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.
F02C 7/28 - Agencement des dispositifs d'étanchéité
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages
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
General Electric Deutschland Holding GmbH (Allemagne)
General Electric Company Polska Sp. z o.o. (Pologne)
Inventeur(s)
Osama, Mohamed
Drozd, Bartlomiej
Yagielski, John Russell
Abrégé
A propulsor is provided including a gas turbine engine having a shaft and one or more bearings supporting the shaft, a rotor hub operatively coupled to the shaft and comprising a hub flange, an electric machine comprising a stator assembly and a rotor assembly, a rotor connection member operatively coupled to the rotor assembly of the electric machine and comprising a connection flange, and an insulated joint for operatively coupling the rotor assembly with the shaft. The insulated joint includes a plurality of insulative layers, at least one of the plurality of insulative layers extending between the hub flange and the connection flange to interrupt common mode electric current from flowing between the rotor assembly and the shaft.
B64D 27/351 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant la récupération d'énergie
B64D 27/357 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles
B64D 27/359 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des condensateurs
B64D 35/024 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride du type en série
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
H02K 7/18 - Association structurelle de génératrices électriques à des moteurs mécaniques d'entraînement, p. ex. à des turbines
H02K 11/02 - Association structurelle de machines dynamo-électriques à des organes électriques ou à des dispositifs de blindage, de surveillance ou de protection pour la suppression des parasites d’origine électromagnétique
H02K 11/21 - Dispositifs pour détecter la vitesse ou la position, ou actionnés par des valeurs de ces variables
H02K 11/40 - Association structurelle à des dispositifs de mise à la terre
H02K 24/00 - Machines adaptées pour la transmission ou réception instantanée du déplacement angulaire de pièces tournantes, p. ex. synchro, selsyn
H03H 1/00 - Détails de réalisation des réseaux d'impédances dont le mode de fonctionnement électrique n'est pas spécifié ou est applicable à plus d'un type de réseau
An additive manufacturing apparatus includes a support configured to support a resin and a constituent material. A support plate includes a window. A stage is configured to hold one or more composite layers of the resin and the constituent material to form a composite component positioned opposite the support plate. A radiant energy device is positioned on an opposite side of the support from the stage and is operable to generate and project radiant energy in a patterned image through the window. An actuator assembly is configured to move the stage in a Z-axis direction and a Y-axis direction.
A gas turbine engine includes a fan section, a compressor section, a combustion section, and a turbine section in serial flow arrangement, and defining an engine centerline extending between a forward direction and an aft direction. A disk includes a slot for mounting a composite airfoil to the disk. An axial retainer couples to the disk and secures the composite airfoil to the disk. A compliant portion positioned at the composite airfoil abuts the composite airfoil during operation of the gas turbine engine to secure the composite airfoil to the disk.
General Electric Company Polska Sp. z o.o. (Pologne)
Inventeur(s)
Nangariyil, Sajinu
Ganiger, Ravindra Shankar
Pillai, Abhilash
Pazinski, Adam Tomasz
Yamarthi, David Raju
Abrégé
A turbine engine is provided. The gas turbine engine defines a radial direction and includes: a rotor; a stator comprising a carrier; a seal assembly disposed between the rotor and the stator, the seal assembly comprising a seal segment, the seal segment having a seal face configured to form a fluid bearing with the rotor; and a seal support assembly, the seal support assembly including a magnet assembly having a magnet coupled to the carrier or the seal segment for biasing the first seal segment along the radial direction.
F01D 11/22 - Réglage actif du jeu d'extrémité des aubes par actionnement mécanique d'éléments du stator ou du rotor, p. ex. par déplacement de sections d'enveloppe par rapport au rotor
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages
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
Provided herein is a calibration block comprising a controlled crack disposed within the volume of a hard to crack material, where the controlled crack has a predetermined location and a predetermined maximum length. Also provided are methods of making a calibration block having a controlled crack. In some aspects, the calibration block comprises a first material and a second material positioned in a volume of the first material. An interface of the first material and the second material has a signal amplitude that is less than about 50% a signal amplitude produced by the controlled crack as detected by an inspection device. The second material includes at least one crack having a predetermined location defined by the position of the second material within the block and predetermined length defined by a size of the second material.
A method for operating a fuel cell assembly, the fuel cell assembly including a fuel cell stack having a solid oxide fuel cell, the solid oxide fuel cell having an anode, a cathode, and an electrolyte, the method including: determining a temperature setpoint for the fuel cell stack, for output products of the fuel cell stack, or both; and controlling a volume of oxidant provided to the anode in response to the determined temperature setpoint to control a temperature of the fuel cell stack, a temperature of the output products of the fuel cell stack, or both.
H01M 8/04111 - Dispositions pour la commande des paramètres des réactifs, p. ex. de la pression ou de la concentration des réactifs gazeux utilisant un assemblage turbine compresseur
B64D 41/00 - Installations génératrices de puissance pour servitudes auxiliaires
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
H01M 4/86 - Électrodes inertes ayant une activité catalytique, p. ex. pour piles à combustible
H01M 8/04014 - Échange de chaleur par des fluides gazeuxÉchange de chaleur par combustion des réactifs
H01M 8/04223 - Dispositions auxiliaires, p. ex. pour la commande de la pression ou pour la circulation des fluides pendant le démarrage ou l’arrêtDépolarisation ou activation, p. ex. purgeMoyens pour court-circuiter les éléments à combustible défectueux
H01M 8/04225 - Dispositions auxiliaires, p. ex. pour la commande de la pression ou pour la circulation des fluides pendant le démarrage ou l’arrêtDépolarisation ou activation, p. ex. purgeMoyens pour court-circuiter les éléments à combustible défectueux pendant le démarrage
H01M 8/04302 - Procédés de commande des éléments à combustible ou des systèmes d’éléments à combustible appliqués pendant des périodes spécifiques appliqués pendant le démarrage
H01M 8/0612 - Combinaison d’éléments à combustible avec des moyens de production de réactifs ou pour le traitement de résidus avec des moyens de production des réactifs gazeux à partir de matériaux contenant du carbone
H01M 8/0637 - Reformage interne direct à l’anode de l’élément à combustible
H01M 8/12 - Éléments à combustible avec électrolytes solides fonctionnant à haute température, p. ex. avec un électrolyte en ZrO2 stabilisé
A gas turbine engine includes a compressor section for compressing air flowing therethrough to provide a compressed air flow, a combustor including a combustion chamber, the combustion chamber configured to combust a mixture of a fuel flow and the compressed air flow to generate combustion products, and a turbine section having at least one turbine driven by the combustion products. The gas turbine engine includes a multicavity damper in fluid communication with the combustion chamber to dampen an instability generated in the combustion chamber by the combustion products. The multicavity damper has a plurality of cavity volumes and the length of each cavity volume is different.
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
9.
INSPECTION SYSTEM AND METHOD USING EDDY CURRENT SENSOR ARRAY
A sensor system may include a sensor array comprising a plurality of eddy current sensor elements, the sensor array having a contact side and a mounting side opposite the contact side. The sensor system may include a sensor mount coupled to the mounting side of the sensor array. The sensor system may include a biasing element configured to bias the contact side of the sensor array against an inspected surface of a component and secure the sensor array to the component while the inspected surface moves relative to the sensor array during data capture.
G01N 27/9093 - Dispositions de support du capteurCombinaisons de capteurs de courants de Foucault et de dispositions auxiliaires pour le marquage ou l’envoi au rebut
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
G01N 27/90 - Recherche ou analyse des matériaux par l'emploi de moyens électriques, électrochimiques ou magnétiques en recherchant des variables magnétiques pour rechercher la présence des criques en utilisant les courants de Foucault
G01N 27/904 - Recherche ou analyse des matériaux par l'emploi de moyens électriques, électrochimiques ou magnétiques en recherchant des variables magnétiques pour rechercher la présence des criques en utilisant les courants de Foucault avec plusieurs capteurs
A fuel injector manifold for a turbine engine includes a fuel manifold ring, a plurality of fuel injectors, and a variable fuel flow system. The fuel manifold flowpath within the fuel manifold ring, the fuel manifold flowpath receiving fuel therein. The plurality of fuel injectors in fluid communication with the fuel manifold flowpath, each of the plurality of fuel injectors having one or more fuel injector flowpaths. The variable fuel flow system disposed within the fuel manifold flowpath, the variable fuel flow system including a closed state, a partially opened state, and a fully opened state to vary a flow of the fuel from the fuel manifold flowpath to the one or more fuel injector flowpaths of each of the plurality of fuel injectors.
