoelectric crystal; and determining one or more fuel characteristics of the fuel based on the vibration parameter. Also disclosed is a fuel characteristic determination system, a method of operating an aircraft, and an aircraft.
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
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 37/00 - Aménagements relatifs à l'alimentation des groupes moteurs en carburant
G01N 21/33 - CouleurPropriétés spectrales, c.-à-d. comparaison de l'effet du matériau sur la lumière pour plusieurs longueurs d'ondes ou plusieurs bandes de longueurs d'ondes différentes en recherchant l'effet relatif du matériau pour les longueurs d'ondes caractéristiques d'éléments ou de molécules spécifiques, p. ex. spectrométrie d'absorption atomique en utilisant la lumière ultraviolette
G01N 21/35 - CouleurPropriétés spectrales, c.-à-d. comparaison de l'effet du matériau sur la lumière pour plusieurs longueurs d'ondes ou plusieurs bandes de longueurs d'ondes différentes en recherchant l'effet relatif du matériau pour les longueurs d'ondes caractéristiques d'éléments ou de molécules spécifiques, p. ex. spectrométrie d'absorption atomique en utilisant la lumière infrarouge
G01N 21/3577 - CouleurPropriétés spectrales, c.-à-d. comparaison de l'effet du matériau sur la lumière pour plusieurs longueurs d'ondes ou plusieurs bandes de longueurs d'ondes différentes en recherchant l'effet relatif du matériau pour les longueurs d'ondes caractéristiques d'éléments ou de molécules spécifiques, p. ex. spectrométrie d'absorption atomique en utilisant la lumière infrarouge pour l'analyse de liquides, p. ex. l'eau polluée
A nuclear power system wherein: a nuclear heat source (10), turbine (15), heat dissipator (20), compressor (25), and electric heater (35) form a fluid flow circuit for channelling a working fluid; the turbine is connected to the compressor and a generator (40) by a connecting shaft (30) such that when the turbine rotates electrical energy is generated; a sensor system (55) is connected to a controller (60), and is configured to acquire data from at least the electric heater, the nuclear heat source, the generator, a battery (50), and a load (45); and the controller is configured to control the distribution of the electrical energy between a load, the battery, and the electric heater, based on the data provided by the sensor system, such that the electrical energy output to the grid can be varied, whilst maintaining a constant temperature, pressure, and massflow of the working fluid entering the turbine.
G21D 3/12 - Régulation de différents paramètres dans l'installation par ajustement du réacteur en réponse uniquement aux changements se produisant dans la demande du moteur
G21C 17/022 - Dispositifs ou dispositions pour la surveillance du réfrigérant ou du modérateur pour la surveillance de réfrigérants ou de modérateurs liquides
An aircraft propulsion system comprises a core gas turbine engine (201) comprising a core compressor (202, 204) configured to provide core air to a core combustor (206) and a core turbine (208, 209) in fluid flow series. An auxiliary compressor (220) is provided, which is separate to the core compressor (202, 204), and configured to provide air to an auxiliary air system (218). A heat exchanger (218) is provided, which is configured to transfer heat from air from the auxiliary compressor (220) to fuel in the main fuel conduit (217) prior to provision to the core combustor (206).
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 37/30 - Circuits de carburant pour carburants particuliers
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
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 nosecone for a ducted fan gas turbine engine is shown. The nosecone includes a body with an outer surface that tapers in axial extent from an apex to a base of the body, the apex defining a position of 0 percent axial extent and the base defining a position of 100 percent axial extent, and in which the outer surface includes a plurality of undulations extending between a first location on the outer surface and a second location on the outer surface, wherein the first location is positioned at from 0 to 50 percent of axial extent and the second location is positioned at from 85 to 100 percent of axial extent.
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes Entrées d'air pour ensembles fonctionnels de propulsion par réaction
A crawler robot comprising a proximal end section, a distal end section and a central core section, through each section passes a central channel; the distal end section comprises a rigid core surrounding a deformable foot, the proximal end section comprises a rigid core surrounding a deformable foot, and the central core section comprises a deformable annular body surrounding an extension spring, surrounding the annular body is at least one balloon, and wherein fluid carrying conduits are connected to the balloon in the central core section, and to the deformable foot in the distal end section and the deformable foot in the proximal end section.
B62D 57/02 - Véhicules caractérisés par des moyens de propulsion ou de prise avec le sol autres que les roues ou les chenilles, seuls ou en complément aux roues ou aux chenilles avec moyens de propulsion en prise avec le sol, p. ex. par jambes mécaniques
B08B 9/043 - Nettoyage des surfaces intérieuresÉlimination des bouchons utilisant des dispositifs de nettoyage introduits dans et déplacés le long des tubes déplacés par liaison mécanique actionnée de l'extérieur, p. ex. poussés ou tirés dans les tubes
F16L 55/26 - Hérissons ou chariots, c.-à-d. dispositifs pouvant se déplacer dans un tuyau ou dans une conduite et portant ou non un moyen de propulsion autonome
A crawler robot has a proximal end section, a distal end section and a core section. Each of the proximal end section, the distal end section and core section has a hollow core. The core section has a flexible body and is surrounded by at least one balloon that is supplied by a fluid carrying conduit. The proximal end section and the distal end section are connected to fluid carrying conduits and have at least three thrust vents each for venting the supplied fluid to generate a thrust force.
A method of operating a gas turbine engine is disclosed, the gas turbine engine including a combustor arranged to combust a fuel and a fuel management system arranged to provide the fuel to the combustor. The fuel management system includes 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 includes controlling the fuel management system so as to raise the fuel temperature to at least 135° C. on entry to the combustor 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
A method of operating an electrical power system for an aircraft. The electrical power system includes a semiconductor-based active power converter and a contactor coupled to and controllable by the semiconductor-based active power converter. The method includes, in response to a determination that there is a fault within the electrical power system, operating in a fault mode which includes: maintaining the semiconductor-based active power converter in a blocked configuration (optionally followed by maintaining the semiconductor-based active power converter in a crow-bar configuration); and subsequently causing the contactor to be opened.
H02H 7/22 - Circuits de protection de sécurité spécialement adaptés aux machines ou aux appareils électriques de types particuliers ou pour la protection sectionnelle de systèmes de câble ou de ligne, et effectuant une commutation automatique dans le cas d'un changement indésirable des conditions normales de travail pour appareillage de distribution, p. ex. système de barre omnibusCircuits de protection de sécurité spécialement adaptés aux machines ou aux appareils électriques de types particuliers ou pour la protection sectionnelle de systèmes de câble ou de ligne, et effectuant une commutation automatique dans le cas d'un changement indésirable des conditions normales de travail pour dispositifs de commutation
There is provided a propulsion machine 10 comprising a fluid duct defined by a wall 42, 44 and a moveable member 34, 36 with a sealing module 90, 90′ therebetween. The moveable member 34, 36 is moveable relative to the wall 42, 44. The sealing module 90, 90′ comprises a mounting structure 92 coupled to the moveable member 34, 36 and an extendable structure 94 having a sealing surface 96. A chamber 98 is defined between the mounting structure 92 and the extendable structure 94 throughout a travel of the extendable structure 94 relative to the mounting structure 92. The sealing module 90, 90′ is configured to receive a pressurized actuation fluid into the chamber 98 to load the sealing surface 96 against an opposing surface 43 of the wall 42, 44 to provide a seal with the opposing surface 43.
Apparatus for active remote detection of leaking hydrogen comprises (i) pressure tubing for enclosing a length of pipework from which leaking hydrogen is to be detected; (ii) a detection vessel containing a hydrogen sensor; and (iii) connecting tubing connecting the interior of the pressure tubing to the interior of the detection vessel.
