A mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gas turbine engine also includes a heat exchanger including an inner peripheral wall and an outer peripheral wall extending between an inlet and an outlet. The inner peripheral wall and the outer peripheral wall define a flow channel therebetween. The heat exchanger includes a plurality of fins disposed in the flow channel and dividing the flow channel into a plurality of flow passages.
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 is provided. The gas turbine engine includes: a thermal management system having a thermal fluid member having a flow of thermal fluid therethrough during operation of the gas turbine engine and a heat exchanger assembly, the heat exchanger assembly including: a core section comprising a plurality of heat exchange members; and a heat exchange manifold including a first direction pressure vessel in fluid communication with the thermal fluid member and a second direction pressure vessel extending from the first direction pressure vessel, the first and second direction pressure vessels each extending in a reference plane, the second direction pressure vessel in fluid communication with the first direction pressure vessel and with at least one of the plurality of heat exchange members.
F28D 1/03 - Appareils échangeurs de chaleur comportant des ensembles de canalisations fixes pour une seule des sources de potentiel calorifique, les deux sources étant en contact chacune avec un côté de la paroi de la canalisation, dans lesquels l'autre source de potentiel calorifique est une grande masse de fluide, p. ex. radiateurs domestiques ou de moteur de voiture avec des canalisations d'échange de chaleur immergées dans la masse du fluide avec des canalisations en forme de plaques ou de laminés
F28F 9/02 - Boîtes de distributionPlaques d'extrémité
3.
METHOD OF OPERATING A LUBRICATION SYSTEM FOR A TURBINE ENGINE
A method of operating a lubrication system for a turbine engine having a fan and one or more rotating components. The lubrication system includes a primary lubrication system and an auxiliary lubrication system. The method includes supplying lubricant to the one or more rotating components from the primary lubrication system during a normal operation of the turbine engine, determining whether at least one of a fuel pressure is less than a fuel pressure threshold or a hydraulic pressure is less than a hydraulic pressure threshold, and activating the auxiliary lubrication system to supply the lubricant to the one or more rotating components from the auxiliary lubrication system when the fuel pressure is less than the fuel pressure threshold or the hydraulic pressure is less than the hydraulic pressure threshold.
A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.
F02C 7/12 - Refroidissement des ensembles fonctionnels
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system is provided that adjusts parameters of the fuel based on data received from the sensors.
F02C 7/22 - Systèmes d'alimentation en combustible
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
A system and method for using a fuel with an engine, an airframe having an aircraft heat load, a fuel tank, and a fuel oxygen reduction unit are provided. The method includes receiving an inlet fuel flow in the fuel oxygen reduction unit for reducing an amount of oxygen in the inlet fuel flow; separating a fuel/gas mixture within the fuel oxygen reduction unit into an outlet gas flow and an outlet fuel flow exiting the fuel oxygen reduction unit; controlling a first portion of the outlet fuel flow to the engine; and controlling a second portion of the outlet fuel flow to the airframe to transfer heat between the second portion of the outlet fuel flow and the aircraft heat load.
A method of forming an acoustic liner that includes bonding a first side of a facesheet to a first adhesive side of a first solid adhesive film. A second solid adhesive film is bonded to a second side of the facesheet at a first adhesive side of the second solid adhesive film. An acoustic screen is bonded to a second adhesive side of the first solid adhesive film and an acoustic core is bonded to a second adhesive side of the second solid adhesive film.
G10K 11/168 - Sélection de matériaux de plusieurs couches de matériaux différents, p. ex. sandwiches
B32B 3/12 - Produits stratifiés comprenant une couche ayant des discontinuités ou des rugosités externes ou internes, ou une couche de forme non planeProduits stratifiés comprenant une couche ayant des particularités au niveau de sa forme caractérisés par une couche discontinue, c.-à-d. soit continue et percée de trous, soit réellement constituée d'éléments individuels caractérisés par une couche d'alvéoles disposées régulièrement, soit formant corps unique dans un tout, soit structurées individuellement ou par assemblage de bandes indépendantes, p. ex. structures en nids d'abeilles
B32B 5/02 - Produits stratifiés caractérisés par l'hétérogénéité ou la structure physique d'une des couches caractérisés par les caractéristiques de structure d'une couche comprenant des fibres ou des filaments
B32B 7/12 - Liaison entre couches utilisant des adhésifs interposés ou des matériaux interposés ayant des propriétés adhésives
B32B 37/12 - Procédés ou dispositifs pour la stratification, p. ex. par polymérisation ou par liaison à l'aide d'ultrasons caractérisés par l'usage d'adhésifs
An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. The fuel nozzle assembly can be configured to increase lateral provision of fuels to reduce flame scrubbing on combustor liners for the combustor.
General Electric Deutschland Holding GmbH (Allemagne)
Inventeur(s)
Osama, Mohamed
Yi, Xuan
Yagielski, John Russell
Huang, Shenyan
Abrégé
A rotor of an electric machine is provided. The rotor includes a plurality of laminates arranged along an axial direction, the plurality of laminates including a first laminate having: a body formed of a first material, the first material being a ferromagnetic material; and a structural element formed integrally with the body of a second material, the second material being a non-ferromagnetic material.
H02K 1/276 - Aimants encastrés dans le noyau magnétique, p. ex. aimants permanents internes [IPM]
H02K 1/02 - Détails du circuit magnétique caractérisés par le matériau magnétique
H02K 15/03 - Procédés ou appareils spécialement adaptés à la fabrication, l'assemblage, l'entretien ou la réparation des machines dynamo-électriques des corps statoriques ou rotoriques comportant des aimants permanents
H02K 21/14 - Moteurs synchrones à aimants permanentsGénératrices synchrones à aimants permanents avec des induits fixes et des aimants tournants avec des aimants tournant à l'intérieur des induits
A seal assembly for a turbomachine includes a rotating component, a non-rotating component having a ring carrier, a face seal ring, and a joint disposed between the ring carrier and the non-rotating component. The non-rotating component is arranged with the rotating component at a sealing interface. The joint is capable of linear and tilt displacement to allow the non-rotating component to displace with respect to the rotating component. The sealing interface defines a gap between the face seal ring and the rotating component. The seal assembly also includes at least one damping element arranged with the ring carrier for maintaining the seal assembly centrally in a radial direction in the turbomachine. Further, the damping element(s) includes at least one mass element.
An altitude insensitive composite coating includes a polymer resin and a plurality of ceramic particles in an electrophoretically formed layer. A fraction of the plurality of ceramic particles in the composite coating is at least 40% by volume. A dielectric strength of the composite coating is at least 30 kilovolts per millimeter (kV/mm) at an air pressure corresponding to high altitude.
C09D 5/00 - Compositions de revêtement, p. ex. peintures, vernis ou vernis-laques, caractérisées par leur nature physique ou par les effets produitsApprêts en pâte
C09D 7/61 - Adjuvants non macromoléculaires inorganiques
A swirler-ferrule assembly includes a radial swirler, a ferrule, a fuel nozzle, and a surface feature. The radial swirler includes a primary swirler vane having a primary air passage and a secondary swirler vane having a secondary air passage. The ferrule may be connected to the radial swirler. The surface feature may be located on the primary swirler vane and/or the ferrule. The surface feature may be configured to direct an air flow through the primary air passage away from a recirculation zone located upstream of the primary swirler vane. The surface feature has a trailing end and a distal end, and the fuel nozzle is axially aligned with the trailing end of the surface feature or is located axially downstream of the trailing end of the surface feature. The surface feature may have a plurality of grooves.
A mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gear includes a first gear that is a split sun gear including a forward sun gear and an aft sun gear separate from the forward sun gear. The forward sun gear and the aft sun gear are each rotationally coupled to a rotating shaft of the gas turbine engine.