General Electric Deutschland Holding GmbH (Allemagne)
Inventeur(s)
Huh, Kum Kang
Datta, Rajib
Rallabandi, Vandana Prabhakar
Yagielski, John Russell
Osama, Mohamed
Abrégé
A power system including a power converter system and an electric machine is provided. In one aspect, the power converter system has first and second switching elements. The electric machine includes a first multiphase winding electrically coupled with the first switching elements and a second multiphase winding electrically coupled with the second switching elements. The first and second multiphase windings are arranged and configured to operate electrically opposite in phase with respect to one another. One or more processors control the first switching elements to generate first pulse width modulated (PWM) signals based on received voltage commands to render a first common mode signal and also control the second switching elements to generate second PWM signals based on received voltage commands to render a second common mode signal. The rendered first and second common mode signals have the same or similar waveform with opposite polarity with respect to one another.
H02M 7/5395 - Transformation d'une puissance d'entrée en courant continu en une puissance de sortie en courant alternatif sans possibilité de réversibilité par convertisseurs statiques utilisant des tubes à décharge avec électrode de commande ou des dispositifs à semi-conducteurs avec électrode de commande utilisant des dispositifs du type triode ou transistor exigeant l'application continue d'un signal de commande utilisant uniquement des dispositifs à semi-conducteurs, p. ex. onduleurs à impulsions à un seul commutateur avec commande automatique de la forme d'onde ou de la fréquence de sortie par modulation de largeur d'impulsions
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 27/12 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz à l'intérieur des ailes ou fixés à celles-ci
H02M 1/44 - Circuits ou dispositions pour corriger les interférences électromagnétiques dans les convertisseurs ou les onduleurs
H02P 25/22 - Enroulements multiplesEnroulements pour plus de trois phases
H02P 27/08 - Dispositions ou procédés pour la commande de moteurs à courant alternatif caractérisés par le type de tension d'alimentation utilisant une tension d’alimentation à fréquence variable, p. ex. tension d’alimentation d’onduleurs ou de convertisseurs utilisant des convertisseurs de courant continu en courant alternatif ou des onduleurs avec modulation de largeur d'impulsions
A gas turbine engine includes a bifurcation and a turbomachine further including an engine cooling system, the engine cooling system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation, a heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly, and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A gas turbine engine is provided, having: a bifurcation; and a turbomachine further including a cooled cooling air (CCA) system, the CCA system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A gas turbine engine having a compressor section for compressing air flowing therethrough to provide a compressed air flow, a combustor including a combustion chamber, the combustion chamber configured to combust a mixture of a fuel flow and the compressed air flow to generate combustion products, a turbine section having at least one turbine driven by the combustion products, and a damper in fluid communication with the combustion chamber to dampen an instability generated in the combustion chamber by the combustion products. The damper defined by nds∝nƒp, ap∝dpn, and ab∝dpν*dpνea, where nds is a number of damper cavities in series, nƒp is a number of discrete frequencies to be damped, ap is an acoustic damping potential, dpn is a neck open area ratio, ab is an acoustic damping broadness, dpν is a damper volume, and dpνea is a damper volume expansion angle.
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes Entrées d'air pour ensembles fonctionnels de propulsion par réaction
A 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
16.
SKIN ASSEMBLY AND METHOD FOR MANUFACTURING THE SAME
A method of constructing a skin assembly comprises forming a first skin panel and a second skin panel; removing material from the first skin panel and the second skin panel, respectively, to form one or more first fingers and one more second fingers. The method also comprises joining the first skin panel and the second skin panel such that the first skin panel and the second skin panel define a plurality of staggered expansion gaps therebetween.
B64C 1/12 - Structure ou fixation de panneaux de revêtement
B32B 3/06 - Caractérisés par des caractéristiques de forme en des endroits déterminés, p. ex. au voisinage des bords pour lier les couches ensembleCaractérisés par des caractéristiques de forme en des endroits déterminés, p. ex. au voisinage des bords pour attacher le produit à quelque chose d'autre p. ex. à un support
B32B 18/00 - Produits stratifiés composés essentiellement de céramiques, p. ex. de produits réfractaires
B64C 1/38 - Constructions adaptées pour réduire les effets de l'échauffement aérodynamique ou d'un échauffement externe d'autre nature
C04B 35/565 - Produits céramiques mis en forme, caractérisés par leur compositionCompositions céramiquesTraitement de poudres de composés inorganiques préalablement à la fabrication de produits céramiques à base de non oxydes à base de carbures à base de carbure de silicium
C04B 35/80 - Fibres, filaments, "whiskers", paillettes ou analogues
C04B 41/91 - Post-traitement des mortiers, du béton, de la pierre artificielle ou des céramiquesTraitement de la pierre naturelle de céramiques uniquement impliquant l'enlèvement d'une partie des matières des objets traités, p. ex. par attaque chimique
C04B 111/00 - Fonction, propriétés ou utilisation des mortiers, du béton ou de la pierre artificielle
A power system operable to implement a power balancing control scheme is provided. In one aspect, a power system includes multiple independent power supplies with independent batteries feeding onto a common power bus. The power supplies regulate the voltage on the common power bus at the same time. The power balancing control scheme, when implemented, causes the load on the common power bus to be shared among the individual power supplies with a specified load distribution. The specified load distribution can be set or determined to balance the State of Charge (SoC) of the batteries over time whilst taking into account the constraints or limits of the elements of the power system.
Embodiments of a propulsion system are provided herein. In some embodiments, a propulsion system for an aircraft may include an electrical power supply; a motor coupled to the electrical power supply, wherein the electrical power supply provides power to the motor; and a fan disposed proximate a rear portion of an aircraft and rotatably coupled to the motor, wherein the fan is driven by the motor.
B64D 27/355 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles à combustible
B64D 27/357 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles
B64D 27/359 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des condensateurs
B64D 35/024 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride du type en série
A propulsion system for an aircraft can include an electric power source and an propulsion assembly having a propulsor. An electric power bus can electrically connect the electric power source to the propulsion assembly. The electric power source can be configured to provide electrical power to the electric power bus. An inverter converter controller can be positioned along the electric power bus and can be electrically connected to the electric power source at a location downstream of the electric power source and upstream of the electric propulsion assembly.
B60L 50/10 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant la puissance de propulsion fournie par des générateurs entraînés par le moteur, p. ex. des générateurs entraînés par des moteurs à combustion
B64C 21/06 - Moyens permettant d'influencer l'écoulement d'air sur les surfaces des aéronefs en agissant sur la couche limite par utilisation de fentes, de conduits, de surfaces poreuses ou de dispositifs similaires en vue de l'aspiration
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 27/12 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz à l'intérieur des ailes ou fixés à celles-ci
B64D 27/18 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à réaction à l'intérieur des ailes ou fixés à celles-ci
B64D 27/32 - Aéronefs caractérisés par des groupes moteurs électriques à l'intérieur des fuselages ou fixés à ceux-ci
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
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 method of operating a rotating detonation combustor includes providing a flow of air through an air inlet to flow into a detonation chamber, providing a flow of fuel from at least one fuel injector into the detonation chamber, mixing the flow of the fuel and the flow of the air in the detonation chamber to generate a fuel-air mixture, detonating the fuel-air mixture in the detonation chamber to generate rotating detonation waves within the detonation chamber, and controlling, during operation of the rotating detonation combustor from a first power operating state to a second power operating state, different from the first power operating state, the air inlet wall to control the flow of the air through the air inlet into the detonation chamber to control a discharge coefficient and an operating mode within the detonation chamber.
F23R 7/00 - Chambres de combustion à combustion intermittente ou explosive
F02C 5/02 - Ensembles fonctionnels de turbines à gaz caractérisés par un fluide énergétique produit par une combustion intermittente caractérisés par l'aménagement de la chambre de combustion dans l'ensemble
F23R 3/38 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible comprenant des moyens d'injection de combustible rotatifs
F23R 3/56 - Chambres de combustion comportant des tubes à flamme rotatifs
A gas turbine engine including a turbomachine having a cooled cooling air (CCA) system is provided. The CCA system includes a cold side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section at a location through a low pressure compressor, between the low pressure compressor and a high pressure compressor, or both; a CCA heat exchanger in thermal communication with the cold side bleed assembly; and a hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
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
22.