G01M 3/28 - Examen de l'étanchéité des structures ou ouvrages vis-à-vis d'un fluide par utilisation d'un fluide ou en faisant le vide par mesure du taux de perte ou de gain d'un fluide, p. ex. avec des dispositifs réagissant à la pression, avec des indicateurs de débit pour tuyaux, câbles ou tubesExamen de l'étanchéité des structures ou ouvrages vis-à-vis d'un fluide par utilisation d'un fluide ou en faisant le vide par mesure du taux de perte ou de gain d'un fluide, p. ex. avec des dispositifs réagissant à la pression, avec des indicateurs de débit pour raccords ou joints d'étanchéité de tuyauxExamen de l'étanchéité des structures ou ouvrages vis-à-vis d'un fluide par utilisation d'un fluide ou en faisant le vide par mesure du taux de perte ou de gain d'un fluide, p. ex. avec des dispositifs réagissant à la pression, avec des indicateurs de débit pour soupapes
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 7/22 - Mesure de la température basée sur l'utilisation d'éléments électriques ou magnétiques directement sensibles à la chaleur utilisant des éléments résistifs l'élément étant une résistance non linéaire, p. ex. une thermistance
A robotic shaping and forming system comprising a plurality of opposing robotic arms, a work platform for supporting a non-planar workpiece and a computer system, the plurality of opposing robotic arms having multiple degrees of freedom and an end effector for holding a tool, and wherein at least one robotic arm being mounted on a radially extending rail, the computer system being connected to the plurality of robotic arms, the computer system controlling the movement of the robotic arms, so that at least a pair of robotic arms work together to shape and form a non-planar workpiece that is mounted upon a work platform.
A method of forming a plurality of components on a build-plate, the method comprising the steps of providing a forming system comprising a plurality of high-energy beam generators mounted on a common support and configured to direct energy onto the build-plate; the build-plate comprising at least one segregation zones segregating the build plate into a plurality of build zones, each build zone associated with one of the high-energy beam generators; forming within each build zone at least one datum from material melted by energy from its associated high-energy beam generator; and forming within each build zone at least one component from material melted by energy from its associated high-energy beam generator; cutting or machining the build plate at the segregation zone or segregation zones to separate the build zones.
A turbine shroud assembly for use with a gas turbine engine includes a first shroud segment, a second shroud segment, and a damping strip seal assembly. The first shroud segment has a first carrier segment arranged circumferentially at least partway around a central axis and a first blade track segment supported by the first carrier segment. The second shroud segment is arranged circumferentially adjacent the first shroud segment. The damping strip seal assembly includes a body segment and a damping segment that extends along a curvilinear path.
A gas turbine includes an engine core with a turbine, a compressor, a combustor 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. A method of operating the turbine includes transferring heat from the oil to the fuel so as to lower the fuel viscosity to a value of less than or equal to 0.58 mm2/s 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 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
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn combustor having 14-22 fuel spray nozzles or 2-6 fuel spray nozzles per unit engine core size. A thrust nvPM emissions index ratio is
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn combustor having 14-22 fuel spray nozzles or 2-6 fuel spray nozzles per unit engine core size. A thrust nvPM emissions index ratio is
EI
maxTO
/
F
maxTO
EI
idle
/
F
i
d
l
e
.
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn combustor having 14-22 fuel spray nozzles or 2-6 fuel spray nozzles per unit engine core size. A thrust nvPM emissions index ratio is
EI
maxTO
/
F
maxTO
EI
idle
/
F
i
d
l
e
.
EIidle is the nvPM emissions index in mg/kg of the gas turbine engine operating at around 7% available thrust for given operating conditions. EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine operating at around 100% available thrust for the given operating conditions. FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN. Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN. The thrust nvPM emissions index ratio is greater than 0.09. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel to the fuel spray nozzles.
A hybrid transmission component includes a composite tube defining a central axis along its length and extending between a first end and a second end opposite to the first end. The composite tube includes a tube inner surface extending circumferentially about the central axis and a tube outer surface that is radially spaced apart from the tube inner surface with respect to the central axis. The composite tube further includes a first tube axial end surface extending between the tube inner surface and the tube outer surface at the first end and a second tube axial end surface extending between the tube inner surface and the tube outer surface at the second end. The tube inner surface comprises a wedge portion disposed at the first end.
F16D 1/08 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour montage d'un organe sur un arbre ou à l'extrémité d'un arbre avec moyeu de serrageAccouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour montage d'un organe sur un arbre ou à l'extrémité d'un arbre avec moyeu et clavette longitudinale
A hydrogen fuel delivery system (300) comprises a fuel line (312) having an inlet (315) and an outlet (316), a liquid fuel pump (307) configured to provide a flow of liquid hydrogen fuel from a hydrogen fuel storage tank (308) to the fuel line inlet (315), a heat exchanger (306) having first and second fluid paths (313, 314), the fuel line (312) passing through the first fluid path (313), a pre-heater line (317) having an inlet (318) connected to the fuel line (312) between the fuel line inlet (315) and the heat exchanger (306), the pre-heater line (317) comprising a first control valve (301) and a burner (305) between the pre-heater line inlet (318) and the heat exchanger (306), the pre-heater line (317) passing through the second fluid path (314) of the heat exchanger (306) towards a pre-heater line outlet (319), a second control valve (302) in the fuel line (312) between the heat exchanger (306) and the fuel line outlet (316), a first temperature sensor (321) configured to measure a first fuel temperature (T1) in the fuel line (312) between the heat exchanger (306) and the second control valve (302), and a control system (400) configured provide a first control signal (CV1) to control operation of the first control valve (301) dependent on an input target temperature (T1Target) compared to the first fuel temperature (T1) and on a measure of fuel flow (mbfuel) through the pre-heater line (317).
B64D 37/30 - Circuits de carburant pour carburants particuliers
B64D 37/34 - Conditionnement du carburant, p. ex. réchauffage
F17C 9/02 - Procédés ou appareils pour vider les gaz liquéfiés ou solidifiés contenus dans des récipients non sous pression avec changement d'état, p. ex. vaporisation
A current limiting device and an electrical power system including a current limiting device are provided. The current limiting device includes: a primary current path extending between a first node N1 and a second node N2 and having a primary current limiter connected therein, the primary current limiter including at least one JFET 101, 1011-N; and a secondary current path extending between the first node N1 and the second node N2 in parallel with the primary current path, the secondary current path including a Transient Voltage Suppressor (TVS). The primary current limiter is configured so that a voltage drop across the primary current limiter increases as a current flowing through the primary current path increases. When the current flowing through the primary current path passes a threshold, the voltage drop across the primary current limiter passes a breakdown voltage of the TVS.
A current limiting device and an electrical power system including a current limiting device are described. The current limiting device includes: an integer number, N, of JFETs 1011-N, each JFET of the N JFETs having a source terminal (S), a drain terminal (D) and a gate terminal (G). N≥2. Each of the N JFETs 1011-N has an index n=(1, . . . , N). For n=(1, . . . , N−1), the source terminal of the nth JFET is connected to the drain terminal of the (n+1)th JFET. The source terminal of the Nth JFET 101N is connected to the gate terminal of each of the N JFETs 1011-N.