Devices, systems, methods, and kits of parts for monitoring operation of an electron beam additive manufacturing systems are disclosed. A monitoring system includes one or more measuring devices positioned on the at least one wall in the interior of a build chamber of the additive manufacturing system. Each one of the one or more measuring devices includes a piezoelectric crystal. The monitoring system further includes an analysis component communicatively coupled to the one or more measuring devices. The analysis component is programmed to receive information pertaining to a frequency of oscillation of the piezoelectric crystal. A collection of material on the one or more measuring devices during formation of an article within the build chamber causes a change to the frequency of oscillation of the piezoelectric crystal that is detectable by the analysis component and usable to determine a potential build anomaly of the article.
A system for cooling oil in an aircraft turbine engine includes an intermediate support casing configured to be located between a low-pressure compressor and a high-pressure compressor of the aircraft turbine engine. The system further includes a heat exchanger for cooling the oil by heat exchange with air. The heat exchanger is at least partially integrated into the intermediate support casing.
A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan; a turbomachine drivingly coupled to the fan, the turbomachine comprising an outer casing; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip; and an accessory gearbox positioned at least partially inward of the outer casing of the turbomachine.
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
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux commande du rapport des débits des différents flux
Composite components and methods for adding a composite material to a composite component are provided. For example, a method comprises positioning a composite material segment against the composite component to form a component layup; applying an insulating material around at least a portion of the component layup to form an insulated layup; and densifying the insulated layup, where the composite component was previously densified before positioning the composite material segment against the composite component. In some embodiments, the composite material is ceramic matrix composite (CMC) and the composite material segment is a plurality of CMC plies. The composite component may be a CMC gas turbine engine component that comprises an original CMC component and a new CMC material segment joined to the original CMC component through the transfer of silicon between the original CMC component and the new CMC material segment during melt infiltration.
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
C04B 37/00 - Liaison des articles céramiques cuits avec d'autres articles céramiques cuits ou d'autres articles, par chauffage
Methods are provided for forming a coating on a surface of a substrate. The method may include: applying a negative charge to the surface of the substrate; electrophoretically depositing a slurry layer onto the surface of the substrate; and densifying the slurry layer on the surface of the substrate at a sintering temperature to form a sintered layer of the coating. The slurry layer may include a plurality of EBC material particles, a cationic polyelectrolyte, a plurality of polymeric binder particles, and a solvent. The plurality of EBC material particles may comprise barium strontium aluminosilicate (BSAS), mullite, silicon, rare earth compounds, or combinations thereof.
C09D 5/44 - Compositions de revêtement, p. ex. peintures, vernis ou vernis-laques, caractérisées par leur nature physique ou par les effets produitsApprêts en pâte pour des applications électrophorétiques
C25D 15/00 - Production électrolytique ou électrophorétique de revêtements contenant des matériaux incorporés, p. ex. particules, "whiskers", fils
A gas turbine engine is provided. The gas turbine engine includes: a thermal management system having a thermal fluid member having a flow of thermal fluid therethrough during operation of the gas turbine engine and a heat exchanger assembly, the heat exchanger assembly including: a core section comprising a plurality of heat exchange members; and a heat exchange manifold including a first direction pressure vessel in fluid communication with the thermal fluid member and a second direction pressure vessel extending from the first direction pressure vessel, the first and second direction pressure vessels each extending in a reference plane, the second direction pressure vessel in fluid communication with the first direction pressure vessel and with at least one of the plurality of heat exchange members.
F28D 7/16 - Appareils échangeurs de chaleur comportant des ensembles de canalisations tubulaires fixes pour les deux sources de potentiel calorifique, ces sources étant en contact chacune avec un côté de la paroi d'une canalisation les canalisations étant espacées parallèlement
A CVFDR system of an aircraft includes a cockpit voice and flight data recorder (CVFDR) communicatively coupled, via a data communication network, to a set of flight recorder modules. The CVFDR receives a first voltage from a remote first power source. In the event of an interruption of the first voltage, the CVFDR receives a second voltage from a local second power source for a predetermined period.
A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A temperature sensor is provided within the engine and configured to detect a gas temperature within the engine core. A set of blades is circumferentially arranged in the turbine section. A blade in the set of blades includes an outer wall bounding an interior, a cooling conduit within the interior, and a plurality of film holes fluidly coupled to the cooling conduit.
A fuel cell assembly is provided, including a fuel cell stack having a fuel cell, the fuel cell having a cathode and an anode; and a multi-gas sensor configured to sense gas composition data of a flow of output products from the cathode, gas composition data of a flow of output products from the anode, gas composition data of a fluid surrounding the fuel cell, or a combination thereof to determine fuel cell leakage diagnostic information.
B64D 27/355 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles à combustible
F02C 7/22 - Systèmes d'alimentation en combustible
G01M 3/26 - 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
A computer implemented method for servicing an engine including receiving information including an initial condition profile, CP1, of the engine; forming a workscope associated with a servicing operation of the engine in view of the initial condition profile, CP1; servicing the engine in view of the workscope; determining at least in part an updated condition profile, CP2, of the engine in view of information acquired during the service; and storing the updated condition profile, CP2, for use in a subsequent service operation.
B60S 5/00 - Maintenance, entretien, réparation ou révision des véhicules
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
G07C 5/08 - Enregistrement ou indication de données de marche autres que le temps de circulation, de fonctionnement, d'arrêt ou d'attente, avec ou sans enregistrement des temps de circulation, de fonctionnement, d'arrêt ou d'attente
A mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gas turbine engine includes an electric power system including at least one electric machine. The electric power system includes a plurality of power converters and a plurality of power distribution management units. At least two of the plurality of power converters or the plurality of power distribution management units are integrated together in a single housing.
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 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F16H 57/02 - Boîtes de vitessesMontage de la transmission à l'intérieur
F16H 57/025 - Support des boîtes de vitesses, p. ex. bras de couple, ou attachement à d'autres dispositifs
H02M 1/00 - Détails d'appareils pour transformation
H02M 1/32 - Moyens pour protéger les convertisseurs autrement que par mise hors circuit automatique
H02M 7/00 - Transformation d'une puissance d'entrée en courant alternatif en une puissance de sortie en courant continuTransformation d'une puissance d'entrée en courant continu en une puissance de sortie en courant alternatif
25.
SELF-HEALING ADAPTIVE CONTROLLER FOR INDUSTRIAL ASSET ABNORMALITY ACCOMMODATION
A training data store may contain training data associated with monitoring node values during normal operation of an industrial asset and simulated abnormal data. An offline model tuning platform accesses the training data from normal operation of the industrial asset and the simulated abnormal data in the training data store. Based on the training data from normal operation of the industrial asset, the simulated abnormal data, an abnormal operating condition, and a constrained optimization solution, controller tuning parameters are created for at least one tuned data-driven adaptive controller such that an operating condition of the industrial asset will move from the abnormal operating condition to a normal operation condition through a stable trajectory. An online monitoring platform receives a stream of current monitoring node values and, when the abnormal operating condition is detected, utilizes the controller tuning parameters to implement the at least one tuned data-driven adaptive controller.
G05B 13/02 - Systèmes de commande adaptatifs, c.-à-d. systèmes se réglant eux-mêmes automatiquement pour obtenir un rendement optimal suivant un critère prédéterminé électriques
A turbine engine having a longitudinal centerline axis. The turbine engine including a fan comprising a plurality of fan blades that rotate about the longitudinal centerline axis and a rotational component coupled to the fan. The turbine engine including a fluid circuit for supplying fuel or lubricant to the turbine engine and a hydraulic fan brake coupled to the fluid circuit to prevent rotation of the rotational component, thus preventing rotation of the fan. The hydraulic fan brake including a hydraulic cylinder fluidly coupled to the fluid circuit and a valve coupled to the hydraulic cylinder and having a first valve position that disengages the hydraulic fan brake to allow rotation of the rotational component and a second valve position that engages the hydraulic fan brake to prevent rotation of the rotational component.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F02C 9/26 - Commande de l'alimentation en combustible
F04C 2/08 - Machines ou pompes à piston rotatif du type à engrènement extérieur, c.-à-d. avec un engagement des organes coopérants semblable à celui d'engrenages dentés
F15B 13/04 - Dispositifs de distribution ou d'alimentation du fluide caractérisés par leur adaptation à la commande de servomoteurs pour utilisation avec un servomoteur unique
27.