SYSTEM AND METHOD FOR FORMING A CERAMIC SANDWICH-STRUCTURED COMPOSITE COMPONENT
A method for forming a ceramic sandwich-structured composite component includes applying one or more ceramic matrix composite (CMC) plies to a surface of a tool, and positioning a ceramic foam precursor with respect to the tool such that the ceramic foam precursor and the one or more CMC plies is disposed in a sealed interior cavity of the tool. The ceramic foam precursor and the one or more CMC plies are cured within the sealed interior cavity to form a green preform such that the ceramic foam precursor expands within the sealed interior cavity to apply a pressure to the one or more CMC plies to consolidate the one or more CMC plies. The expanded ceramic foam precursor forms a ceramic core of the green preform bonded to the one or more CMC plies. The green preform is removed from the tool and sintered to ceramify the green preform.
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 method of inspecting a composite component having a plurality of composite plies.
A method of inspecting a composite component having a plurality of composite plies.
The method includes capturing an image of a cross section of the plurality of composite plies, the image including light reflection patterns corresponding to a plurality of reinforcing fibers within the composite plies of the composite component, identifying key points of the composite component on the image to define an area of interest, generating a plurality of profile lines in the image, processing along each of the plurality of profile lines to identify patterns corresponding to a layup of the plurality of reinforcing fibers, grouping identified patterns from the plurality of profile lines, comparing the identified patterns to an as-expected template of the plurality of composite plies in the composite component to evaluate a manufacturing process of the composite component, and determining whether the identified patterns match the as-expected template.
A ceramic matrix composite (CMC) component and method of forming the CMC component with multiple layers of impregnated matrix fibers, at least one of the multiple layers including at least one CMC prepreg formed from a plurality of twisted tows, twisted through at least one turn/meter.
A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having comprising a cooled cooling air (CCA) system, the CCA system comprising: a hot side bleed assembly; a CCA heat exchanger in thermal communication with the hot side bleed assembly; and a cold side bleed assembly defining an inlet in fluid communication with a cold location of the gas turbine engine and an outlet, the cold side bleed assembly in thermal communication with the CCA heat exchanger for cooling an airflow through the hot side bleed assembly, the outlet of the cold side bleed assembly in fluid communication with the working gas flowpath at a location downstream of the turbine section, with a core cowl vent of the turbomachine, with a passage over the turbomachine, or a combination thereof.
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 gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a cooled cooling air (CCA) system, the CCA system including: a CCA heat exchanger; a cold side bleed assembly; a hot side bleed assembly; and a valve assembly operable with the cold side bleed assembly and with the hot side bleed assembly, the valve assembly structured to modulate airflows through the cold side bleed assembly and through the hot side bleed assembly in tandem.
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
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
28.
METHOD OF MANUFACTURING A COMPOSITE COMPONENT FOR A GAS TURBINE ENGINE
A method of manufacturing a composite component having an outer shell, an inner hub, and a plurality of struts connecting the outer shell and the inner hub. An outer shell hoop preform and an inner hub hoop preform are woven and installed on a mold tooling structure, a plurality of outer shell pi-joint members are integrally woven to the outer shell hoop preform, and a plurality of inner hub pi-joint members are integrally woven to the inner hub hoop preform. A plurality of strut preforms are connected between respective ones of the outer shell pi-joint members and the inner hub pi-joint members. A matrix material is injected into the mold tooling structure and a curing process is applied to the mold tooling structure to obtain the composite component.
B29C 70/24 - Façonnage de matières composites, c.-à-d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p. ex. des inserts comprenant uniquement des renforcements, p. ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins trois directions formant une structure tridimensionnelle
B29K 105/08 - Présentation, forme ou état de la matière moulée contenant des agents de renforcement, charges ou inserts de grande longueur, p. ex. ficelles, mèches, mats, tissus ou fils
General Electric Company Polska sp. z o.o. (Pologne)
Inventeur(s)
Ganiger, Ravindra Shankar
Prasad, Santosh Kumar
Kuropatwa, Michal Tomasz
Abrégé
Compressor bleed slots with variable wall structures are disclosed herein. An example apparatus disclosed herein is to be coupled to a wall of a bleed slot of a compressor of a gas turbine engine, the bleed slot defining a flow path, the apparatus comprising a member to be coupled to the wall, and a plate coupled to the member, the plate having a first geometry at a first ambient condition at a first time, the flow path having a first area when the plate has the first geometry, and a second geometry at a second ambient condition at a second time, the flow path having a second area when the plate has the second geometry, the first area greater than the second area, the first time after the second time.
F04D 27/00 - Commande, p. ex. régulation, des pompes, des installations ou des systèmes de pompage spécialement adaptés aux fluides compressibles
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
F04D 29/56 - Moyens de guidage du fluide, p. ex. diffuseurs réglables
30.
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.
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
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
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
31.
TURBOMACHINERY ENGINES WITH HIGH-SPEED LOW-PRESSURE TURBINES
A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine includes four rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 2.0-5.1. Additionally (or alternatively) the low-pressure turbine includes an area-EGT ratio within a range of 1.05-1.6.
F04D 29/32 - Rotors spécialement adaptés aux fluides compressibles pour pompes à flux axial
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux comprenant une soufflante avant
An insertion tool for performing an operation on equipment, the insertion tool including: a plurality of segments, each segment of the plurality of segments including a body comprising: a first hinge; and a second hinge, the first hinge of a first segment being coupled to the second hinge of a second segment adjacent to the first segment through an interface, wherein the interface comprises a powder gap, a multi-modal interface, a compliance feature, a displace-to-lock configuration, an interference fit, or any combination thereof, wherein the insertion tool is configured to be selectively rigidizable using a strength member interfacing with the plurality of segments.
Outlet guide vane mounts with an airfoil structure are disclosed herein. The airfoil structure includes an airfoil platform, a composite material, a load spreader block, and a shim material, the composite material and a load spreader block positioned on the airfoil platform, the composite material having a first stiffness and the load spreader block having a second stiffness, and a shim material positioned between the composite material and the load spreader block, the shim material having a third stiffness different from at least one of the first stiffness or the second stiffness.