H02H 9/02 - Circuits de protection de sécurité pour limiter l'excès de courant ou de tension sans déconnexion sensibles à un excès de courant
H02H 3/02 - Circuits de protection de sécurité pour déconnexion automatique due directement à un changement indésirable des conditions électriques normales de travail avec ou sans reconnexion Détails
H02H 3/087 - Circuits de protection de sécurité pour déconnexion automatique due directement à un changement indésirable des conditions électriques normales de travail avec ou sans reconnexion sensibles à une surcharge pour des systèmes à courant continu
A gas turbine engine for an aircraft includes a fan system having a reverse travelling wave first flap mode, Fan RTW, and including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode comprising a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first reverse whirl rotor dynamic mode, Rotor RW, and a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A backward whirl frequency margin of:
A gas turbine engine for an aircraft includes a fan system having a reverse travelling wave first flap mode, Fan RTW, and including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode comprising a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first reverse whirl rotor dynamic mode, Rotor RW, and a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A backward whirl frequency margin of:
the
lowest
frequency
of
either
mode
Fan
RTW
or
Rotor
RW
at
the
MTO
speed
the
MTO
speed
A gas turbine engine for an aircraft includes a fan system having a reverse travelling wave first flap mode, Fan RTW, and including a fan located upstream of the engine core; a fan shaft; and a front engine structure arranged to support the fan shaft and having a front engine structure nodding mode comprising a pair of modes at similar, but not equal, natural frequencies in orthogonal directions; and a gearbox. An LP rotor system including the fan system and a gearbox output shaft arranged to drive the fan shaft has a first reverse whirl rotor dynamic mode, Rotor RW, and a first forward whirl rotor dynamic mode, 1FW. The engine has a maximum take-off speed, MTO. A backward whirl frequency margin of:
the
lowest
frequency
of
either
mode
Fan
RTW
or
Rotor
RW
at
the
MTO
speed
the
MTO
speed
may be in the range from 15 to 50%.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
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
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
F04D 25/02 - Ensembles comprenant des pompes et leurs moyens d'entraînement
F04D 29/40 - Carters d'enveloppeTubulures pour le fluide énergétique
A method of training a conditional Generative Adversarial Network (cGAN) is disclosed. The cGAN has a generator and a discriminator. The method comprises obtaining a collection of images of components having a plurality of input parameter values defining component design properties and a plurality of output parameter values defining component performance attributes. For each image, it is determined whether the image represents a feasible design or an infeasible design. The images are then categorised into categories representing feasible designs into a plurality of feasible categories using a Pareto front ranking process, in which each one of the component performance attributes is defines a corresponding objective function, and the images representing infeasible designs are categorised into an infeasible category. The cGAN is then using the images and their categorisation.
G06V 10/774 - Génération d'ensembles de motifs de formationTraitement des caractéristiques d’images ou de vidéos dans les espaces de caractéristiquesDispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant l’intégration et la réduction de données, p. ex. analyse en composantes principales [PCA] ou analyse en composantes indépendantes [ ICA] ou cartes auto-organisatrices [SOM]Séparation aveugle de source méthodes de Bootstrap, p. ex. "bagging” ou “boosting”
G06V 10/82 - Dispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant les réseaux neuronaux
34.
EFFICIENT GAS TURBINE ENGINE INSTALLATION AND OPERATION
A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. 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.
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
F02C 3/02 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant la pression des gaz d'échappement dans un échangeur de pression pour comprimer l'air comburant
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 3/107 - 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
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
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
F04D 29/68 - Lutte contre la cavitation, les tourbillons, le bruit, les vibrations ou phénomènes analoguesÉquilibrage en agissant sur les couches limites
A method of training a conditional Generative Adversarial Network (cGAN) is disclosed. The cGAN has a generator and a discriminator. The method comprises obtaining a collection of images of components, each image having a physical parameter value relating to the component associated therewith, for each one of the collection of images of components, embedding a plurality of graphical encodings into the image that encode the associated physical parameter value, and training the cGAN using the collection of images of components with their embedded plurality of graphical encodings.
A seal structure for a gas turbine engine. The seal structure comprises a seal element, a retainer having an outer transversal diameter along a transversal axis and engaging the seal element. The retainer is connected to a supporting structure and comprises an external upper surface having a central convex portion and a peripheral curved portion.
The present disclosure relates to a method for recovering magnet material from a samarium cobalt, SmCo, magnet, the method comprising: initiating a hydrogen decrepitation process within a reaction vessel, wherein the hydrogen decrepitation process comprises: increasing a concentration of hydrogen in the reaction vessel, and maintaining the reaction vessel at either: a temperature of less than 70° C. and at a pressure of more than 10 bar, or at a temperature of more than 70° C. and at a pressure of less than 5 bar, to cause hydrogen decrepitation of the SmCo magnet disposed in the reaction vessel and produce SmCo-hydride material; and initiating a degasification process within a degasification vessel, wherein the degasification process comprises: removing gas from the degasification vessel and maintaining the degasification vessel at a temperature within a range of 150° C. to 300° C., to de-gas SmCo-hydride material disposed in the degasification vessel and produce SmCo material.
A cold plate for transferring heat from a device to a liquid coolant is shown. The cold plate comprises a plurality of coolant flow channels extending from a common inlet to a common outlet. The plurality of coolant flow channels is defined by a triply periodic minimal surface comprising a plurality of cells. Each one of the plurality of cells has an associated set of geometric parameters that are dependent upon the disposition of the cell in relation to the common inlet and the common outlet to produce a non-uniform lattice structure.
A cooling ring for a combustor system having an inner wall, an outer wall spaced apart from the inner wall, and a discharge nozzle disposed downstream of the inner wall includes an upstream portion disposed adjacent to the outer wall, a downstream portion spaced apart from the upstream portion, and a middle portion connecting the upstream portion to the downstream portion. The middle portion includes a plurality of first apertures and a plurality of second apertures. Each first aperture extends from a first inner surface portion to an outer surface portion of the middle portion along a first aperture axis and is configured to supply a cooling fluid to a cavity. Each second aperture extends from the second inner surface portion to the outer surface portion along a second aperture axis and is configured to supply the cooling fluid to the discharge nozzles.
Apparatus for heating a flow of hydrogen includes (i) a tank having a tank wall and first region containing a volume of incompressible fluid, the tank wall having a fluid input and a fluid output therein for establishing a flow of the incompressible fluid through the first region, (ii) an input conduit system having one or more outputs within the first region and (iii) an output conduit coupling a second region of the tank configured to collect gaseous hydrogen in operation of the apparatus to an apparatus output, the tank wall of the tank otherwise being closed. In operation of the apparatus, heat lost from the volume of incompressible fluid to hydrogen input to the apparatus is replaced by heat within incompressible fluid provided to the fluid input, thus maintaining the temperature of the body of incompressible fluid.
A composite gear includes an annular metallic gear hub including a bore and defining a central axis. The composite gear includes a composite web disposed circumferentially around the annular metallic gear hub with respect to the central axis. It includes a metallic gear rim disposed circumferentially around the composite web with respect to the central axis. The metallic gear rim includes an annular portion disposed adjacent to the composite web and a plurality of gear teeth angularly spaced apart from each other and extending outwardly from the annular portion distal to the composite web. The composite gear includes an interface disposed between the metallic gear rim and the composite web. The composite web includes a plurality of fibres inclined against each other by an angle from 30 to 60 degrees in order to form a mesh or woven fabric.
A pressure relief device for a gaseous fuel including a cylinder having a fuel inlet in a wall, an outlet at an open first end and an end wall at a closed opposing second end; a piston disposed within the cylinder, the piston having a first end adjacent the outlet of the cylinder defining an outlet orifice and a second opposing end slidably mounted within the cylinder between first and second end stops; and a spring mounted between the second end of the piston and the end wall of the cylinder, the spring arranged to bias the second end of the piston towards the second end stop, wherein movement of the piston along a longitudinal axis of the cylinder between the first and second end stops varies the outlet orifice to vary a flow of gaseous fuel between the fuel inlet and the outlet orifice.
F02M 21/02 - Appareils pour alimenter les moteurs en combustibles non liquides, p. ex. en combustibles gazeux stockés sous forme liquide en combustibles gazeux
F02M 49/02 - Appareils d'injection dans lesquels les pompes sont entraînées ou dont les injecteurs sont actionnés par la pression dans le cylindre moteur ou par contact du piston moteur utilisant la pression du cylindre, p. ex. pression de fin de compression
F02M 61/04 - Injecteurs de combustible non couverts dans les groupes ou comportant des clapets
EImaxTO,SAFEImaxTO,FFWf,maxTOf,maxTO is the mass flow rate of fuel provided to the plurality of fuel spray nozzles in kg/s when the gas turbine engine is operating at around 100% available thrust for the given operating conditions. The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine in kg/s is less than 4. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.