GAS TURBINE ENGINE HAVING A HEAT EXCHANGER LOCATED IN AN ANNULAR DUCT
A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.
A hybrid electric aircraft equipped with gyroscopic stabilization control is provided. In one aspect, a hybrid electric aircraft includes a turbo-generator having a gas turbine engine and an electric generator operatively coupled thereto for generating electrical power. The turbo-generator defines a rotation axis. The aircraft also includes one or more electrically-driven propulsors for producing thrust for the aircraft. In addition, the aircraft includes a pivot mount operatively coupled with the turbo-generator. To provide gyroscopic stabilization control of the aircraft, the pivot mount is controlled to adjust the rotation axis of the turbo-generator relative to a prime stability axis of the aircraft. Additionally or alternatively, a rotational speed of the turbo-generator can be changed to provide gyroscopic stabilization control of the aircraft.
B64C 17/06 - Stabilisation des aéronefs non prévue ailleurs par gravité ou par appareil actionné par inertie par appareil gyroscopique
B64C 29/00 - Aéronefs capables d'atterrir ou de décoller à la verticale, p. ex. aéronefs à décollage et atterrissage verticaux [ADAV, en anglais VTOL]
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 27/14 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz à l'intérieur des fuselages ou fixés à ceux-ci
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur ou l'énergie de ressorts
B64D 31/02 - Dispositifs amorçant la mise en œuvre
29.
BEARINGS WITH GROOVES AND METHODS OF PRODUCING THE SAME
Bearings with grooves and methods of producing the same are disclosed. Examples disclosed herein include a bearing having a bearing surface with a groove, the bearing positioned adjacent to a shaft, the groove facing the shaft, and a negative thermal expansion (NTE) material positioned in the groove, the NTE material at least partially filling the groove.
The present disclosure is directed to an additive manufacturing system configured to generate an electron beam directed toward a target to generate x-ray flux. The x-ray flux is directed toward the component through at least one plate with a pinhole. Interactions between the component and the x-ray flux generate x-ray radiation. The at least one detector is configured to detect the x-ray radiation through a pinhole. An analysis component is configured to generate an image comprising a three-dimensional component based on the x-ray radiation detected by the at least detector.
G01N 23/046 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p. ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et formant des images des matériaux en utilisant la tomographie, p. ex. la tomographie informatisée
B29C 64/393 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
31.
VIBRATION SENSOR AND METHODS FOR DEPOSITION MONITORING
Apparatuses and methods are provided herein that are useful to monitoring a coating deposition process. In some embodiments, a method of monitoring a coating process for a target object involves applying a coating to at least a portion of a beam during at least part of a coating deposition process to obtain a coated beam. The method also includes exciting the coated beam such that the coated beam vibrates. The method also includes monitoring the vibration response of the coated beam and determining at least one of a deposition rate or a drying rate for the target object based on the change in the frequency response of the coated beam.
A system comprises an energy source and detector. The energy source is configured to emit energy from a focal spot. The detector comprises an array of pixels. Each of the pixels of the array of pixels has a centerline extending longitudinally through the pixel. The centerline of each pixel is focally aligned with the focal spot of the energy source. Each of the pixels receives selected amounts of the energy from the energy source that have passed through an object.
G01N 23/04 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p. ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et formant des images des matériaux
The present disclosure is generally related to a fault isolation sensor system for use with determining whether a potential fault is more likely to be on a rotor side or a stator side. The measurement signals are transmitted from the rotor antenna to the stator antenna, and then from the stator antenna to a controller. The controller is configured to monitor the measurement signals. If the measurement signal is outside of a predetermined range or past a predetermined threshold, then the stator antenna can be interrogated with an interrogation signal with a reflected signal being compared with the interrogation signal and a ratio thereof being used to identify the potential side of the fault.
G01R 31/11 - Localisation de défauts dans les câbles, les lignes de transmission ou les réseaux en utilisant des méthodes de réflexion d'impulsion
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
H02K 11/20 - Association structurelle de machines dynamo-électriques à des organes électriques ou à des dispositifs de blindage, de surveillance ou de protection pour la mesure, la surveillance, les tests, la protection ou la coupure
34.
METHODS AND APPARATUS TO IMPROVE FAN OPERABILITY CONTROL USING SMART MATERIALS
Systems, apparatus, articles of manufacture, and methods are disclosed to improve fan operability control using smart materials. An engine comprising an engine surface in an airflow path, a sensor positioned on the engine surface, and a smart-material-based feature positioned on the engine surface, the smart-material-based feature triggered to modify the airflow path when the sensor outputs an indication of a detected deviation from a reference value of an operating parameter of the engine.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
A combustor liner for a combustor of a gas turbine includes an outer liner having a plurality of outer liner segments, and an inner liner having a plurality of inner liner segments. Each segment of the inner liner and the outer liner includes at least one slotted dilution opening therethrough extending in a circumferential direction, and each slotted dilution opening includes a deflector wall extending radially from the respective liner into a dilution zone of a combustion chamber between the outer liner and the inner liner. The at least one slotted dilution opening may be a curved slot (either concave or convex) dilution opening, and the curved slot dilution openings for each segment of the outer liner and the inner liner may be connected so as to provide a wavy slotted dilution opening extending annularly through the outer liner and the inner liner.
A turbine engine and method for controlling nitrogen oxides present within a combustor of the turbine engine. The turbine engine having a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. The combustion section having a combustor liner having a first end, a second end, opposing the first end, and at least partially defining a combustion chamber extending between the first and second ends. A dome assembly is mounted to the combustor liner at the first end and defines a dome inlet of the combustion chamber. There are multiple sets of dilution holes including a first set of dilution holes provided in the combustor liner downstream from the dome inlet and a second set of dilution holes provided in the combustor liner between the first set of dilution holes and the dome inlet.
An aircraft defining a vertical direction, a longitudinal direction, and a transverse direction is provided. The aircraft includes: a fuselage; a propulsion system comprising a power source and a plurality of vertical thrust electric fans driven by the power source; and a wing extending from the fuselage in the transverse direction. The wing includes a support structure that comprises a plurality of first support members and a plurality of second support members. The plurality of first support members extending at least partially between the plurality of second support members. The plurality of vertical thrust electric fans arranged between the plurality of first support members and the plurality of second support members.
A fuel nozzle valve includes a fuel nozzle valve liner having a channel with an opening for allowing fuel to flow therethrough and a seat. A plunger has a stud and a base substantially perpendicular to the stud, the plunger being configured to move relative to the fuel nozzle valve liner to seal or to open the opening of the fuel nozzle valve. The fuel nozzle valve further includes a metal resilient member configured to contact the base of the plunger and the seat of the fuel nozzle valve liner to seal the opening of the fuel nozzle valve when the plunger is moved to seal the fuel nozzle valve.
A fuel oxygen reduction unit is provided for reducing an oxygen content of a flow of liquid fuel to an engine. The fuel oxygen reduction unit includes: a stripping gas supply line for providing a flow of stripping gas; a contactor defining a liquid fuel inlet, a stripping gas inlet and a fuel/gas mixture outlet, the stripping gas supply line in airflow communication with the stripping gas inlet; a means for modulating the flow of stripping gas through the stripping gas supply line; and a controller operable with the means for modulating the flow of stripping gas through the stripping gas supply line to modulate the flow of stripping gas through the stripping gas supply line in response to an engine operability parameter.
A turbine engine that includes an engine core having at least a compressor section and a combustion section. The combustion section includes a combustor. The combustor section or combustor includes a fuel-air mixing assembly fluidly coupled to the compressor section. The fuel-air mixing assembly includes an outer wall, a center body at least partially circumscribed by the outer wall, and an annular flow passage between the outer wall and center body. At least one fuel orifice includes a fuel outlet fluidly coupled to the annular flow passage.