A turbine engine has a compression section, a combustion section, a turbine section in serial flow arrangement and defining a working airflow path. The combustion section has a circumferential casing, a combustor, and a fuel nozzle. The circumferential casing defines an interior. The turbine engine includes a deswirler assembly. The deswirler assembly couples the working airflow path of the compression section to the interior. The turbine engine includes a fuel supply system.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
F01D 9/06 - Conduits d'admission du fluide à l'injecteur ou à l'organe analogue
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 7/22 - Systèmes d'alimentation en combustible
F02C 7/228 - Division du fluide entre plusieurs brûleurs
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
09 - Appareils et instruments scientifiques et électriques
35 - Publicité; Affaires commerciales
37 - Services de construction; extraction minière; installation et réparation
41 - Éducation, divertissements, activités sportives et culturelles
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Aircraft engines, aircraft gas turbine engines, aircraft turbo-prop engines, aircraft turbo-jet engines, rocket engines not for land vehicles and replacement parts, fittings, and accessories for the foregoing; additive manufacturing machines being 3d printers; propulsion systems for aeronautical vehicles comprised of aeronautic engines, turbines other than for land vehicles, centrifugal pumps, electric pumps, industrial chemical reactors, nacelles, thrust reversers, engine air intake assembly, engine fan casings, exhaust cone in the nature of engine exhaust tips; aeronautical vehicle propulsion system components, namely, aeronautic engines, turbines other than for land vehicles, electric pumps, thrusters for machinery, industrial chemical reactors, nacelles, thrust reversers, engine air intake assemblies, engine fan casings, and exhaust cones in the nature of engine exhaust tips; replacement parts for aeronautical vehicle propulsion systems; cyclone separators; Pneumatic starters for reciprocating and turbine engines for industrial uses and replacement parts thereof Aviation sensor systems, namely, navigation systems for aircraft, aircraft altitude indicators, aircraft speed indicators; flight control systems, namely, apparatus for checking flight parameters; downloadable software for operating engines and navigation systems, and for data analysis in the aviation industry; downloadable software for use in operating machinery used the additive manufacturing industry; unmanned aircraft systems, namely, aircraft engines; unmanned traffic management systems, namely, downloadable computer software for managing operations in the field and automated traffic light apparatus; downloadable software for controlling the operation of 3d printers and machines for additive manufacturing; downloadable cloud-based software for controlling the operation of 3d printers and machines for additive manufacturing; downloadable avionic flight management software for checking flight parameters, namely, systems for maintaining efficient cruise altitude and systems for achieving lower fuel costs, reduced emissions, and less noise; cockpit voice and data recorder Business consulting services in the aviation and additive manufacturing industry; inventory management for engines; sustainability consulting services in the aviation and additive manufacturing industry; fuel management services in the aviation industry; business advisory services relating to organization, administrative and commercial management of industrial or commercial companies with an air fleet; business advisory services relating to administrative management of engine fleets, systems, equipment and parts of aircraft; compilation and statistical study of data relating to the management of a fleet of systems, equipment and parts for aeronautical and space vehicles and management of maintenance of an aircraft fleet; collection of data in a central computer database file; management and compilation of databases; analysis of commercial data, and business data gathering, systematization, management and processing; exploitation of statistical data, namely, statistical analysis and reporting services for business purposes; provision of statistical information in a computer database for business purposes; provision and compilation of statistical data for business purposes; commercial project management services in the framework of construction projects; all these services used and/or intended for the aeronautical and space field; Customer loyalty program for commercial, promotional, and/or advertising purposes for customers in the commercial aviation industry; Consumer product information, namely, providing configuration data and engine information regarding the OEM assessment of operational history, namely, parts, repairs, configurations, and maintenance practices, for specifically identified aircraft engines and gas turbine engines for the purpose of enabling consumers to make informed purchasing decisions Technical support services, namely, providing technical advice in the field of repair, maintenance, modification and overhaul services for airplanes, aircraft engines, gas turbine engines and rocket engines and replacement parts; providing configuration data and engine information data for airplanes, aircraft engines and gas turbine engines regarding operational, servicing and repair history; services of repair, maintenance, and reconditioning of propulsion systems for aerospace vehicles, engines, thrusters, nacelles, thrust reversers and their components; under-wing repair, mechanical repair revision, servicing and maintenance of all kinds of propulsion systems for aeronautical vehicles and their component parts; standardization, namely, refurbishment and standard replacement for repair of propulsion systems for aeronautical vehicles, engines, thrusters, nacelles, thrust reversers and their component parts; consultancy and assistance services relating to the repair, mechanical revision repair, servicing, standardization, namely, mechanical repair and maintenance of propulsion systems for aeronautical vehicles, engines, thrusters, nacelles, thrust reversers and their component parts Training services in the field of aviation and the additive manufacturing industry Engineering design services for additive manufacturing and aviation industries; research and development for additive manufacturing and aviation industries; analysis of technical data, technical advice and technical project study, all in the field of aeronautics; design and development of software and computer programming; inspection of aircraft engines; prototyping services, namely, new product design, development and testing in the field of additive manufacturing management; providing online non-downloadable software for operating engines and navigation systems, and for data analysis in the aviation industry; Providing temporary use of non-downloadable software for controlling the operation of 3D printers and machines for additive manufacturing; Software as a service (SAAS) featuring software for managing flight operations, aviation technical operations, aircraft and aircraft engine maintenance, and aviation analytics
A gas turbine engine includes a compressor section, combustion section, and turbine section is serial flow arrangement. A fuel injector supplies a mixture of fuel and air for combustion within the combustion section. An outer wall defines a mixing passage extending along a stream-wise direction including a first mixing region and a second mixing region. A first fuel passage supplies a first fuel to the first mixing region and an air passage supplies a supply of air to the first mixing region. A second fuel passage supplies a second fuel to the second mixing region.
F23D 14/36 - Brûleurs spécialement conçus pour être utilisés avec des moyens comprimant le combustible gazeux ou l'air de combustion dans lesquels le compresseur et le brûleur forment une unité
F23D 14/64 - Dispositifs mélangeursTubes mélangeurs avec injecteurs
F23D 14/70 - Chicanes ou dispositifs analogues pour créer des turbulences
F23D 14/84 - Diffusion de la flamme ou autres moyens pour lui donner une forme particulière
F23D 17/00 - Brûleurs pour la combustion simultanée ou alternative de combustibles gazeux, liquides ou pulvérulents
37.
FLUID MANAGEMENT SYSTEM AND METHODS FOR ADDITIVE MANUFACTURING SYSTEMS
The present disclosure relates to fluid management systems and methods of calibrating the same within additive manufacturing systems. A fluid management system includes a pump and at least one fluid circuit, each fluid circuit comprising a plurality of fluid pathways, each of the plurality of fluid pathways comprising at least one flow-regulating valve and at least one actuating valve. At least a portion of each of the plurality of fluid pathways are fluidly connected by at least one actuating valve. The pump is operable to provide a fluid to each of the plurality of fluid pathways, wherein the fluid has a flowrate within each of the plurality of fluid pathways. Each flow-regulating valve is adjustable to increase or decrease the flowrate of each of the plurality of fluid pathways such that each flowrate of the plurality of fluid pathways is substantially the same.
An insertion tool is provided. The tool includes a flexible section a rigidization actuator configured to actuate the flexible section between a rigidized state and a relaxed state, wherein in the rigidized state, a fluid path is formed within the flexible section, an end effector coupled to the flexible section, and an end effector actuator comprising at least one pneumatic turbine configured to transform fluid force of a fluid from the fluid path into torque to drive a rotation of the end effector.
A turbine engine seal configured for use between a turbine engine rotor and a turbine engine static component of a turbine engine can include a seal construction having a negative thermal expansion (NTE) layer located on one or both of the turbine engine rotor and turbine engine static component. The NTE layer can include a NTE reactive component comprising a material with a negative thermal expansion coefficient. When the turbine engine rotor rubs against the turbine engine static component, heat is generated and the NTE reactive component can experience an increase in temperature from a first temperature to a second temperature. The increase in temperature causes a dimension of the NTE reactive component to decrease which consequently forms a hydrodynamic pocket useful to generate a lift force that urges separation between the turbine engine rotor and turbine engine static component. The seal construction can include a lattice compliant layer.
A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). The high-pressure compressor defines a high-pressure compressor exit area (AHPCExit) in square inches. The gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT/(AHPCExit2×1000). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.
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
41.
COMPOSITE COMPONENTS AND METHODS FOR PREVENTING FLOW FROM INFILTRATED COMPONENT DURING RE-INFILTRATION
A composite component may include an infiltrated segment infiltrated with a molten material during a prior infiltration process, a green segment that is uninfiltrated, and a barrier segment having a microstructure different from the infiltrated segment, the green segment, or both. The microstructure of the barrier segment may be configured to slow a flow of material between the infiltrated segment and the green segment during a subsequent infiltration process.
A fuel nozzle has a fuel nozzle body, a set of fuel jets, and a compressed air resonator. The fuel nozzle body has a central channel defining a channel centerline. The set of fuel jets extend through the fuel nozzle body. The set of fuel jets are fluidly coupled to the central channel to define an injecting section of the central channel. The compressed air resonator defines a resonating section of the compressed air channel.
A combustor for a turbine engine includes a wall defining a combustion chamber, and an acoustic damper. The acoustic damper includes a housing in fluid communication with the combustion chamber through an opening provided in the wall, the housing defining a cavity and having one or more neck holes in fluid communication with the combustion chamber, and a mechanism configured to vary a damping acoustic frequency of the acoustic damper so as to vary the damping acoustic frequency of the acoustic damper to align with an acoustic frequency of acoustic vibrations generated in the combustion chamber to attenuate an acoustic instability within the combustion chamber.
A lubrication system for a turbine engine includes a reservoir that stores a lubricant, a primary lubricant supply circuit including a primary supply pump fluidly coupled to the reservoir, and an auxiliary lubricant supply circuit including an auxiliary supply pump fluidly coupled to the reservoir. A clutch is mechanically coupled to the auxiliary supply pump, and the clutch configured to engage the auxiliary supply pump and a shaft of the turbine engine when activated. The lubrication system further includes a pressure sensor that monitors a lubricant pressure within the primary lubricant supply circuit. When the lubricant pressure within the primary lubricant supply circuit falls below a predetermined lubricant threshold, the clutch is activated to engage the auxiliary supply pump and the shaft of the turbine engine.
An insertion tool is provided. The tool includes a flexible section comprising a plurality of rigidizable links, an end effector actuator, and an end effector coupled to the flexible section. A flexible is shaft inserted through the flexible section, wherein torque is transferred from the end effector actuator to the distal end via the flexible shaft to cause a rotation of the end effector. A tool may include a tool-less disconnect interface between the end effector and the flexible shaft.