F02C 3/24 - 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 liquide aux température et pression normales
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
maxTO,SAFmaxTO,FFmaxTO,FF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
EIidleidle EImaxTOmaxTO FmaxTOmaxTO Fidleidle idle is the thrust of the gas turbine engine at around 7% available thrust in kN; the thrust nvPM emissions index ratio is greater than 0.09; and the gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
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 7/228 - 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
46.
MODULAR MECHANICAL TEST APPARATUS FOR CMC COMPONENT
Rolls-Royce High Temperature Composites, Inc. (USA)
Rolls-Royce North American Technologies, Inc. (USA)
Rolls-Royce plc (Royaume‑Uni)
Inventeur(s)
Erlitz, Kristopher
Bledsoe, Ronald Adolphus
Tran, Thomas
Baker, Jason David
Traudes, Daniel Jozef
Pattison, Stephen John
Abrégé
A test apparatus for a ceramic matrix composite (CMC) component. The test apparatus includes a first support member configured to mechanically support a CMC component from a first side. The CMC component includes a T-joint and a pinhole. The test apparatus includes a second support member configured to mechanically support the CMC component from a second side opposite the first side. The first support member and the second support member are configured to be forced toward each other to cause the CMC component to fail at the pinhole and not at the T-joint.
G01N 3/08 - Recherche des propriétés mécaniques des matériaux solides par application d'une contrainte mécanique par application d'efforts permanents de traction ou de compression
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in range of 14-22 or a number of fuel spray nozzles per unit engine core size in range 2 to 6. An MTO nvPM emissions index ratio is:
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in range of 14-22 or a number of fuel spray nozzles per unit engine core size in range 2 to 6. An MTO nvPM emissions index ratio is:
EI
maxTO
,
SAF
EI
maxTO
,
FF
×
W
f
,
maxTO
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in range of 14-22 or a number of fuel spray nozzles per unit engine core size in range 2 to 6. An MTO nvPM emissions index ratio is:
EI
maxTO
,
SAF
EI
maxTO
,
FF
×
W
f
,
maxTO
where: EImaxTO,SAF and EImaxTO,FF are respectively the nvPM emissions index in mg/kg when operating around 100% available thrust for given operating conditions if fuel provided to fuel spray nozzles includes either sustainable aviation fuel (SAF) or fossil-based hydrocarbon fuel; Wf,maxTO is mass flow rate of fuel provided to fuel spray nozzles in kg/s when gas turbine engine is operating at around 100% available thrust for given operating conditions. MTO nvPM emissions index is less than 2. The gas turbine engine is configured to provide fuel including SAF to fuel spray nozzles. Also disclosed is method of operating gas turbine engine.
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
EI
cruise
(
lean
)
,
SAF
EI
cruise
(
lean
)
,
FF
,
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
EI
cruise
(
lean
)
,
SAF
EI
cruise
(
lean
)
,
FF
,
where EIcruise(lean),SAF is
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
EI
cruise
(
lean
)
,
SAF
EI
cruise
(
lean
)
,
FF
,
where EIcruise(lean),SAF is
EI
maxTO
,
SAF
+
EI
climb
,
SAF
2
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
EI
cruise
(
lean
)
,
SAF
EI
cruise
(
lean
)
,
FF
,
where EIcruise(lean),SAF is
EI
maxTO
,
SAF
+
EI
climb
,
SAF
2
and EIcruise(lean),FF is
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
EI
cruise
(
lean
)
,
SAF
EI
cruise
(
lean
)
,
FF
,
where EIcruise(lean),SAF is
EI
maxTO
,
SAF
+
EI
climb
,
SAF
2
and EIcruise(lean),FF is
EI
maxTO
,
FF
+
EI
climb
,
FF
2
.
An aircraft gas turbine engine includes a combustor, with a combustion chamber and fuel spray nozzles to inject fuel into the chamber. The nozzles include first and second subsets. The first subset of nozzles is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:3 to 1:6. A lean cruise nvPM emissions index ratio is
EI
cruise
(
lean
)
,
SAF
EI
cruise
(
lean
)
,
FF
,
where EIcruise(lean),SAF is
EI
maxTO
,
SAF
+
EI
climb
,
SAF
2
and EIcruise(lean),FF is
EI
maxTO
,
FF
+
EI
climb
,
FF
2
.
EImaxTO,SAF is nvPM emissions index operating at 100% available thrust and EIclimb,SAF is the nvPM emissions index operating at 85% available thrust if fuel includes sustainable aviation fuel. EImaxTO,FF is the nvPM emissions index operating at 100% available thrust and EIclimb,FF is the nvPM emissions index at 85% available thrust if fuel is a fossil-based hydrocarbon. The lean cruise nvPM emissions index ratio is less than 1.
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises:
a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises:
a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is defined as:
E
I
idle
×
W
f
,
idle
E
I
maxTO
×
W
f
,
maxTO
where: EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg/s at around 100% available thrust for the given operating conditions. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed are methods of operating a gas turbine engine.
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:
maxTO,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:
EI
maxTO
/
F
maxTO
EI
idle
/
F
idle
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:
EI
maxTO
/
F
maxTO
EI
idle
/
F
idle
where: EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN for the given operating conditions; and Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN for the given operating conditions. The thrust nvPM emissions index ratio is greater than 0.001. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
An inspection system for inspecting a surface includes a macromanipulator including an inspection end configured to be disposed proximal to the surface. The inspection system further includes a micromanipulator coupled to the inspection end. The. micromanipulator includes a housing, a pair of guide rails at least partially disposed within and fixedly coupled to the housing, a probe support slidably coupled to the pair of guide rails, an actuating arm disposed within the housing and coupled to the probe support via a scotch yoke mechanism, and an actuating mechanism configured to rotate the actuating arm relative to the housing. The inspection system further includes a probe coupled to the probe support for inspecting the surface.
An annular brush seal includes a bristle pack formed of bristles and a bristle pack support, which supports the bristle pack. The bristle pack has an upstream face, a downstream face and a sealing face. The sealing face includes an upstream portion and a downstream portion that is adjacent to the upstream portion. A length of at least a fraction of the bristles of the upstream portion is shorter than a length of the bristles of the downstream portion to form a circumferentially repeating pattern at an interface between the upstream portion of the sealing surface and the upstream surface.
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A fuel-flow nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A fuel-flow nvPM emissions index ratio is defined as:
EI
idle
×
W
f
,
idle
EI
max
TO
×
W
f
,
maxTO
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A fuel-flow nvPM emissions index ratio is defined as:
EI
idle
×
W
f
,
idle
EI
max
TO
×
W
f
,
maxTO
where: EIidle is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; and EImaxTO is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; Wf,idle is the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% available thrust for the given operating conditions; and Wf,maxTO is the rate of fuel flow to the fuel spray nozzles in kg/s at around 100% available thrust for the given operating conditions. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.08. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
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
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:
EI
maxTO
,
SAF
EI
maxTO
,
FF
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:3 to 1:6. A MTO nvPM emissions index ratio is defined as:
EI
maxTO
,
SAF
EI
maxTO
,
FF
where: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed are methods of operating the gas turbine engine.
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:
EI
maxTO
/
F
maxTO
EI
idle
/
F
idle
A gas turbine engine for an aircraft is disclosed. The gas turbine engine comprises: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. A thrust nvPM emissions index ratio is defined as:
EI
maxTO
/
F
maxTO
EI
idle
/
F
idle
where: EIidle is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; EImaxTO is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions; FmaxTO is the thrust of the gas turbine engine at around 100% available thrust in kN for the given operating conditions; and Fidle is the thrust of the gas turbine engine at around 7% available thrust in kN for the given operating conditions. The thrust nvPM emissions index ratio is greater than 0.001. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
A gas turbine engine for an aircraft. The gas turbine engine comprising: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is defined as:
A gas turbine engine for an aircraft. The gas turbine engine comprising: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is defined as:
EI
maxTO
,
SAF
EI
maxTO
,
FF
×
W
f
,
maxTO
A gas turbine engine for an aircraft. The gas turbine engine comprising: a combustor, comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor is operable in a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is defined as:
EI
maxTO
,
SAF
EI
maxTO
,
FF
×
W
f
,
maxTO
where: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); EImaxTO,FF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and Wf,maxTO is the mass flow rate of fuel provided to the plurality of fuel spray nozzles in kg/s when the gas turbine engine is operating at around 100% available thrust for the given operating conditions. The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine in kg/s is less than 2. The gas turbine engine is configured to provide fuel comprising a SAF to the plurality of fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
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 7/22 - Systèmes d'alimentation en combustible
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
59.