General Electric Company Polska sp. z o.o. (Pologne)
Inventeur(s)
Jasiczek, Michal
Sinha, Shatil
Keshavan, Hrishikesh
Gemeinhardt, Gregory Carl
Graves, John Harvey
Lin, Wendy Wenling
Abrégé
A geopolymer composite material includes a geopolymer resin, a fibrous reinforcement material embedded in the geopolymer resin for reinforcing the geopolymer resin, and a protective coating material covering the fibrous reinforcement material embedded in the geopolymer resin to provide a coated fibrous reinforcement material. Further, the geopolymer resin and the coated fibrous reinforcement material are combined together to form a prepreg containing the geopolymer resin being pre-impregnated into the coated fibrous reinforcement material and being curable into a solid matrix.
C04B 20/00 - Emploi de matières comme charges pour mortiers, béton ou pierre artificielle prévu dans plus d'un groupe et caractérisées par la forme ou la répartition des grainsTraitement de matières spécialement adapté pour renforcer leur propriétés de charge dans les mortiers, béton ou pierre artificielle prévu dans plus d'un groupe de Matières expansées ou défibrillées
C04B 28/00 - Compositions pour mortiers, béton ou pierre artificielle, contenant des liants inorganiques ou contenant le produit de réaction d'un liant inorganique et d'un liant organique, p. ex. contenant des ciments de polycarboxylates
C04B 40/00 - Procédés, en général, pour influencer ou modifier les propriétés des compositions pour mortiers, béton ou pierre artificielle, p. ex. leur aptitude à prendre ou à durcir
A lubrication system for a turbine engine. The turbine engine includes a fan having a fan shaft and one or more rotating components. The lubrication system includes one or more tanks that stores lubricant therein, a primary lubrication system, and an auxiliary lubrication system. The primary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components during normal operation of the turbine engine. The auxiliary lubrication system includes an auxiliary pump that is coupled to the fan shaft. The auxiliary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components based on a pressure of the lubricant in the primary lubrication system. Rotation of the fan shaft causes the auxiliary pump to pump the lubricant to the one or more rotating components.
An apparatus for servicing internal components of an aircraft engine includes a flexible hollow tube; a latching mechanism connected to or incorporated with the flexible hollow tube; and a servicing device to be inserted through the flexible hollow tube. The servicing device is freely moveable through the flexible hollow tube and decoupled from the latching mechanism. The flexible hollow tube is shaped and configured so as to enable proximate positioning of the flexible hollow tube with respect to a rotatable component of an aircraft engine allowing attachment of the flexible hollow tube to the rotatable component via the latching mechanism after the flexible hollow tube is inserted through an entry port of the aircraft engine.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
G01N 21/95 - Recherche de la présence de criques, de défauts ou de souillures caractérisée par le matériau ou la forme de l'objet à analyser
44.
Engine component assembly with ceramic matrix composite component and connection pin
A turbine engine component assembly includes a first part comprising a ceramic matrix composite material and having a first flange defining a first borehole, a second part defining a second borehole, and a pin inserted, in an axial direction, through the first borehole and the second borehole to connect the first part to the second part. The pin includes a first contact portion positioned at a first axial location of the pin corresponding to the first borehole and an elongated portion extending axially from the first contact portion. The first contact portion includes a first chamfered section, a second chamfered section, and a first contact surface between the first chamfered section and the second chamfered section. The first chamfered section and the second chamfered section slope away from the first contact surface with decreasing diameters.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F01D 25/00 - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
An inlet duct for a nacelle of a ducted fan engine includes an inlet portion having an inlet lip and a hardwall portion, a highlight plane defined at an upstream end of the inlet portion, and an acoustic liner downstream of the inlet portion, a hardwall-acoustic liner interface defined at an interface of the hardwall portion and the acoustic liner. The inlet portion and the acoustic liner are coupled at an inlet-acoustic liner interface extending circumferentially about an inner circumferential surface of the inlet duct. The inlet lip varies circumferentially and axially with respect to the highlight plane and the inlet centerline axis, and includes a plurality of inlet lip crests arranged along the highlight plane, and a plurality of inlet lip troughs arranged downstream of the highlight plane. The hardwall-acoustic liner interface extends circumferentially about the inner circumferential surface and is arranged axially parallel to the highlight plane.
F02C 7/045 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs destinés à supprimer le bruit
Hybrid electric propulsion systems includes a combustion engine and an electric motor. The hybrid electric propulsion systems may include or utilize a non-transitory computer-readable medium comprising computer-executable instructions, which when executed by a processor associated with the hybrid electric propulsion system, cause the processor to perform a method that includes determining an occurrence of a thrust asymmetry in the hybrid electric propulsion system, and controlling the electric motor to decrease an efficiency of the electric motor for a transient time period sufficient to reduce a torque output of the combustion engine to match an electrical load on the combustion engine.
B64D 31/06 - Dispositifs amorçant la mise en œuvre actionnés automatiquement
B60W 10/06 - Commande conjuguée de sous-ensembles de véhicule, de fonction ou de type différents comprenant la commande des ensembles de propulsion comprenant la commande des moteurs à combustion
B60W 10/08 - Commande conjuguée de sous-ensembles de véhicule, de fonction ou de type différents comprenant la commande des ensembles de propulsion comprenant la commande des unités de traction électrique, p. ex. des moteurs ou des générateurs
B60W 20/10 - Commande de l'apport de puissance de chacun des moteurs primaires pour répondre à la demande de puissance requise
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur ou l'énergie de ressorts
B64D 37/00 - Aménagements relatifs à l'alimentation des groupes moteurs en carburant
B64D 45/00 - Indicateurs ou dispositifs de protection d'aéronefs, non prévus ailleurs
A turbine engine having a compressor section, a combustor section, a turbine section, and a rotatable drive shaft. A bypass conduit couples the compressor section to the turbine section. At least one centrifugal separator is fluidly coupled to the bypass stream, where the at least one centrifugal separator includes a body, a center body, a separator inlet, and a separator outlet fluidly coupled with the turbine section to output a reduced-particle stream that is provided to the turbine section for cooling. The centrifugal separator includes an angular velocity increaser, a flow splitter, a first outlet passage defined by an inner annular wall that receives the reduced-particle stream, and an angular velocity decreaser located downstream of the flow splitter. A second outlet passage receives the concentrated-particle stream.
B01D 45/16 - Séparation de particules dispersées dans des gaz ou des vapeurs par gravité, inertie ou force centrifuge en utilisant la force centrifuge produite par le mouvement hélicoïdal du courant gazeux
B04C 3/00 - Appareils dans lesquels la direction axiale du tourbillon ne change pas
B04C 3/06 - Structures des entrées ou sorties de la chambre où se produit le tourbillon
F01D 5/18 - Aubes creusesDispositifs de chauffage, de protection contre l'échauffement ou de refroidissement des aubes
F01D 9/06 - Conduits d'admission du fluide à l'injecteur ou à l'organe analogue
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 7/052 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes comportant des dispositifs séparateurs de poussière
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
F02K 3/04 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux
A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
F02K 3/065 - 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 des soufflantes avant et arrière
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
A system (e.g., an ultrasound imaging system) is provided. The system includes an ultrasound probe having a cable, and an ultrasound probe holder configured to receive the ultrasound probe. The system further includes a housing supported by a base. The housing includes a connector port and a cable manage passage. The cable manage passage positioned at an upper end of the housing distal to the base. The cable extending through the cable manage passage, and is attached to the connector port. The ultrasound probe holder is coupled to a front side of the housing.
A composite airfoil assembly for a turbine engine includes an airfoil defining an airfoil interior, and an inner support structure at least partially located within the airfoil interior. The inner support structure can include intertwined fibers defining a three-dimensional structure. A laminate overlay surrounds at least a portion of the inner support structure.
A composite airfoil assembly for a turbine engine includes an airfoil defining an airfoil interior, and an inner support structure at least partially located within the airfoil interior. The inner support structure can include a first core, a second core, and at least one pin. A laminate overlay surrounds at least a portion of the inner support structure.