A gas turbine engine includes a turbomachine having an engine core including a high-pressure compressor, a combustion section, a high-pressure turbine, and a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine. The engine core has a length (LCORE), and the high-pressure compressor has an exit stage diameter (DCORE). The high-pressure compressor defines a high-pressure compressor exit area (AHPCExit) in square inches. The gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT/(AHPCExit2×1000). The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of LCORE/DCORE is from 2.1 to 4.3.
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 turbine engine includes a cooling air duct for cooling air positioned radially between a core air flow path for core air and a bypass airflow passage for bypass air. A heat exchanger is positioned in the cooling air duct to transfer heat from a heat source from within the turbine engine. The heat exchanger may be a condenser. The turbine engine may further include a steam system that extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flow path, the steam system including the condenser to transfer heat from the combustion gases to the cooling air and to condense the water from the combustion gases. The turbine engine may further include a booster fan to increase the pressure of the cooling air and the core air.
F02C 3/30 - Addition d'eau, de vapeur ou d'autres fluides aux composants combustibles ou au fluide de travail avant l'échappement de la turbine
F01D 25/32 - Recueil de l'eau de condensationDrainage
F02K 3/077 - 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 l'ensemble fonctionnel étant du type multi-flux, c.-à-d. ayant au moins trois flux
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
An engine component for a turbine engine. The engine component has a composite structure and a cover structure. The composite structure has a composite structure outer wall, a composite structure edge, and a channel. The channel is provided along the composite structure edge. The cover encases at least a portion of the composite structure outer wall. The cover has a main body and an extension.
A turbomachine comprises a nozzle segment including an inner shroud defining a bottom surface and a nozzle flange defining a forward side surface and an aft side surface. A floating rotor seal is coupled to the nozzle flange via a carrier flange. The carrier flange includes a forward wall and an aft wall. The nozzle flange is positioned between the forward and aft walls and a flowpath is defined therebetween. A seal pocket is defined in one of the forward wall or the aft wall and is in fluid communication with the flowpath. At least one linear seal segment is partially disposed within the seal pocket. The linear seal segment is configured to form a seal against the nozzle flange or the bottom surface in response to pressurization of the seal pocket via a working fluid in the flowpath.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages
F01D 11/16 - Régulation ou commande du jeu d'extrémité des aubes, c.-à-d. de la distance entre les extrémités d'aubes du rotor et le corps du stator par des moyens auto-réglables
Collars to retain multiple fan blades for use with an aircraft engine are disclosed herein. An example gas turbine engine includes a plurality of fan blades and a collar coupled to the plurality of fan blades. The collar includes a first mating piece and a second mating piece coupled to the first mating piece. The first and second mating pieces define an opening configured to receive a first root of a first fan blade of the plurality of fan blades, and the first and second mating pieces define a slot to retain a second root of a second fan blade of the plurality of fan blades.
A variable pitch airfoil assembly for an engine includes a disk having an annular shape extending about an axial direction, an airfoil coupled to the disk via a platform, and at least one damping element disposed between the platform and the disk. The airfoil extends outwardly from the disk in a radial direction and is rotatable relative to the disk about a pitch axis. The at least one damping element is configured to provide vibration damping by friction between the at least one damping element, the disk, and the platform while also allowing for a pitch change of the airfoil.
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
A liner for a combustion section of a gas turbine engine includes a base portion and a stiffening portion. The base portion includes a plurality of plies of a composite material including a first ply having a fiber direction aligned with a circumferential direction of the liner and a second ply adjacent to the first ply, the second ply having a fiber direction angled away from the circumferential direction. The stiffening portion is disposed on the base portion and includes a plurality of plies of the composite material including a first ply having a fiber direction aligned with the circumferential direction, a second ply adjacent to the first ply the second ply having a fiber direction aligned with the circumferential direction, and a third ply adjacent to the second ply, the third ply having a fiber direction angled away from the circumferential direction.
A method for predicting build line locations in a part before additively manufacturing the part, includes obtaining a sliced three-dimensional model of a part for additive manufacturing, generating, for each neighboring pair of layers in the plurality of layers, a face count difference, generating, for each of the neighboring pair of layers, a surface area difference, predicting, for each of the neighboring pair of layers, that the first layer from each of the neighboring pair of layers comprises a presence of a build line based on a determination that the face count difference is less than zero and the surface area difference is greater than zero; storing a list of predicted build line layers comprising the one or more layers predicted to comprise the presence of the build line; and adjusting dimensions of the part for additive manufacturing corresponding to a layer in the list of predicted build line layers.
An insertion tool includes a housing, an elongated section at least partially within the housing, a bendable section coupled to the elongated section at, and an actuator. The actuator is configured to actuate the bendable section, via causing an axial displacement of the elongated section within the housing, from a retracted state at least partially positioned within the housing to an extended state outside of the housing. The insertion tool also includes a tensioning assembly configured to tension the bendable section into a predefined shape in the extended state. The elongated section is coupled to the housing via a rotation interface configured to cause a rotation of the bendable section during the actuation of the bendable section from the retracted state to the extended state.
A gear assembly for use with a turbomachine comprises a sun gear, a plurality of planet gears, and a ring gear. The gear assembly is connected to an input shaft and an output shaft. The sun gear is configured to rotate about a longitudinal centerline of the gear assembly, and is driven by the input shaft. A component of the gear assembly drives the output shaft. The gear assembly further comprises an output shaft reversal mechanism configured to reverse the rotational direction of the output shaft.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
A variable area turbine nozzle assembly includes a guide vane including an outer centering pin defining a tab. An inner support ring is spaced radially outward from the guide vane and defines an opening and a protrusion. The protrusion is configured to engage with the tab of the outer centering pin. An outer support ring extends circumferentially around the inner support ring and defines an aperture. The outer support ring has a second coefficient of thermal expansion that is greater than or less than the first coefficient of thermal expansion. At least one linkage joins the inner support ring to the outer support ring and is configured to rotate the inner support ring circumferentially about an axial centerline of the variable area turbine nozzle assembly in response to a change in operational temperature of a combustion gas thus causing the guide vane to rotate.
Additive manufacturing apparatuses, components of additive manufacturing apparatuses, and methods of using such manufacturing apparatuses and components are disclosed. An additive manufacturing apparatus may include a recoat head for distributing build material in a build area, a print head for depositing material in the build area, one or more actuators for moving the recoat head and the print head relative to the build area, and a cleaning station for cleaning the print head.
B29C 64/165 - Procédés de fabrication additive utilisant une combinaison de matériaux solides et liquides, p. ex. une poudre avec liaison sélective par liant liquide, catalyseur, inhibiteur ou absorbeur d’énergie
B22F 10/14 - Formation d’un corps vert par projection de liant sur un lit de poudre
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
B22F 12/17 - Moyens de chauffage auxiliaires pour chauffer la chambre ou la plate-forme de formation
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
60.
APPARATUS AND SYSTEMS FOR SEPARATING PHASES IN LIQUID HYDROGEN PUMPS
Methods, apparatus, systems, and articles of manufacture are disclosed herein that include a cryogenic pump system comprising: a cryogenic liquid tank; a cryogenic pump including a suction adapter, the suction adapter connected to the cryogenic liquid tank via a liquid supply line and a gaseous return line; and a phase separator connected downstream of the cryogenic liquid tank and upstream of the cryogenic pump, the phase separator including a filtration structure integrated into the liquid supply line to separate vapor from cryogenic liquid, the phase separator connected to the gaseous return line to direct the vapor to the cryogenic liquid tank.
B01D 39/20 - Autres substances filtrantes autoportantes en substance inorganique, p. ex. papier d'amiante ou substance filtrante métallique faite de fils métalliques non-tissés
F17C 1/00 - Récipients sous pression, p. ex. bouteilles de gaz, réservoirs de gaz, cartouches échangeables
A coating including a plurality of indicator oxide nanoparticles, a binder, and a wetting agent. A sulfidation corrosion mitigation coating including: a sulfidation corrosion mitigation material, a binder, and a plurality of indicator oxide nanoparticles. An article including a metal alloy substrate having the sulfidation corrosion mitigation coating thereon is also provided. The sulfidation corrosion mitigation coating can include a first indicator layer containing indicator oxide nanoparticles disposed on the surface of the metal alloy substrate. Methods for inspection of an article having a coating containing a plurality of indicator oxide nanoparticles is also provided.