Emissions of non-volatile particulate matter from gas turbine engines combusting sustainable aviation fuel and fossil-based hydrocarbon fuel
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as:
maxTO,FF are respectively the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles includes a sustainable aviation fuel (SAF) or is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel including a SAF to the fuel spray nozzles. Also disclosed is a method of operating a gas turbine engine.
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An nvPM emissions index ratio is defined as:
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An nvPM emissions index ratio is defined as:
EI
idle
×
W
f
,
idle
EI
maxTO
×
W
f
,
maxTO
A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An nvPM emissions index ratio is defined as:
EI
idle
×
W
f
,
idle
EI
maxTO
×
W
f
,
maxTO
where: EIidle and EImaxTO are respectively the nvPM emissions index in mg/kg of gas turbine engine if operating at around 7% or 100% available thrust for given operating conditions; Wf,idle and Wf,maxTO are respectively the rate of fuel flow to fuel spray nozzles in kg/s at around 7% or 100% available thrust for given operating conditions. The fuel-flow nvPM emissions index ratio is less than 0.08. The gas turbine engine is configured to provide fuel including sustainable aviation fuel to fuel spray nozzles. Also disclosed is a method of operating the gas turbine engine.
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn (RQL) combustor having 14-22 fuel spray nozzles or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A first idle-MTO nvPM emissions index ratio
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn (RQL) combustor having 14-22 fuel spray nozzles or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A first idle-MTO nvPM emissions index ratio
EI
idle
EI
maxTO
.
A gas turbine engine for an aircraft includes a rich burn, quick quench, lean burn (RQL) combustor having 14-22 fuel spray nozzles or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A first idle-MTO nvPM emissions index ratio
EI
idle
EI
maxTO
.
EIidle is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions. EImaxTO is the nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions. The first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 0.8. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.
A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater flow rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is
A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater flow rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is
EI
maxTO
,
SAF
EI
maxTO
,
FF
×
W
f
,
maxTO
.
A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater flow rate than each of the second subset. A ratio of nozzles in the first subset to the second subset is 1:2 to 1:5. An MTO nvPM emissions index ratio-modified fuel flow is
EI
maxTO
,
SAF
EI
maxTO
,
FF
×
W
f
,
maxTO
.
EImaxTO,SAF is nvPM emissions index in mg/kg of the engine operating at around 100% available thrust with sustainable aviation fuel. EImaxTO,FF is nvPM emissions index in mg/kg of the engine operating at around 100% available thrust with fossil-based hydrocarbon. Wf,maxTO is mass flow rate of fuel to the nozzles in kg/s operating at around 100% available thrust. The MTO nvPM emissions index ratio-modified fuel flow in kg/s is less than 2.
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 7/22 - Systèmes d'alimentation en combustible
A fuel system for a hydrogen fuelled aircraft propulsion system comprises a fuel tank configured to store hydrogen fuel, a main fuel conduit configured to provide fuel to a combustor of a gas turbine engine, a fuel heater comprising a catalytic combustor configured to catalytically combust a portion of the hydrogen fuel prior to delivery to the combustor and a heat exchanger configured to exchange heat between exhaust gases from the fuel heater and hydrogen fuel in the fuel conduit.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
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 for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A first idle-MTO nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. A first idle-MTO nvPM emissions index ratio is defined as:
EIidle/EImaxTO
where: EIidle is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine if operating at around 7% available thrust for given operating conditions; and EImaxTO is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine if operating at around 100% available thrust for the given operating conditions. The first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 0.8. The gas turbine engine is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles. A method of operating the gas turbine engine is also disclosed.
A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each the second subset of fuel spray nozzles. A a ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:
A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each the second subset of fuel spray nozzles. A a ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:
EI
idle
EI
maxTO
A gas turbine engine has a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. A combustor is operable with the first subset of fuel spray nozzles supplied with fuel at a greater fuel flow rate than each the second subset of fuel spray nozzles. A a ratio of the first subset of fuel spray nozzles to the second subset of fuel spray nozzles is 1:2 to 1:5. A first idle-MTO nvPM emissions index ratio is:
EI
idle
EI
maxTO
where: EIidle is the system loss corrected nvPM emissions index in mg/kg operating at around 7% available thrust; and EImaxTO is the system loss corrected nvPM emissions index in mg/kg operating at around 100% available thrust; and the first idle-MTO nvPM emissions index ratio of the gas turbine engine is less than 60. A sustainable aviation fuel (SAF) can be provided to the fuel spray nozzles.
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as:
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as:
EI
maxTO
,
SAF
EI
maxTO
,
FF
A gas turbine engine for an aircraft. The gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as:
EI
maxTO
,
SAF
EI
maxTO
,
FF
where: EImaxTO,SAF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EImaxTO,FF is the system loss corrected nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel comprising a SAF to the fuel spray nozzles. Also disclosed is a method of operating a gas turbine engine.
A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset from 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset from 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
EI
idle
×
W
f
,
idle
RI
maxTO
×
W
f
,
maxTO
.
A gas turbine engine includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The nozzles include a first and second subset. Each of the nozzles of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of nozzles in the first subset to nozzles in the second subset from 1:3 to 1:6. A fuel-flow nvPM emissions index ratio is
EI
idle
×
W
f
,
idle
RI
maxTO
×
W
f
,
maxTO
.
EIidle and EImaxTO are the system loss corrected nvPM emissions index in mg/kg operating at around 7% and 100% available thrust for the given operating conditions, respectively. Wf,idle and Wf,maxTO are the rate of fuel flow to the fuel spray nozzles in kg/s at around 7% and 100% available thrust for the given operating conditions, respectively. The fuel-flow nvPM emissions index ratio of the gas turbine engine is less than 0.3.
A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is
A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is
EI
maxTO
,
SAF
EI
maxTO
,
FF
.
A gas turbine engine for aircraft includes a combustor with a combustion chamber and fuel spray nozzles to inject fuel into the combustion chamber. The fuel spray nozzles include a first subset and a second subset of nozzles. Each of the first subset is supplied with fuel at a greater rate than each of the second subset. A ratio of the first subset to the second subset is 1:2 to 1:5. A MTO nvPM emissions index ratio is
EI
maxTO
,
SAF
EI
maxTO
,
FF
.
EImaxTO,SAF is nvPM emissions index in mg/kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles includes sustainable aviation fuel. EImaxTO,FF IS nvPM emissions index in mg/kg of the engine when operating at around 100% available thrust if fuel provided to the fuel spray nozzles is fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio is less than 1.
A method for modifying the mechanical and surface properties of a component. The method involves: removably attaching a component to at least one component support that is located within a vibratory trough of a component treatment apparatus, the at least one component support being a support shaft upon which the component is removably mountable within the vibratory trough; supplying the vibratory trough with treatment media; moving the component support or the vibratory trough so that the component is immersed into the treatment media; vibrating the vibratory trough to provide a substantially uniform surface treatment of the component, the vibratory trough being movable by at least one trough vibrating mechanism whose actuation is controlled by a controller in response to signals received from at least one sensor located on or within the vibratory trough; removing the component from the treatment media; and detaching the component from the component support.