A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (Li), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.
F02K 3/115 - Chauffage du flux dérivé à l'aide d'un échange indirect de chaleur
F02K 3/04 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux
A snake-arm robot and a servicing device are mechanically coupled. The mechanical coupling is accomplished by a longitudinal insertion of the snake-arm robot into the servicing device or the servicing device into the snake-arm robot. An actuator moves the snake-arm robot through a passage within an engine until the snake-arm robot reaches a desired location. The movement of the snake-arm robot concurrently moves the servicing device through the passage. Subsequently, the snake-arm robot is de-coupled from the servicing device and the snake-arm robot is removed from the engine while leaving the servicing device in place within the engine.
B25J 13/08 - Commandes pour manipulateurs au moyens de dispositifs sensibles, p. ex. à la vue ou au toucher
B23P 6/00 - Remise en état ou réparation des objets
B25J 9/06 - Manipulateurs à commande programmée caractérisés par des bras à articulations multiples
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
G01M 15/02 - Détails ou accessoires pour appareils de test
A turbine engine includes a rotor, a stator having an aft wall, and a seal assembly having a plurality of seal segments disposed between the rotor and the stator. The rotor, the stator, and the seal assembly are arranged together to define a high pressure region and a low pressure region. The turbine engine also includes at least one biasing member engaged with one or more of the plurality of seal segments. The plurality of seal segments include a primary seal segment and a secondary seal segment connected together via a flexible joint. As such, the flexible joint allows for angular misalignment between the primary seal segment and the secondary seal segment, thereby allowing the primary seal segment to move with the rotor while the secondary seal segment maintains contact with the aft wall of the stator.
This disclosure is directed to seal assemblies for a turbomachine. The seal assemblies include one or more paired rotors and stators and at least one interface between the rotors and the stators. The components of the stator may be axially and radially movable by vibrations and other mechanical interference. The stators comprise a sealing element, a seal housing, and a stator interface connected to the engine housing. In some examples, seal assembly includes a damping element to isolate one or more of the rotating components from vibrations mechanical interference that might misalign the rotating components from the stationary components while the turbomachine is operational. In some examples, the damping element is positioned between the seal housing and the stator interface. In other examples, the damping element is positioned between the stator interface and the engine housing.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
A rotary component for a gas turbine engine includes a plurality of rotor blades operably coupled to a rotating shaft extending along the central axis and an outer casing arranged exterior to the plurality of rotor blades in a radial direction of the gas turbine engine. The outer casing defines a gap between a blade tip of each of the plurality of rotor blades and the outer casing. The outer casing includes a plurality of features formed into an interior surface thereof. Each of the plurality of features includes one or more design parameters that are perturbed about a mean design parameter for stall performance so as to provide a circumferential variation in wake strengths associated with the plurality of rotor blades, thereby reducing operational noise of the gas turbine engine.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
F02C 7/045 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs destinés à supprimer le bruit
A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.
A composition of matter is generally provided, in one embodiment, a titanium alloy comprising 5 wt % to 8 wt % aluminum; 2.5 wt % to 5.5 wt % vanadium; 0.1 wt % to 2 wt % of one or more elements selected from the group consisting of iron and molybdenum; 0.01 wt % to 0.2 wt % carbon; up to 0.3 wt % oxygen; silicon and copper; and titanium. A turbine component is also generally provided, in one embodiment, that comprises an article made from a titanium alloy. Additionally, methods are also generally provided for making an alloy component having a beta transus temperature and a titanium silicide solvus temperature.
C22F 1/18 - Métaux réfractaires ou à point de fusion élevé ou leurs alliages
B21K 3/04 - Fabrication de pièces de moteurs ou de machines similaires, non couverte par Fabrication d'hélices ou d'organes similaires d'aubes, p. ex. de turbinesRefoulement des pieds d'aubes
Example variable bleed valve assemblies for a gas turbine engine are disclosed herein, including a port extending radially outward from a compressor section of the gas turbine engine, the port defining a variable bleed valve cavity, the port to resonate at a resonant frequency based on an operating condition of the gas turbine engine, and an acoustic suppressor positioned on a wall of the port, the acoustic suppressor extending circumferentially along the wall by a length greater than a cross-sectional width of the acoustic suppressor, the acoustic suppressor defining a resonant cavity based on the length and the cross-sectional width, the acoustic suppressor including a perforated portion, the acoustic suppressor tuned to resonate at the resonant frequency based on the length and the perforated portion.
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
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
60.
GEARBOX ASSEMBLY LUBRICATION SYSTEM FOR A TURBINE ENGINE
A lubrication system for a gearbox assembly for a turbine engine. The gearbox assembly includes a gear assembly including one or more gears and one or more bearings. The lubrication system includes a sump. The sump is a primary reservoir that has a first lubricant level. The lubrication system also includes a secondary reservoir in the gearbox assembly. The secondary reservoir has a second lubricant level that is greater than the first lubricant level. A plurality of drain ports includes a first drain port and a second drain port. The lubrication system fills the secondary reservoir with lubricant between the first lubricant level and the second lubricant level and a portion of the lubricant drains though the second drain port. The one or more gears collects the lubricant to supply the lubricant from the secondary reservoir to the one or more gears or to the one or more bearings.
F16H 57/04 - Caractéristiques relatives à la lubrification ou au refroidissement
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 turbine engine has a longitudinal centerline axis. The turbine engine includes a fan, a turbo-engine, and a unidirectional brake. The fan includes a plurality of fan blades that rotate in a first direction about the longitudinal centerline axis. The turbo-engine includes a combustor that combusts compressed air and fuel to generate combustion gases and a low-pressure turbine including a low-pressure shaft. The low-pressure turbine receives the combustion gases to rotate the low-pressure turbine. The fan is coupled to the low-pressure shaft such that rotation of the low-pressure shaft causes the fan to rotate in the first direction. A unidirectional brake is coupled to the low-pressure shaft to prevent rotation of the low-pressure shaft and, thus, the fan in a second direction opposite the first direction.
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 defines an axial direction and a radial direction and comprises a turbomachine having an unducted primary fan, a core engine a combustor casing enclosing a combustor and defining an outer surface, a core cowl surrounding at least a portion of the core engine. The outer surface of the core cowl defines a peak cowl diameter (D) in the radial direction, and the outer surface of the combustor casing defines a maximum combustor casing diameter (d) along the radial direction. The core engine defines an overall core axial length (L) along the axial direction and an under-core cowl axial length (L1) along the axial direction. The gas turbine engine defines a core cowl diameter ratio (CDR) equal to the peak cowl diameter (D) divided by the maximum combustor casing diameter (d) and a core cowl length ratio (CLR) equal to the under-core cowl axial length (L1) divided by the overall core axial length (L). The CDR is between 2.7 and 3.5 and the CLR is between 0.25 and 0.50.
Methods and systems for calibrating a model are provided herein. In some embodiments, the methods include receiving, via a control circuit, test data for an operational parameter of a real-world system, such as an engine, from operational tests. The control circuit also receives model data for the operational parameter from simulations performed via a model of the engine. The control circuit then compresses the test data and the model data to generate compressed test data and compressed model data and fusing the compressed test data with the compressed model data to generate fused data. The control circuit performs parallel Bayesian inference simulations using the fused data to identify at least one value for a tuning parameter of the model. The control circuit may identify and select tuning parameters to match the model data with test data, the model data may be one or more model outputs (i.e., output parameters).
G06F 30/27 - Optimisation, vérification ou simulation de l’objet conçu utilisant l’apprentissage automatique, p. ex. l’intelligence artificielle, les réseaux neuronaux, les machines à support de vecteur [MSV] ou l’apprentissage d’un modèle
64.