C09D 1/00 - Compositions de revêtement, p. ex. peintures, vernis ou vernis-laques, à base de substances inorganiques
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
Jet engine thermal transport bus pumps are disclosed. Disclosed herein is an aircraft comprising a gas turbine engine configured to burn fuel at a fuel flow rate to generate an engine power (Pengine), the fuel characterized by a first specific heat capacity (cp_fuel) and a net heat of combustion (NHCfuel); and a thermal management system configured to transfer heat from a working fluid to a heat sink fluid, the working fluid characterized by a second specific heat capacity (cp_pump) and a first density (ρpump), the thermal management system including a pump configured to generate a pump power (Ppump) to pressurize the working fluid, and wherein
Jet engine thermal transport bus pumps are disclosed. Disclosed herein is an aircraft comprising a gas turbine engine configured to burn fuel at a fuel flow rate to generate an engine power (Pengine), the fuel characterized by a first specific heat capacity (cp_fuel) and a net heat of combustion (NHCfuel); and a thermal management system configured to transfer heat from a working fluid to a heat sink fluid, the working fluid characterized by a second specific heat capacity (cp_pump) and a first density (ρpump), the thermal management system including a pump configured to generate a pump power (Ppump) to pressurize the working fluid, and wherein
POW
=
P
pump
(
c
p
_
pump
c
p
_
water
)
(
ρ
water
ρ
pump
)
2
,
FFR
=
(
P
engine
N
H
C
fuel
)
(
c
p
_
fuel
c
p
_
pump
)
,
Jet engine thermal transport bus pumps are disclosed. Disclosed herein is an aircraft comprising a gas turbine engine configured to burn fuel at a fuel flow rate to generate an engine power (Pengine), the fuel characterized by a first specific heat capacity (cp_fuel) and a net heat of combustion (NHCfuel); and a thermal management system configured to transfer heat from a working fluid to a heat sink fluid, the working fluid characterized by a second specific heat capacity (cp_pump) and a first density (ρpump), the thermal management system including a pump configured to generate a pump power (Ppump) to pressurize the working fluid, and wherein
POW
=
P
pump
(
c
p
_
pump
c
p
_
water
)
(
ρ
water
ρ
pump
)
2
,
FFR
=
(
P
engine
N
H
C
fuel
)
(
c
p
_
fuel
c
p
_
pump
)
,
0.008≤POW/FFR5/3≤12, FFR is between 0.05 pounds-mass per second and 16 pounds-mass per second, and ρwater and cp_water is the density and specific heat capacity of water, respectively.
A rotor blade for a gas turbine engine is provided. The rotor blade includes a blade body formed of a first material; and a spar within a portion of the blade body, the spar formed of a second material that is different than the first material, the spar having an elongate body including a notch. The notch, weakened geometric feature, or other reduction in cross-section defines a frangible portion of the spar that is used to control a fracture of a rotor blade.
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
Turbomachine (1) comprising an unducted propeller (14) propelling a tertiary flow (13), a fan (12) and a compressor (4) compressing a primary flow (F1), as well as an annular passage (19) for the flow of a secondary flow (F2) downstream of the fan (12); the annular passage (19) accommodating an annular row of rectifier vanes (22) and at least one heat exchanger (24) downstream of the row of vanes (22); a plurality of diffusion corridors being provided upstream of the at least one exchanger (24), each corridor being delimited circumferentially by an intrados and by an extrados of two circumferentially adjacent vanes (22), and by at least one fin carried by at least one of the two circumferentially adjacent vanes (22).
A gas turbine engine is provided, including: an accessory system; and a turbomachine comprising a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section collectively defining in part a working gas flowpath, the turbomachine further including: a reverse bleed system comprising a reverse bleed duct and an RBS blower in fluid communication with the reverse bleed duct, the reverse bleed duct in fluid communication with the working gas flowpath; and an accessory cooling system including a cooling duct defining an inlet in fluid communication with the reverse bleed duct, the accessory cooling duct including a cooling tip oriented towards the accessory system to provide an airflow onto the accessory system.
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 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A gas turbine engine is provided. The gas turbine engine includes: a turbomachine comprising a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section collectively defining in part a working gas flowpath, the turbomachine further including; a reverse bleed system comprising a reverse bleed duct and an RBS blower in fluid communication with the reverse bleed duct, the reverse bleed duct in fluid communication with the working gas flowpath; and an active clearance control (ACC) system including an inlet, a heat transfer assembly arranged around the turbine of the turbine section, and an ACC duct assembly extending from the inlet to the heat transfer assembly, the inlet of the ACC system in fluid communication with the reverse bleed duct.
A variable area turbine nozzle assembly includes a guide vane including an outer centering pin defining a tab. An inner support ring is spaced radially outward from the guide vane and defines an opening and a protrusion. The protrusion is configured to engage with the tab of the outer centering pin. An outer support ring extends circumferentially around the inner support ring and defines an aperture. The outer support ring has a second coefficient of thermal expansion that is greater than or less than the first coefficient of thermal expansion. At least one linkage joins the inner support ring to the outer support ring and is configured to rotate the inner support ring circumferentially about an axial centerline of the variable area turbine nozzle assembly in response to a change in operational temperature of a combustion gas thus causing the guide vane to rotate.
A disconnector system for disconnecting a drive shaft of a drive mechanism from rotating equipment, upon a failure of the drive mechanism or rotating equipment, includes a disconnector mechanism having a disconnector shaft disposed in a casing and moveable relative thereto, between a first position and a second position, and a cam surface on a distal end of the arm configured to engage a slidable coupler. The movement of the disconnector shaft can be triggered by an operation of a solenoid, or by a displacement of the solenoid responsive to a melting of a meltable element.
F16D 9/02 - Accouplements avec organe de sécurité pour le désaccouplement par des moyens thermiques, p. ex. un élément fusible
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
69.
GAS TURBINE ENGINE WITH FORWARD SWEPT OUTLET GUIDE VANES
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, the fan comprising a plurality of fan blades; a turbomachine drivingly coupled to the fan, the turbomachine comprising a compressor section with a low pressure compressor, a turbine section with a low pressure turbine, a reduction gearbox, and an outer casing, the low pressure turbine drivingly coupled to the low pressure compressor across the reduction gearbox; 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 at a location downstream of the plurality of fan blades, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip.
F02C 3/045 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur les passages du compresseur et de la turbine se trouvant sur un même rotor
B64C 33/02 - AilesMécanismes d'actionnement des ailes
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c.-à-d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p. ex. du type à double flux comprenant une soufflante avant
70.
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.
F02K 3/115 - Chauffage du flux dérivé à l'aide d'un échange indirect de chaleur
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 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
71.
Methods and apparatus to control a surface of an aircraft engine
Methods and apparatus to control a surface of an aircraft engine are disclosed. An example system to control a surface in an aircraft engine comprises a first valve to vary a flow of cold fluid from a thermal transfer bus (TTB) to an active surface control (ASC) system based on an operating condition of the aircraft engine, the ASC system positioned adjacent to the surface, the first valve positioned upstream from the surface, and a second valve to vary a flow of hot fluid from the TTB to the ASC system based on the operating condition, the second valve positioned downstream from the surface.
A reciprocating motion insertion tool include a flexible section, an end effector actuator, a connector within the flexible section and coupled to the end effector actuator, and an end effector coupled to a distal end of the flexible section, wherein the end effector is configured to move in a reciprocating motion at a select reciprocation rate when driven by the end effector actuator.
A gas turbine engine includes a compressor section, combustion section, and turbine section is serial flow arrangement. A fuel injector supplies a mixture of fuel and air for combustion within the combustion section. A first annular structure defines a central passage and a longitudinal axis within the fuel injector. A second annular structure is spaced from and in annular arrangement about the first annular structure to define an outer passage in annular arrangement between the first annular structure and the second annular structure.
F23D 14/36 - Brûleurs spécialement conçus pour être utilisés avec des moyens comprimant le combustible gazeux ou l'air de combustion dans lesquels le compresseur et le brûleur forment une unité
F23D 14/24 - Brûleurs à gaz sans prémélangeur, c.-à-d. dans lesquels le combustible gazeux est mélangé à l'air de combustion à l'arrivée dans la zone de combustion avec des conduits d'alimentation en air et en gaz séparés, p. ex. avec des conduits disposés parallèlement ou se croisant au moins un des fluides étant soumis à un mouvement tourbillonnant
74.