B24B 31/06 - Machines ou dispositifs pour polir ou travailler par abrasion des surfaces "au tonneau", ou au moyen d'autres appareils, dans lesquels les pièces à travailler ou les produits abrasifs sont libresAccessoires à cet effet impliquant l'emploi de récipients oscillants ou vibrants
B24B 19/14 - Machines ou dispositifs conçus spécialement pour une opération particulière de meulage non couverte par d'autres groupes principaux pour meuler des aubes de turbine, des pales d'hélice ou similaires
70.
DEVICE AND APPARATUS FOR HEATING A FLOW OF HYDROGEN
A heating device comprises a thermally-conductive outer tube, and a combustor arranged to combust air and gaseous hydrogen input to the device and introduce resulting combustion products into a first end of the outer tube. The device includes an igniter arranged to ignite the air and gaseous hydrogen. Combustion products produced by the combustor flow from the combustor to an output of the device along a flow path such that they are in contact with the internal surface of the outer tube over at least a portion of the flow path. Apparatus for heating a flow of hydrogen comprises a conduit for conducting the flow of hydrogen from a principal input port to a principal output port, the apparatus further comprising the heating device located at least partially within the conduit between the principal input port and the principal output port.
F23K 5/00 - Alimentation en d'autres combustibles ou distribution d'autres combustibles pour les appareils à combustion
B64D 37/30 - Circuits de carburant pour carburants particuliers
B64D 37/34 - Conditionnement du carburant, p. ex. réchauffage
F23D 14/02 - Brûleurs à gaz avec prémélangeurs, c.-à-d. dans lesquels le combustible gazeux est mélangé à l'air de combustion en amont de la zone de combustion
71.
TRAINING AND USING CONDITIONAL GENERATIVE ADVERSARIAL NETWORKS
Methods and techniques are disclosed for training Conditional Generative Adversarial Networks using images of acceptable and non-acceptable components. Methods and techniques are also disclosed for using Conditional Generative Adversarial Networks to assess the acceptability or non-acceptability of a component using an image thereof.
An assessment of power electronics converters for electrical power systems. Example embodiments include a method of assessing a power electronics converter in an electrical propulsion system, the power electronics converter connected between a DC power supply and an AC electrical machine connected to a rotatable mechanical load, the method comprising: i) operating the converter to drive the electrical machine to apply a torque to the mechanical load; ii) recording a first set of electromagnetic emissions from the converter; iii) comparing the first set of electromagnetic emissions with a second set of electromagnetic emissions to determine a difference between the first and second set of electromagnetic emissions; and iv) providing an assessment of the converter based on the difference between the first and second sets of electromagnetic emissions.
G01R 31/42 - Tests d'alimentation d'alimentations en courant alternatif
B60L 3/00 - Dispositifs électriques de sécurité sur véhicules propulsés électriquementContrôle des paramètres de fonctionnement, p. ex. de la vitesse, de la décélération ou de la consommation d’énergie
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
This disclosure relates to detection of a bearing fault in an electrical propulsion system by shaft voltage measurement. Example embodiments include a method of monitoring an electrical propulsion system (100) comprising a first power electronics converter (1021) connected between a first DC power source (1031) and a first AC electric motor (1011) connected to drive a first shaft (1051) mounted to a first bearing (1061), the method comprising: operating the first power electronics converter (1021) to drive the first AC electric motor (1011); contacting a first electrical contact (1091) with the first shaft (1051); measuring a first voltage between the first electrical contact (1091) and a first common reference; comparing the first voltage with a predetermined threshold; and determining a fault in the first bearing (1061) if the measured first voltage exceeds the predetermined threshold.
F16C 19/52 - Paliers à contact de roulement pour mouvement de rotation exclusivement avec dispositifs affectés par des conditions anormales ou indésirables
74.
HYDROGEN FUELLED AIRCRAFT PROPULSION SYSTEM OPERATING METHOD
A method of operating a fuel system of a hydrogen fuelled aircraft propulsion system includes exposing one or more fuel system components to a hydrogen embrittlement inhibiting gas including one of oxygen and carbon monoxide at a concentration above an embrittlement inhibition concentration.
A method of removing a coating system from a component that is coated with the coating system. The method involves: (a) immersing the component in a caustic solution; (b) maintaining the component in the caustic solution at atmospheric pressure for a time ≤1.5 hours at a temperature ≥150° C. and ≤250° C.; (c) removing the component; (d) rinsing the component in water; (e) water jet blasting the component to remove the ceramic top coat layer and any thermally-grown oxide; (f) immersing the component in an acid solution; (g) ultra-high pressure water jetting the component; (h) aluminising the component to convert any diffused Pt within the bond coat layer to Pt—Al; (i) acid stripping and grit blasting the component; (j) immersing the component in a solution of nitric acid and/or sulphamic acid; and (k) ultra-high pressure water jetting the component to remove any Pt—Al.
F01D 5/00 - AubesOrganes de support des aubesDispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
B24C 1/04 - Méthodes d'utilisation de jet abrasif en vue d'effectuer un travail déterminéUtilisation d'équipements auxiliaires liés à ces méthodes pour travailler uniquement certaines parties déterminées, p. ex. pour graver la pierre ou le verre
C23G 1/14 - Nettoyage ou décapage de matériaux métalliques au moyen de solutions ou de sels fondus avec des solutions alcalines
76.
INSPECTION ASSEMBLY, SYSTEM, AND METHOD FOR A GAS TURBINE ENGINE
An inspection system for a gas turbine engine having an engine core, a casing, and a static aerodynamic fairing includes an inspection assembly including: an elongate member extending along a longitudinal axis from a first end to a second end and configured to be at least partially inserted through the static aerodynamic fairing into a portion of the engine core radially inboard of the static aerodynamic fairing; a coupler disposed proximal to the first end and configured to removably couple the elongate member to the casing; and a sensor member pivotally coupled to the elongate member at the second end and including a sensor. The sensor member is pivotable between an insertion position and a sensing position. The inspection system further includes: a guidance member located within the engine core and including a guidance surface; and a locking mechanism operatively coupled to the sensor member.
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
The disclosure relates to synchronising measurement data for engine health monitoring when using more than one engine health monitoring system. Example embodiments include a method of synchronising sensor data from first and second engine health monitoring, EHM, systems installed on an engine, each EHM system having a plurality of sensors configured to measure parameters of the engine, the method comprising: acquiring first and second sensor data from the plurality of sensors at the first and second EHM systems; logging the first sensor data against a first time signal from the first EHM systems and the second sensor data against a second time signal from the second EHM system; acquiring a side channel signal from a side channel sensor mounted to a cable or connector connected to the first EHM system; deriving a timing signal from the side channel signal; determining a timing difference between the first and second time signals from the derived timing signal; and adjusting the second time signal for the second sensor data to align the second sensor data in time with the first sensor data.
The disclosure relates to synchronising measurement data for engine health monitoring when using more than one engine health monitoring system. Example embodiments include a method of synchronising sensor data from first and second engine health monitoring, EHM, systems installed on an engine, each EHM system having a plurality of sensors configured to measure parameters of the engine, the method comprising:
The disclosure relates to synchronising measurement data for engine health monitoring when using more than one engine health monitoring system. Example embodiments include a method of synchronising sensor data from first and second engine health monitoring, EHM, systems installed on an engine, each EHM system having a plurality of sensors configured to measure parameters of the engine, the method comprising:
acquiring first and second sensor data from the plurality of sensors at the first and second EHM systems, the first sensor data including a measured engine speed; logging the first sensor data against a first time signal from the first EHM system and the second sensor data against a second time signal from the second EHM system; acquiring a synchronisation signal from a synchronisation sensor mounted to a part of the engine; deriving an engine speed from a frequency of the synchronisation signal; determining a timing difference between the first and second time signals from the derived engine speed and the measured engine speed; and adjusting the second time signal for the second sensor data to align the second sensor data in time with the first sensor data.
A gas turbine engine includes an engine core including a turbine, compressor, combustor to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan to drive the fan; an oil loop system to supply oil to the gearbox; and a heat exchange system with an air-oil heat exchanger through which the oil flows; a fuel-oil heat exchanger through which the oil and the fuel flow; and an air valve to control a flow rate of air through the air-oil heat exchanger. A method of operating the gas turbine engine includes determining at least one fuel characteristic; and controlling the air valve based on the fuel characteristic so as to adjust the flow rate of air through the air-oil heat exchanger.