METHOD OF GENERATING POWER WITH AN ELECTRIC MACHINE
General Electric Deutschland Holding GmbH (Allemagne)
Inventeur(s)
Zatorski, Darek Tomasz
Ostdiek, David Marion
Osama, Mohamed
Solomon, William Joseph
Abrégé
A method is provided of generating electric power with an electric machine. The method includes rotating a rotor of the electric machine relative to a stator of the electric machine with a shaft of a gas turbine engine during an operating condition of the gas turbine engine, the gas turbine engine being a three-stream gas turbine engine defining an axial direction, the three-stream gas turbine engine comprising: the shaft, a primary fan operatively coupled with the shaft, a mid-fan positioned downstream of the primary fan and operatively coupled with the shaft, a low pressure turbine operatively coupled with the shaft, wherein rotating the rotor of the electric machine relative to the stator of the electric machine comprises generating an electric machine power during the operating condition and generating a low pressure turbine power during the operating condition.
Example bearing lubrication system and methods of operating the same are disclosed herein. An example closed loop system to provide a lubricant to a fluid pump includes a lubrication flow network disposed within the fluid pump; a sensor fluidly coupled to the fluid pump to measure a condition of a fluid that is to enter the lubrication flow network; a first transport bus fluidly coupled to the lubrication flow network, the first transport bus to transport an inert gas; a controller to actuate a valve fluidly coupled to the first transport bus, the controller to transmit signals to the valve based on the condition of the fluid to cause the valve to open or close; and a separator fluidly coupled between an outlet of the fluid pump and the first transport bus, the separator to separate the fluid and the inert gas.
F16N 7/40 - Installations à huile ou autre lubrifiant non spécifié, à réservoir ou autre source portés par la machine ou l'organe machine à lubrifier avec pompe séparéeInstallations centralisées de lubrification à circuit fermé
F16N 29/00 - Dispositifs particuliers dans les installations ou les systèmes de lubrification indiquant ou détectant des conditions indésirablesUtilisation des dispositifs sensibles à ces conditions dans les installations ou les systèmes de lubrification
F16N 39/02 - Dispositions pour conditionner des lubrifiants dans les circuits de lubrification par refroidissement
66.
Combustion section with a primary combustor and a set of secondary combustors
A turbine engine with a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. A combustion section for the turbine engine, having a primary combustor liner including an inner liner and an outer liner annular about an engine centerline. A dome wall extending between the inner liner and the outer liner. A set of primary dome inlets located in the dome wall and circumferentially arranged about the engine centerline. A set of secondary combustors fluidly coupled to a primary combustion chamber, the set of secondary combustors including a first mini combustor and a second mini combustor.
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
F23R 3/42 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la disposition ou la forme des tubes à flamme ou des chambres de combustion
67.
Turbine engine with fan bypass water injection to augment thrust
A gas turbine has a core turbine engine, and a fan having a plurality of fan blades. A nacelle surrounds the fan and at least a portion of the core turbine engine. The nacelle defines an inlet arranged upstream of the fan, and a bypass flow passage downstream of the fan defined between the nacelle and the core turbine engine. The gas turbine also includes a bypass flow passage water injection system that includes (a) at least one water injection nozzle assembly arranged to inject water into at least one of the inlet of the nacelle, or into the bypass flow passage, and (b) a water injection supply system arranged to supply water from a storage tank to the at least one water injection nozzle assembly. Water is provided by the bypass flow passage water injection system during a high power operating states of the gas turbine to augment thrust.
F02K 3/02 - 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
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
A lock lug for inhibiting movement of a plurality of rotor blades includes a body and an engagement mechanism. The body is sized and configured to be received within the rim slot of the rotor disk and defines a dovetail receiving aperture. The engagement mechanism extends from the body and has a retracted configuration configured to allow entry and exit of a dovetail of at least one or more of the plurality of rotor blades into and out of the dovetail receiving aperture and an extended configuration configured to block the dovetail from entering the dovetail receiving aperture.
A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.
A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.
A gas turbine engine is provided. The gas turbine engine includes a turbomachine comprising a low speed spool; a rotor assembly coupled to the low speed spool; an electric machine mechanically coupled to the low speed spool at a connection point of the low speed spool; and a clutch positioned in the torque path of the low speed spool between the connection point and the rotor assembly
F01D 5/18 - Aubes creusesDispositifs de chauffage, de protection contre l'échauffement ou de refroidissement des aubes
F01D 5/30 - Fixation des aubes au rotorPieds de pales
F01D 7/00 - Rotors à aubes réglables en marcheLeur commande
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F01D 15/12 - Combinaisons avec des transmissions mécaniques
F01D 17/16 - Organes de commande terminaux disposés sur des parties du stator faisant varier l'aire effective de la section transversale des injecteurs ou tuyères de guidage en obturant les injecteurs
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
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
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
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
F02C 9/22 - Commande du débit du fluide de travail par étranglementCommande du débit du fluide de travail par réglage des aubes par réglage des aubes de turbine
F02K 1/66 - Inversion du flux de la soufflante en inversant les aubes du ventilateur
F02K 1/76 - Commande ou régulation des inverseurs de poussée
72.
POWDER RECLAMATION SYSTEM FOR MULTIPLE METAL POWDER PROCESSING DEVICES
A powder reclamation system is provided. The powder reclamation system includes a support structure; a filter housing movable relative to the support structure, the filter housing defining an inlet and an outlet; a raw reclaimed powder hopper in communication with the inlet of the filter housing; a first reclamation passageway in communication with the raw reclaimed powder hopper and configured to be in communication with a first metal powder processing device to recover a first unused portion of a first powder from the first metal powder processing device to the raw reclaimed powder hopper; and a second reclamation passageway in communication with the raw reclaimed powder hopper and configured to be in communication with a second metal powder processing device to recover a second unused portion of a second powder from the second metal powder processing device to the raw reclaimed powder hopper.
B07B 1/46 - Éléments de structure constitutifs des tamis en généralNettoyage ou chauffage des tamis
B07B 1/54 - Nettoyage par des dispositifs batteurs
B07B 9/00 - Combinaisons d'appareils à cribler ou tamiser ou à séparer des solides par utilisation de courants de gazDisposition générale des installations, p. ex. schéma opératoire
B22F 10/00 - Fabrication additive de pièces ou d’objets à partir de poudres métalliques
B22F 12/00 - Appareils ou dispositifs spécialement adaptés à la fabrication additiveMoyens auxiliaires pour la fabrication additiveCombinaisons d’appareils ou de dispositifs pour la fabrication additive avec d’autres appareils ou dispositifs de traitement ou de fabrication
General Electric Company Polska sp. z o.o. (Pologne)
Inventeur(s)
Owoeye, Eyitayo James
Ganiger, Ravindra Shankar
Pazinski, Adam Tomasz
Abrégé
A gas turbine engine is provided having a plurality of outlet guide vanes, each defining an internal thermal fluid passageway. The engine defining an Outlet Guide Vane Cooling Capacity greater than 0.01 and less than 13, wherein OGVCC equals:
A gas turbine engine is provided having a plurality of outlet guide vanes, each defining an internal thermal fluid passageway. The engine defining an Outlet Guide Vane Cooling Capacity greater than 0.01 and less than 13, wherein OGVCC equals:
[
HTSA
OGV
×
BPR
(
BPR
+
1
)
×
C
air
×
(
T
inlet
-
T
air
)
×
v
flight
×
D
fan
Fn
Total
×
v
tip
speed
×
Δ
H
]
1
/
3
,
and
wherein
HTSA
OGV
=
N
vane
×
D
fan
2
×
(
1
-
R
OGV
_
ratio
2
)
2
×
sin
(
180
/
N
vane
)
×
sin
θ
OGV
×
f
OGV
.
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
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
74.
SYSTEM AND METHOD FOR EFFICIENTLY DETERMINING A PHASE SHIFT IN A PROPULSION SYSTEM
A propulsion system includes at least two propulsors. The at least two propulsors each comprising a fan having a plurality of fan blades. A controller includes memory and one or more processors. The memory stores instructions that when executed by the one or more processors cause the system to perform the following: determine a pairwise phase difference between one propulsor of the at least two propulsors and another propulsor of the at least two propulsors; generate a reference phase angle; determine a target phase shift for each propulsor of the at least two propulsors; and adjust a speed of each propulsor of the at least two propulsors based on the target phase shift until the pairwise phase difference is equal to the reference phase angle.