LUBRICATION MAINTENANCE SYSTEMS AND METHODS OF CHANGING LUBRICATION IN A TURBINE ENGINE
A lubrication maintenance system for a gearbox assembly of a turbine engine includes a reservoir that stores a lubricant and a lubrication pump fluidly coupled to the reservoir that circulates the lubricant through the lubrication maintenance system. A heat exchanger is fluidly coupled to the lubrication pump and the gearbox assembly of the turbine engine, and a plurality of sumps are fluidly coupled to the heat exchanger. The lubrication pump is fluidly coupled to the gearbox assembly and each of the plurality of sumps, such that the lubrication pump scavenges circulated lubricant from the gearbox assembly and each of the plurality of sumps and recycles the circulated lubricant to the reservoir.
A composite airfoil comprising an airfoil portion and a composite ply. The airfoil portion has an outer wall extending between a root and a tip, and between a leading edge and a trailing edge. The composite ply has a first set of fibers and a second set of fibers. Each fiber of the first set of fibers along a first centerline axis. Each fiber of the second set of fibers extends along a second centerline axis.
A method of generating an engine health assessment based on one or more index values. The method includes receiving, by a first computer system, a first set of data including environmental data, and a second set of data including operating data having an operating time and location data from an aircraft having an engine, calculating localized environmental data from the first and second sets of data, generating one or more index values based on the localized environmental data, receiving, by a second computer system, the one or more index values, and a third set of data including operational data, and performing engine health assessment using an analytical model, based on the one or more index values, and the third set of data, to estimate a health of the engine, determining a future health or condition of the engine, and outputting recommended actions to perform maintenance to the engine.
77.
SYSTEMS AND METHODS FOR LIMITING VOID FORMATION IN CERAMIC MATRIX COMPOSITE COMPONENTS
A method for limiting void formation in a melt-infiltrated ceramic matrix composite (MI-CMC) component includes arranging one or more infiltrant feedstocks in fluid communication with a targeted area of the MI-CMC component. The one or more infiltrant feedstocks have a nominal melting point at or below a nominal melting point of an alloy within the MI-CMC component. The method includes heating the one or more infiltrant feedstocks to a first temperature at or above the nominal melting point of the one or more infiltrant feedstocks to form a molten phase. The method also includes infiltrating the targeted area of the MI-CMC component with the molten phase. As such, the molten phase reacts with a solid phase in the targeted area of the MI-CMC component. Further, the method includes cooling the MI-CMC component to a second temperature that is below the first temperature to solidify the molten phase.
A method of forming a preform for a composite component. The method includes laying up a plurality of plies to form an initial preform having an initial shape and partially bonding adjacent plies of the plurality of plies to each other to form a bonded preform with the initial shape. The bonded preform includes a high adherence region and a low adherence region, the bonding between adjacent plies being greater in the high-adherence region than the low-adherence region. The method also includes forming the bonded preform from the initial shape to a final shape to generate a shaped preform. The final shape has a low-contour region formed from the high adherence region and a high-contour region formed from the low-adherence region.
In one embodiment, a method includes determining a pitch value based on at least an energy goal. The energy goal may be at least one of bowed rotor motoring or an engine start. The method includes adjusting a pitch of a plurality of variable-pitch propeller blades based on the pitch value. The plurality of variable-pitch propeller blades are part of a variable-pitch propeller. The method includes, in response to a rotation of the plurality of variable-pitch propeller blades about a central axis based on the pitch of the plurality of variable-pitch propeller blades, activating an electric brake to reduce a speed of the rotation. The method further includes transferring electrical energy, generated by the electric brake reducing the speed of the rotation, to a component.
B64C 11/30 - Mécanismes de changement de pas des pales
B64D 31/06 - Dispositifs amorçant la mise en œuvre actionnés automatiquement
B64D 35/04 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions caractérisée par le fait que la transmission entraîne plusieurs hélices ou rotors
B64D 41/00 - Installations génératrices de puissance pour servitudes auxiliaires
A blended wing aircraft is provided, including a body having a fuselage and a pair of wings extending outward from the fuselage; and an aircraft engine defining an outlet and including a thrust reverser assembly, the thrust reverser assembly including a deployable structure extending less than 360 degrees around the outlet.
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
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 29/04 - Nacelles, carénages ou capotages des groupes moteurs montés dans le fuselage
An electric propulsion system includes a fan rotor assembly having a first fan portion and a second fan portion. A fan cowling defines a first flow passage and a second flow passage. An electric drive mechanism drives the fan rotor assembly. An electric drive cooling system includes a liquid coolant that provides cooling to at least a part of the electric drive mechanism, and a heat exchanger through which the liquid coolant flows is arranged within the second flow passage. The first fan portion provides a first flow of air through the first flow passage to provide at least one of a lifting force or a thrust force to the electric propulsion system, and the second fan portion provides a second flow of air through the second flow passage to provide a flow of cooling air therethrough that is in thermal communication with the heat exchanger.
A turbine engine including a turbo-engine, a gearbox assembly, a propulsor, and a lubrication system. The turbo-engine includes a compressor section, a combustor, a turbine section, and an input shaft. The gearbox assembly includes a first gear, a plurality of second gears, and a third gear. The propulsor has an output shaft drivingly coupled to the input shaft through the gear assembly. The lubrication system is characterized by a Gearbox Lubrication System Parameter (GLSP) between 0.2 and 140 when a mass flow rate of the lubricant is linear with a lubricant pump speed, where the GLSP is given by:
The GLSP is between 0.2 and 70 when the mass flow rate of the lubricant is modulated, where the GLSP is given by:
.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 3/107 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur avec plusieurs rotors raccordés par transmission de puissance
A method of manufacturing a composite panel is provided. The method includes applying a composite face sheet to a first side of a core structure, the core structure comprising a plurality of first ceramic particles each having a first particle size that is within a first particle size range and the composite face sheet comprising a plurality of second ceramic particles each having a second particle size that is within a second particle size range, wherein the second particle size range is smaller than the first particle size range and densifying the composite panel through infiltration, wherein the infiltration comprises transport of an infiltrant through the core structure and into the composite face sheet.
An airfoil includes a non-uniform weave structure that is two-dimensional or three-dimensional, the non-uniform weave structure including a plurality of reinforcing fibers, the non-uniform weave structure having a first region with a first stiffness and a second region with a second stiffness higher than the first stiffness, wherein the plurality of reinforcing fibers include higher density fibers in the second region and lower density fibers in the first region so as to increase a stiffness of the airfoil at the second region of the airfoil in a desired orientation to achieve a desired aeromechanics response of the airfoil.
A powder removal apparatus includes an extraction housing comprising a sidewall that is sized and configured to extend around a powder bed of a build module and a top wall that is sized and configured to extend between opposite sides of the sidewall and over the powder bed. The sidewall and top wall are configured to form a chamber portion of a turbulence chamber. The top wall has a vacuum exit opening that is configured to fluidly connect to a vacuum source. The sidewall has a plurality of sidewall inlet flow channels that extend from an inlet opening at an exterior side of the sidewall to an outlet opening at an interior side of the sidewall. A side exit channel is configured to extend along the top wall from a collector opening in communication with the chamber portion toward the vacuum exit opening.
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
B29C 64/255 - Enceintes pour le matériau de construction, p. ex. récipients pour poudre
B29C 71/00 - Post-traitement d'objets sans modification de leur formeAppareils à cet effet
B33Y 30/00 - Appareils pour la fabrication additiveLeurs parties constitutives ou accessoires à cet effet
B33Y 40/20 - Posttraitement, p. ex. durcissement, revêtement ou polissage
A de-powdering system includes one or more sidewalls defining a support chamber configured to contain an additive manufacturing build where the additive manufacturing build includes one or more objects disposed within a powder build material. A fluidization mechanism is fluidically couplable to a fluid source and includes one or more flow channels fluidically coupled to the support chamber. The fluid source is actuatable to provide a fluid from the fluid source to the support chamber and inject the fluid into the support chamber via the one or more flow channels. The one or more flow channels are oriented to introduce a swirling flow of the fluid into the support chamber to fluidize at least a portion of the powder build material within the support chamber.
An additive manufacturing apparatus includes a stage configured to hold a component. A radiant energy is device operable to generate and project radiant energy in a patterned image. An actuator is configured to change a position of the stage relative to the radiant energy device. A deposition assembly is upstream of the stage and configured to deposit a resin on a resin support. The deposition assembly includes a reservoir housing configured to retain a volume of resin between the upstream wall and the downstream wall. The deposition assembly also includes an application device operably coupled with the reservoir housing. A computing system is operably coupled with the application device. The computing system is configured to intermittently initiate a flush operation between successive layers of the component, wherein the application device is moved from a first position to a second position during the flush operation.