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 for an aircraft comprises an engine core. The engine comprises a fan located upstream of the engine core. The engine comprises a nacelle surrounding the fan and the engine core and defining a bypass duct, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine comprises a plurality of actuators and a fuel supply system arranged to supply fuel for combustion in the combustor, and to supply fuel to fueldraulically drive at least one actuator. The fuel comprises at least 25% SAF by volume, and the fuel supply system is arranged such that a peak differential pressure of the fuel across the at least one fueldraulic actuator during cruise conditions is at least 2400 kPa.
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
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
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 7/22 - Systèmes d'alimentation en combustible
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
F02C 9/26 - Commande de l'alimentation en combustible
81.
COMBUSTORS AND METHODS OF MITIGATING THERMOACOUSTIC INSTABILITIES IN A GAS FLOW
There is provided a heat engine 10 comprising a combustor 120 and a controller 290. The combustor 120 includes a fuel injection port 128 and a pilot 130. The controller 290 is configured to mitigate thermoacoustic instabilities in a gas flow A, B within the heat engine 10 by selectively energising 320 the pilot 130 to ignite fuel discharged from the fuel injection port 128. There is also provided a method 300 of mitigating thermoacoustic instabilities in a gas flow A, B within a heat engine 10 comprising a combustor 120 including a fuel injection port 128 and a pilot 130. The method 300 comprises selectively energising 320 the pilot 130 to ignite fuel discharged by the fuel injection port 128.
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
Gearboxes for aircraft gas turbine engines, in particular arrangements for journal bearings such gearboxes, and related methods of operating such gearboxes and gas turbine engines, including a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F16C 17/03 - Paliers à contact lisse pour mouvement de rotation exclusivement pour charges radiales uniquement avec segments supportés obliquement, p. ex. paliers Michell
F16H 1/28 - Transmissions à engrenages pour transmettre un mouvement rotatif avec engrenages à mouvement orbital
F16H 57/02 - Boîtes de vitessesMontage de la transmission à l'intérieur
F16H 57/08 - Parties constitutives générales des transmissions des transmissions à organes à mouvement orbital
A cabin blower control system for controlling a cabin blower system includes a drive unit and a controller including a memory and a processor. The memory stores a predetermined blower dataset for a blower unit of the cabin blower system. The blower unit includes at least one compressor. The processor performs the following steps: receive a desired mass flow rate of the outlet airflow to meet a current loading on the cabin blower system; receive an inlet temperature and an inlet pressure of the inlet airflow, a compressor speed of the compressor, and a current operating condition of the blower unit; determine an estimated power consumption, a current power consumption, and an estimated operating condition of the blower unit; determine a desired speed of the compressor to operate the compressor at the desired speed.
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
An electrical power system and an aircraft including an electrical power system are provided. The electrical power system includes: an electrical power source; a DC electrical network; a power converter including at least one input terminal and first and second DC output terminals, the at least one input terminal connected to the electrical power source, the first and second DC output terminals connected to the DC electrical network; a DC link capacitor connected between the first and second DC output terminals; a power semiconductor switch connected in series with the DC link capacitor between the first and second DC output terminals; and a control unit configured to respond to a fault in the DC electrical network by opening the power semiconductor switch, whereby discharge of the DC link capacitor is prevented.
B64D 31/16 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques
H02H 1/04 - Dispositions pour prévenir la réponse à des conditions transitoires anormales, p. ex. à la foudre
H02H 3/24 - Circuits de protection de sécurité pour déconnexion automatique due directement à un changement indésirable des conditions électriques normales de travail avec ou sans reconnexion sensibles à une baisse ou un manque de tension
H02J 1/02 - Dispositions pour réduire les harmoniques ou les ondulations
H02M 1/14 - Dispositions de réduction des ondulations d'une entrée ou d'une sortie en courant continu
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
H02M 3/335 - Transformation d'une puissance d'entrée en courant continu en une puissance de sortie en courant continu avec transformation intermédiaire en courant alternatif par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrodes de commande pour produire le courant alternatif intermédiaire 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 7/219 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continu sans possibilité de réversibilité par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs dans une configuration en pont
A hydraulic system for a hydrogen fueled gas turbine engine. The hydraulic system comprises a closed-loop hydraulic fluid conduit (232), a hydrogen-hydraulic fluid heat exchanger (230) configured to exchange heat between cryogenic hydrogen fuel and hydraulic fluid within the hydraulic fluid conduit (232), and a second fuel-hydraulic heat exchanger (518) configured to exchange heat between a second, non-hydrogen fuel and hydraulic fluid in the hydraulic fluid conduit (232).
F15B 15/14 - Dispositifs actionnés par fluides pour déplacer un organe d'une position à une autreTransmission associée à ces dispositifs caractérisés par la structure de l'ensemble moteur le moteur étant du type à cylindre droit
F02C 6/00 - Ensembles fonctionnels multiples de turbines à gazCombinaisons d'ensembles fonctionnels de turbines à gaz avec d'autres appareilsAdaptations d'ensembles fonctionnels de turbines à gaz à des applications particulières
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F02C 7/22 - Systèmes d'alimentation en combustible
A method includes operating a gas turbine engine. The gas turbine engine includes a staged combustor having an arrangement of fuel spray nozzles in which fuel flow is biased to a subset of the nozzles adjacent one or more ignitors during a re-light procedure. The method includes providing fuel to the combustor having an aromatic content of 10% or lower by volume. Also disclosed is a gas turbine engine.
A nozzle body for a fuel injector includes an outer air circuit, a fuel circuit, a dome air circuit, and a first prefilmer radially disposed between the outer air circuit and the dome air circuit. The first prefilmer is inclined to a central axis of the nozzle body by an inclination angle from 0 degree to 10 degrees, such that a fuel discharged by the fuel circuit at least partially impinges on the first prefilmer. An axial gap between a shroud throat of a shroud of the nozzle body and a trailing edge of the first prefilmer is between 1 mm and 2 mm, such that the fuel impinging on the first prefilmer is guided towards a dome air discharged by the dome air circuit.
F23R 3/14 - Aménagements de l'entrée d'air pour l'air primaire créant un tourbillon au moyen d'ailettes de tourbillonnement
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
F23R 3/30 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible comprenant des dispositifs destinés à prévaporiser le combustible
F02C 7/22 - Systèmes d'alimentation en combustible
A power electronics converter including: an input terminal; first and second DC output terminals; a branch having first and second semiconductor switches connected in series between the first and second DC output terminals, the input terminal connected to a node between the first and second semiconductor switches; a DC link capacitor connected between the first and second DC output terminals; a differential current sensor arranged to measure a differential current signal through the first or second DC terminals; and a controller configured to provide switching signals to each of the first and second switches of the power electronics converter, wherein the controller is further configured to detect a fault in a DC network connected between the first and second DC output terminals upon detection of a peak in an output from the differential current sensor.
H02H 7/125 - Circuits de protection de sécurité spécialement adaptés aux machines ou aux appareils électriques de types particuliers ou pour la protection sectionnelle de systèmes de câble ou de ligne, et effectuant une commutation automatique dans le cas d'un changement indésirable des conditions normales de travail pour convertisseursCircuits de protection de sécurité spécialement adaptés aux machines ou aux appareils électriques de types particuliers ou pour la protection sectionnelle de systèmes de câble ou de ligne, et effectuant une commutation automatique dans le cas d'un changement indésirable des conditions normales de travail pour redresseurs pour convertisseurs ou redresseurs statiques pour redresseurs
H02M 1/00 - Détails d'appareils pour transformation
H02M 1/08 - Circuits spécialement adaptés à la production d'une tension de commande pour les dispositifs à semi-conducteurs incorporés dans des convertisseurs statiques
H02M 7/25 - 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 agencés pour la marche en série, p. ex. pour la multiplication de la tension
89.