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
F02K 3/00 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant
75.
Circumferential Row of Vanes for a Gas Turbine Engine and Having Non-Uniform Vane Spacing
A circumferential row of vanes for a gas turbine engine, the circumferential row of vanes has non-uniform spacing. The circumferential row of vanes includes a plurality of stator vanes arranged circumferentially about an inner ring. The plurality of stator vanes include a first group of stator vanes having a first spacing between adjacent stator vanes of the first group of stator vanes and a second group of stator vanes having a second spacing between adjacent stator vanes of the second group of stator vanes. The second spacing is from two percent to eleven percent greater than a nominal uniform vane spacing or from two percent to eleven percent lesser than the nominal uniform vane spacing, the nominal uniform vane spacing being defined by a total number of the plurality of stator vanes. An engine includes the circumferential row of vanes.
Example apparatus, systems, and methods for rapid active clearance control of inter-stage and mid-stage seals are disclosed. An example apparatus to control clearance for a turbine engine comprises a case surrounding at least part of the turbine engine and defining an opening therethrough; a nozzle, the nozzle including a reference pressure sensor and a static pressure sensor on a tip of the nozzle; an actuator including a multilayer stack of material, a rod coupled to the first actuator and coupled to the nozzle through the opening in the case, the rod to move the nozzle based on contraction or expansion of the multilayer stack of material; and a controller to calculate and set the clearance between the rotor and the nozzle by supplying an electrical current to the multilayer stack to cause the multilayer stack to at least one of expand or contract.
A gas turbine engine includes a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining a working gas flowpath, a compressor of the compressor section comprising an aft-most compressor stage; a stage of stator vanes located downstream of the aft-most compressor stage; a stator case including a seal pad; and a spool drivingly coupled to the compressor, the spool and the stator case together defining a rotor cavity in fluid communication with the working gas flowpath, the spool comprising a seal tooth assembly, the seal tooth assembly including a seal support extension, a seal tooth extending from the seal support extension toward the seal pad, and a dampener operable with the seal support extension.
A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.
A composite airfoil assembly for a gas turbine engine. The composite airfoil assembly includes a composite airfoil defined by a core and a skin. The composite airfoil assembly further includes cladding. The core defines a core exterior, where the skin is applied to at least a portion of the core exterior. The cladding is prepared before being coupled or adhered to the composite airfoil.
A propulsion system includes at least two propulsors. The at least two propulsors each include a fan and a controller having one or more processors configured to implement controller logic. The controller logic includes a phase angle control scheme and a speed control scheme. In implementing the controller logic, the one or more processors are configured to: determine an actual pairwise phase difference between a pair of propulsors of the at least two propulsors; generate a reference phase angle for the pair of propulsors; compare the actual pairwise phase difference to the reference phase angle to generate a phase error; provide the phase error to a phase controller module to generate an output based on the phase error; and adjust a speed of at least one propulsor of the at least two propulsors based on the output to drive the phase error towards zero.
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
F02K 3/00 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant
General Electric Company Polska sp. z o.o. (Pologne)
Inventeur(s)
Kumar, Rajesh
Subramanian, Sesha
Raghuvaran, Vaishnav
Ganiger, Ravindra Shankar
Pazinski, Adam Tomasz
Abrégé
A gas turbine engine includes a compressor section comprising a compressor, a combustion section, and a turbine section arranged in serial flow order and defining a working gas flowpath, the compressor comprising an aft-most compressor stage; a spool drivingly coupled to the compressor; a stage of stator vanes located downstream of the aft-most compressor stage; and a stator case, the spool and the stator case together defining a rotor cavity in fluid communication with the working gas flowpath, the stage of stator vanes including a first stator vane defining a fluid passage, and the stator case defining a plenum and a supplemental airflow passage, the plenum in fluid communication with the fluid passage in the first stator vane, the supplemental airflow passage in fluid communication with the plenum and the rotor cavity for proving an airflow to the rotor cavity
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
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 7/141 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel du fluide de travail
82.
COMBUSTION SECTION WITH A PRIMARY COMBUSTOR AND A SET OF SECONDARY COMBUSTORS
A turbine engine with a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. The combustion section including a primary combustor liner having an inner liner and an outer liner. A dome wall and a dome inlet are located in the dome wall. At least one opening is located in the outer liner downstream from the dome inlet. A primary combustion chamber and a set of secondary combustors are fluidly coupled to the primary combustion chamber at the at least one opening.
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
F23R 3/34 - Alimentation de différentes zones de combustion
F23R 3/42 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la disposition ou la forme des tubes à flamme ou des chambres de combustion
An aircraft component inspection system is provided. The system includes an image capture system including: an image sensor system, a positioning system, and a processor configured to determine an inspection recipe based at least on an identifier associated with a component of an aircraft being inspected, identify a plurality of locations for performing image capture during an inspection workflow based on the inspection recipe, provide machine instruction to the positioning system to position the image sensor system relative to the component based on the plurality of locations, cause the image sensor system to capture images at the plurality of locations, and store the images with capture location data in an inspection data database.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
B64F 5/60 - Test ou inspection des composants ou des systèmes d'aéronefs
A printing assembly for an additive manufacturing apparatus includes an energy emitter configured to steer one or more emissions across a build platform to raise a temperature of a build material on the build platform from an initial temperature to a first temperature, the first temperature being less than a threshold temperature set by material dependent metallurgical properties, and a fusing beam emitter configured to generate one or more laser beams to raise the temperature of the build material on the build platform from the first temperature to a second temperature, the second temperature being greater than the threshold temperature.
A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; and an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber, wherein the forward and aft liner segments are coupled together at a moveable interface.
F23R 3/42 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la disposition ou la forme des tubes à flamme ou des chambres de combustion
F02C 3/02 - 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 3/14 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail caractérisés par l'aménagement de la chambre de combustion dans l'ensemble
86.
SYSTEMS AND METHODS FOR ADDITIVELY MANUFACTURING THREE-DIMENSIONAL OBJECTS WITH ARRAY OF LASER DIODES
A system for additively manufacturing a three-dimensional object is provided. The system includes a build platform, an array of laser diodes, each laser diode of the array of laser diodes configured to direct a laser beam toward the build platform, and a controller communicatively coupled to each laser diode of the array of laser diodes such that control signals are communicated from the controller to each laser diode individually.
B29C 64/277 - Agencements pour irradiation utilisant des moyens de rayonnement multiples, p. ex. des micro-miroirs ou des diodes électroluminescentes multiples [LED]
A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
A gas turbine engine including a compressor section and a combustion section in serial flow arrangement along an engine centerline, the combustion section having a combustor liner, a dome wall coupled to the combustor liner, and a dome inlet located in the dome wall, a fuel injector fluidly coupled to the dome inlet, a combustion chamber fluidly coupled to the fuel injector and defined at least in part by the combustor liner and the dome wall, and at least one set of dilution openings located in the dome wall and fluidly coupled to the combustion chamber.
F23R 3/06 - Disposition des ouvertures le long du tube à flamme
F23R 3/16 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la configuration du flux d'air ou du flux de gaz avec des dispositifs à l'intérieur du tube à flamme ou de la chambre de combustion pour influer sur le flux d'air ou de gaz
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for directing incoming objects towards an outer portion of the turbofan engine in communication with the inlet pre-swirl feature.
F02C 7/05 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
F02K 3/00 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant
90.
CLEANING FLUIDS FOR USE IN ADDITIVE MANUFACTURING APPARATUSES AND METHODS FOR MONITORING STATUS AND PERFORMANCE OF THE SAME
Embodiments of the present disclosure are directed to additive manufacturing apparatuses, cleaning stations incorporated therein, and methods of cleaning using the cleaning stations.
B41J 2/165 - Prévention du colmatage des ajutages, p. ex. nettoyage, obturation par un capuchon ou humidification des ajutages
B22F 10/14 - Formation d’un corps vert par projection de liant sur un lit de poudre
B22F 10/38 - Commande ou régulation des opérations pour obtenir des caractéristiques spécifiques du produit, p. ex. le lissage de la surface, la densité, la porosité ou des structures creuses
B22F 12/90 - Moyens de commande ou de régulation des opérations, p. ex. caméras ou capteurs
C09D 9/00 - Produits chimiques pour enlever la peinture ou l'encre
C09D 11/107 - Encres d’imprimerie à base de résines artificielles contenant des composés macromoléculaires obtenus par des réactions faisant intervenir uniquement des liaisons non saturées carbone-carbone à partir d'acides non saturés ou de leurs dérivés
C09D 11/30 - Encres pour l'impression à jet d'encre
91.
METHOD AND DECISION TOOL FOR TRACKING SUSTAINABLE AVIATION FUEL USAGE AND EMISSIONS
A decision tool models emissions for a sustainable aircraft fuel (SAF) for the lifecycle of the fuel. The decision tool includes a controller module configured to receive data related to at least one fuel pathway for the fuel wherein the fuel pathway considers emissions from initial feedstock production to fuel burn during flight and arrival. The decision tool determines at least one fuel pathway for the fuel used to fuel the aircraft during a flight, models an emission score for the at least one fuel pathway, and then outputs the emission score where a user can purchase the fuel or make other decisions based upon the fuel pathway provided by the decision tool.
Methods and apparatus to remove liquid from a housing are disclosed. An example system includes a pump including a chamber, a shaft positioned at least partially in the chamber, and a bearing to support the shaft, the chamber including a chamber inlet and a chamber outlet, the chamber to hold a fluid in a first state, a first conduit to carry the fluid in a second state of the fluid, the first conduit fluidly coupled to the chamber inlet, a second conduit to carry the fluid in the first state of the fluid, the second conduit fluidly coupled to the chamber outlet, at least one jet pump to deliver a mixture of the fluid in the first state of the fluid and the second state of the fluid to a third conduit, and a heat exchanger coupled to the first conduit upstream of the first inlet.
F16N 33/00 - Dispositions mécaniques pour le nettoyage des dispositifs de lubrificationÉgouttoirs ou autres dispositifs particuliers pour débarrasser les parties de machines du lubrifiant
A composite airfoil assembly for a turbine engine. The composite airfoil assembly has an airfoil outer surface defining opposing pressure and suction sides, which extend between a leading edge and a trailing edge. The composite airfoil assembly includes a woven core, a skin applied to at least a portion of a core exterior, and a cladding located adjacent the trailing edge, the leading edge, a root, or a tip, where a portion of a skin outer surface defines a first portion of the airfoil outer surface and a cladding outer surface defines a second portion of the airfoil outer surface.
A composite airfoil assembly for a turbine engine including an airfoil extending in a radial direction between a root and a tip to define a span length. The airfoil including a composite core and a set of skins overlying the composite core. The set of skins include an outer skin defining at least a portion of an exterior surface of the airfoil. A dovetail extends radially below the root, the dovetail including the composite core and the set of skins, the dovetail further including a fan skin overlying the outer skin proximate the root.
A fuel system including: a liquid hydrogen fuel tank for holding a first portion of hydrogen fuel in a liquid phase; a gaseous hydrogen fuel tank for holding a second portion of hydrogen fuel in a gaseous phase; and a fuel delivery assembly including a liquid hydrogen delivery assembly in fluid communication with the liquid hydrogen fuel tank, the liquid hydrogen delivery assembly having a pump for pumping, in the liquid phase, the first portion of hydrogen fuel through the liquid hydrogen delivery assembly; a gaseous hydrogen delivery assembly in fluid communication with the gaseous hydrogen fuel tank, the gaseous hydrogen delivery assembly extending in a parallel arrangement with the liquid hydrogen delivery assembly; and a regulator assembly in fluid communication with both the liquid hydrogen delivery assembly and the gaseous hydrogen delivery assembly for providing gaseous hydrogen fuel to the engine when installed in the vehicle.
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 7/236 - Systèmes d'alimentation en combustible comprenant au moins deux pompes
F02C 9/26 - Commande de l'alimentation en combustible
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p. ex. à la température, à la pression, à la vitesse du rotor
A gas turbine engine is provided, comprising: a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order; and a fan section having a fan drivingly coupled to the turbomachine and an airflow surface rotatable with the fan and exposed to a fan airflow provided to and through the fan during operation of the gas turbine engine, the airflow surface defining a plurality of boundary layer openings configured to ingest a boundary layer of the fan airflow over the airflow surface during operation of the gas turbine engine.
Acoustic inspection systems and methods are described herein useful to determining microstructural characteristics, such as microtexture regions (MTRs), of a material sample. In some embodiments, a method of determining a material characteristic of a material sample includes transmitting acoustic waves from transducer having a concave face that includes one or more piezoelectric elements through a coupling medium and to a surface of a material sample to produce surface acoustic waves along a portion of a surface of the material sample. The one or more piezoelectric elements of the concave face may operate as at least one of an acoustic transmitter or an acoustic receiver. The method also includes receiving the surface acoustic waves reflected from the surface of the material sample at the concave face. The method then includes determining at least one material characteristic of the material sample based on a property of the surface acoustic waves.
G01N 29/28 - Recherche ou analyse des matériaux par l'emploi d'ondes ultrasonores, sonores ou infrasonoresVisualisation de l'intérieur d'objets par transmission d'ondes ultrasonores ou sonores à travers l'objet Détails pour établir le couplage acoustique
Methods and apparatus are disclosed for a hydrogen aircraft with cryo-compressed storage. An example fuel distribution system includes a cryogenic vessel, the cryogenic vessel part of a cryo-compressed hydrogen delivery assembly, a compressed natural gas tank and a thermosiphoning loop to maintain a pressure of the cryogenic vessel using an automatic valve.
B64D 37/30 - Circuits de carburant pour carburants particuliers
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/232 - Soupapes pour combustibleSystèmes ou soupapes de drainage
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p. ex. à la température, à la pression, à la vitesse du rotor
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
A seal assembly for a rotary machine. The seal assembly includes a rotor and a stator. The rotor is rotatable about a rotational axis and has a rotor seal face. The stator has a stator seal face. The stator seal face is positioned opposite the rotor seal face and faces the rotor seal face with a gap therebetween. A portion of one of the rotor and the stator is formed of (i) a shape memory alloy or (ii) a first metal and a second metal with the second metal having a coefficient of thermal expansion different from the first metal. The seal assembly is characterized by a seal clearance compliance ratio (SCCR) from 20% to 90%.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages par obturation non contact, p. ex. du type labyrinthe
F03G 7/06 - Mécanismes produisant une puissance mécanique, non prévus ailleurs ou utilisant une source d'énergie non prévue ailleurs utilisant la dilatation ou la contraction des corps produites par le chauffage, le refroidissement, l'humidification, le séchage ou par des phénomènes similaires
A combustor for a turbine engine includes a combustion chamber including an outer liner and an inner liner, and an annular dome. A plurality of first mixing assemblies includes a pilot mixer and a first main mixer, the first mixing assemblies disposed through the annular dome. The pilot mixer injects a pilot mixer fuel-air mixture axially into a first combustion zone, and the first main mixer injects a first main mixer fuel-air mixture radially into the first combustion zone. A plurality of second mixing assemblies includes a second main mixer, the second mixing assemblies being axially aft of the plurality of first mixing assemblies. The second main mixer injects a second main mixer fuel-air mixture radially into a second combustion zone that is axially aft of, and separate from, the first combustion zone.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
F02C 9/44 - Commande de l'alimentation en combustible sensible à la vitesse de l'aéronef, p. ex. commande du nombre de Mach, optimisation de la consommation en combustible