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
90.
SYSTEM AND METHOD FOR WET TREATMENT OF A COMPONENT
A system has a chamber configured to receive and at least partially enclose at least one gas turbine engine component, at least one component support configured to support the at least one gas turbine engine component, a plurality of tanks configured to store a corresponding plurality of fluids, an electrode disposed at least partially within the chamber, a power source, at least one fluid application device configured to apply a fluid, at least one delivery valve for selectively fluidly coupling the at least one fluid application device to the plurality of tanks, at least one port configured to collect the fluid applied by the at least one fluid application device, and at least one recovery valve for selectively fluidly coupling the at least one port to the plurality of tanks.
An electrical power system 10, comprising: a rotary electrical machine 12 configured to output AC; a diode-bridge rectifier 13 having an AC input connected to the electrical machine 12 and a DC output (DC+, DC−); an active filter circuit 14 comprising a plurality of power semiconductor switches 141L-H, 142L-H connected in a bridge configuration between first and second output terminals 14out, the first and second output terminals 14out connected to the DC output (DC+, DC−) of the diode-bridge rectifier 13; and a controller 15 configured to control a switching operation of the plurality of power semiconductor switches 141L-H, 142L-H of the 10 active filter circuit 14 to control an output voltage Vfilt across the first and second output terminals 14out of the active filter circuit 14.
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
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
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
93.
APPARATUS FOR SUPPORTING AT LEAST A PART OF AN ENGINE
Apparatus for supporting at least a part of an engine, the apparatus comprising: an engine stand configured to support at least a part of an engine; an engine cover defining an aperture; a dehumidifier coupled to the aperture of the engine cover, the dehumidifier being configured to remove water vapour from air within the engine cover.
F16M 5/00 - Bâtis pour machines, c.-à-d. moyens de tenir les moteurs ou machines sur leurs fondations
F24F 3/14 - Systèmes de conditionnement d'air dans lesquels l'air conditionné primaire est fourni par une ou plusieurs stations centrales aux blocs de distribution situés dans les pièces ou enceintes, blocs dans lesquels il peut subir un traitement secondaireAppareillage spécialement conçu pour de tels systèmes caractérisés par le traitement de l'air autrement que par chauffage et refroidissement par humidificationSystèmes de conditionnement d'air dans lesquels l'air conditionné primaire est fourni par une ou plusieurs stations centrales aux blocs de distribution situés dans les pièces ou enceintes, blocs dans lesquels il peut subir un traitement secondaireAppareillage spécialement conçu pour de tels systèmes caractérisés par le traitement de l'air autrement que par chauffage et refroidissement par déshumidification
F24F 13/10 - Organes de réglage de l'écoulement d'air, p. ex. persiennes, grilles, volets ou plaques directrices mobiles, p. ex. registres
F24F 13/22 - Moyens pour éviter la condensation ou pour évacuer le condensat
A method of analysing a spray of liquid from a nozzle comprises: receiving a fluorescent liquid composition comprising fluorophores of a first type and fluorophores of a second type, wherein the fluorophores of the first type are excitable by absorption of electromagnetic radiation in a first absorption wavelength band and are configured to emit electromagnetic radiation, following excitation, in a first emission wavelength band, and wherein the fluorophores of the second type are excitable by absorption of electromagnetic radiation in a second absorption wavelength band and are configured to emit electromagnetic radiation, following excitation, in a second emission wavelength band, wherein the first emission wavelength band overlaps with the second absorption wavelength band; ejecting the fluorescent liquid composition from the nozzle to generate a spray; projecting, within a sheet plane, a sheet of light through the spray, wherein the light comprises wavelengths within the first absorption wavelength band and within the second absorption wavelength band; capturing, in a side scattering orientation, light scattered by the spray within the sheet plane and determining a first intensity corresponding to an intensity of the captured light within the first emission wavelength band and a second intensity corresponding to an intensity of the captured light within the second emission wavelength band; and determining a characteristic of the spray based on the first intensity and the second intensity.
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
96.
SHROUD OF GAS TURBINE ENGINE AND METHOD OF MANUFACTURING THEREOF
A shroud for a gas turbine engine, the shroud including: an arcuate base extending circumferentially about a central axis; a pair of opposing flanges extending from respective axial ends of the arcuate base; and a pair of tangs spaced apart from the arcuate base and extending towards each other. Moreover, the arcuate base, the pair of flanges, and the pair of tangs define a slot. Each tang defines a width extending from the respective flange and a thickness 10 orthogonal to the width. Each tang includes an undulating pattern extending along a whole of the thickness and at least a portion of the width. The undulating pattern of each tang is configured to at least partially engage with the complementary mating features of the plurality of stator vanes. Further, a method of manufacturing the shroud by using a metal sheet is disclosed.
A gas turbine engine includes: first spool including first compressor and first turbine drivingly coupled by first main shaft; second spool including a second compressor and second turbine drivingly coupled by second main shaft; combustion equipment; electrical machine arrangement including: first rotor drivingly coupled to first main shaft and carrying array of rotor elements; second rotor drivingly coupled to second main shaft and carrying array of rotor elements; and stator arrangement including array of stator coils arranged around an axis, axis being coaxial with an axis of rotation of first rotor and an axis of rotation of second rotor. Stator arrangement is axially translatable relative to first and second rotors, between: a first axial position in which array of stator coils interacts with array of rotor elements of first rotor; and a second axial position in which array of stator coils interacts with array of rotor elements of second rotor.
A power electronics converter including: an input terminal; first and second DC output terminals; a branch including first and second semiconductor switches connected in series between the first and second DC output terminals, the input terminal connected to a node between the first and second semiconductor switches; a DC link capacitor connected between the first and second DC output terminals; and a resistive damping element connected in series with the DC link capacitor, wherein a damping factor of a circuit including the DC link capacitor, the resistive damping element and an inductance of the circuit with a short between the output terminals is at least 1.
H02M 1/14 - Dispositions de réduction des ondulations d'une entrée ou d'une sortie en courant continu
H02M 7/04 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continu sans possibilité de réversibilité par convertisseurs statiques
H02M 7/162 - 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 thyratron ou thyristor exigeant des moyens d'extinction utilisant uniquement des dispositifs à semi-conducteurs dans une configuration en pont
99.
HEATING DEVICE, HEATING APPARATUS AND APPARATUS FOR PROVIDING A FLOW OF GASEOUS HYDROGEN
A heating device comprises a device input and a device output, a conduit for conducting a flow of fluid from the device input to the device output, a hydrogen-burning torch arranged to heat the flow of fluid between the device input and the device output by flame-heating, and a spark ignitor arranged to light the torch. The heating device may be used to heat a flow of supercritical hydrogen derived from a store of liquid hydrogen to provide a flow of gaseous hydrogen. The device may be used in a fuel delivery system for a hydrogen-burning gas turbine engine, providing a weight saving compared to a hydrogen fuel delivery system having a fuel pre-heater which includes a heat-exchanger.
F24H 1/10 - Chauffe-eau instantanés, c.-à-d. dans lesquels il n'y a production de chaleur que lorsque l'eau s'écoule, p. ex. avec contact direct de l'eau avec l'agent chauffant
100.
Gas turbine engine with relative clocking of bifurcations
A gas turbine engine for an aircraft comprises: an air intake comprising a lip, a most upstream portion of which defining a highlight plane; an engine core comprising a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a fan located upstream of the engine core and adapted to rotate about an engine main axis, the fan comprising a plurality of fan blades having a respective leading edge, trailing edge, and tip, a forward-most portion of the tip leading edge of each fan blade defining a fan inlet plane; an air intake arranged upstream of, and configured to direct air to, the fan; a plurality of fan outlet guide vanes (FOGVs) arranged downstream of the fan in a bypass duct of the gas turbine engine; and upper and lower bifurcations arranged in the bypass duct and extending along respective radial directions.