B29C 64/124 - Procédés de fabrication additive n’utilisant que des matériaux liquides ou visqueux, p. ex. dépôt d’un cordon continu de matériau visqueux utilisant des couches de liquide à solidification sélective
A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes 18-26 fan blades. The low-pressure turbine includes four rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 2.0-5.1. Additionally (or alternatively) the low-pressure turbine includes an area-EGT ratio within a range of 1.2-1.3.
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 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
89.
BUILD MATERIAL ESCAPEMENT ASSEMBLY AND ADDITIVE MANUFACTURING SYSTEMS INCLUDING SAME
A build material escapement assembly for an additive manufacturing system includes a retaining plate defining an outer perimeter and an inner perimeter, a retractable plate, and a top plate coupled to the retractable plate, the top plate including a top plate perimeter and being actuatable between a retracted position and an extended position. A diaphragm is coupled to the top plate via the retractable plate and further includes an exposed area extending between the inner perimeter of the retaining plate and the top plate perimeter of the top plate. The retractable plate is actuated in a lateral direction as the top plate is moved from the retracted position to the extended position, and a width of the exposed area of the diaphragm is greater than a travel distance that the retractable plate is actuated in the lateral direction.
B29C 64/307 - Manipulation du matériau destiné à être utilisé en fabrication additive
B29C 64/153 - Procédés de fabrication additive n’utilisant que des matériaux solides utilisant des couches de poudre avec jonction sélective, p. ex. par frittage ou fusion laser sélectif
An additive manufacturing system for additively manufacturing a three-dimensional object includes a beam generation device and a first lens array disposed downstream from the beam generation device. The first lens array divides an energy beam received from the beam generation device into a plurality of beam segments. An optical modulator is disposed downstream from the first lens array and is modulated to reflect or transmit one or more beamlets from the plurality of beam segments incident on the optical modulator. A second lens array is disposed downstream from the optical modulator where the one or more beamlets are incident upon the second lens array. A focusing lens assembly is disposed downstream from the second lens array. The one or more beamlets projected from the second lens array become incident on the focusing lens assembly, and the focusing lens assembly converges the one or more beamlets in a target plane.
B29C 64/268 - Agencements pour irradiation par faisceaux laserAgencements pour irradiation par faisceaux d’électrons [FE]
B33Y 30/00 - Appareils pour la fabrication additiveLeurs parties constitutives ou accessoires à cet effet
G02F 1/01 - Dispositifs ou dispositions pour la commande de l'intensité, de la couleur, de la phase, de la polarisation ou de la direction de la lumière arrivant d'une source lumineuse indépendante, p. ex. commutation, ouverture de porte ou modulationOptique non linéaire pour la commande de l'intensité, de la phase, de la polarisation ou de la couleur
91.
AUTOMATED DE-POWDERING OF ADDITIVE MANUFACTURING BUILD
A de-powdering system for additively manufactured objects includes an enclosure defining a cavity configured to support an additive manufacturing build. The enclosure includes a wall defining a lower boundary of the cavity, and the wall includes one or more flow channels. A sleeve is disposable in the cavity to at least partially surround the additive manufacturing build. At least one vibration mechanism is coupled to the sleeve and is actuatable to induce vibrations to the additive manufacturing build to loosen at least a portion of a powder build material from one or more objects suspended within the powder build material. The powder build material is removed from the cavity via the one or more flow channels.
B08B 7/02 - Nettoyage par des procédés non prévus dans une seule autre sous-classe ou un seul groupe de la présente sous-classe par distorsion, battage ou vibration de la surface à nettoyer
B08B 13/00 - Accessoires ou parties constitutives, d'utilisation générale, des machines ou appareils de nettoyage
B08B 15/02 - Précautions prises pour empêcher les crasses ou les fumées de s'échapper de la zone où elles sont produitesRamassage ou enlèvement des crasses ou des fumées de cette zone par utilisation de chambres ou de hottes recouvrant cette zone
An apparatus and method for an inspection apparatus for inspecting an engine component. The inspection apparatus includes at least one controller configured to receive a set of inspection parameters based on a detection metric. A non-destructive evaluation (NDE) instrument for scanning a predetermined area of a surface of the engine component according to the set of inspection parameters to generate a data set is included. Further, the inspection apparatus includes a computer configured to apply a detection algorithm to the data set.
G01M 15/04 - Test des moteurs à combustion interne
G01N 23/223 - 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 mesurant l'émission secondaire de matériaux en irradiant l'échantillon avec des rayons X ou des rayons gamma et en mesurant la fluorescence X
A hybrid electric gas turbine engine is provided. The hybrid electric gas turbine engine includes: a turbomachine having a compressor section and a turbine section arranged in serial flow order, the compressor section and turbine section together defining a core air flowpath, the turbomachine defining a core air flowpath exhaust; and an electric machine assembly having an electric machine disposed aft of the core air flowpath exhaust and mechanically connected to the turbine section.
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 gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (FnTotal) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: FnTotal×EGT/(AHPCExit2×1000). The gas turbine engine further includes a blade effective acoustic length (BEAL), an acoustic spacing, and an acoustic spacing ratio (ASR). The ASR can be in a range from 1.5 to 16.0.
A method of mitigating rotor bow in a rotor of a turbine engine. The method includes determining thermal rotor bow in the rotor, determining non-thermal rotor unbalance in the rotor by monitoring the response at the bowed rotor mode of the rotor, and determining a time period for motoring the rotor prior to operation of turbine engine, wherein the time period is based on a combination of the thermal rotor bow and the non-thermal rotor unbalance. The method also includes motoring the rotor for the time period and until the vibration of the rotor is below a predetermined acceptable value.
F01D 19/02 - Démarrage des "machines" ou machines motricesDispositifs de régulation, de commande ou de sécurité en rapport avec les organes de démarrage dépendant de la température des éléments constitutifs, p. ex. du carter de la turbine
A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section. The turbine section includes a band having an upstream end and a downstream end. The band extends between the upstream end and the downstream end, and the band at least partially defines the working gas flow path. A plurality of airfoils extend into the working gas flow path from the band. Each airfoil of the plurality of airfoils includes a leading edge, a trailing edge, a first side, and a second side opposite the first side. Each of the plurality of airfoils is substantially symmetric across an airfoil centerline extending through a center of each of the plurality of airfoils. The band defines a valley portion adjacent the leading edge of each of the plurality of airfoils and a pair of hill portions on opposing sides of the valley portion.
A turbine engine including a fan having a plurality of fan blades, a turbo-engine positioned downstream of the fan and having a core inlet, an engine intake that extends to the core inlet, and a variable engine intake system. Air enters the turbine engine through the engine intake. The variable engine intake system includes a plurality of wind condition sensors for sensing wind conditions about the turbine engine, and a controller. The plurality of wind condition sensors includes a first wind condition sensor on a first side of the turbine engine and a second wind condition sensor on a second side of the turbine engine. The controller adjusts the engine intake based on the wind conditions about the turbine engine from the first wind condition sensor and the second wind condition sensor.
Systems, apparatus, articles of manufacture, and methods are disclosed that include an air foil bearing, the air foil bearing comprising: a thrust disc coupled to a rotor shaft, the thrust disc and rotor shaft to rotate; a thrust pad aligned with a first side of the thrust disc, the thrust pad to engage with the thrust disc as the thrust disc rotates; and a micro lattice structure between the thrust disc and the thrust pad, the micro lattice structure to mitigate the thrust pad engaging with the thrust disc.
A combustor of a turbine engine includes a first combustion zone operable to combust a first fuel and air mixture, a first fuel inlet for providing a first fuel, a first air inlet for providing first zone air, the first fuel and the first zone air combining to form the first fuel and air mixture in the first combustion zone, a second combustion zone operable depending on turbine engine operating parameters, for combusting a second fuel and air mixture, a second fuel inlet providing a second fuel, that operates when the second combustion zone is operating and does not operate when the second combustion zone is not operating, and a second air inlet providing second zone air, the second fuel and the second zone air combining to form the second fuel and air mixture in the second combustion zone, wherein the first fuel and the second fuel are disparate fuels.
A turbine engine operable in a cold start condition to prevent wear on components of a gearbox assembly of the turbine engine. The turbine engine includes a gearbox assembly, a pump for directing lubricant to the gearbox assembly, a supply line heating path comprising a heat exchanger, a recirculation bypass path, a valve being positionable between a first position to direct the flow of the lubricant into the supply line heating path and a second position to direct the flow of the lubricant into the recirculation bypass path, and an electronic control unit configured to position the valve between the first position and the second position based on a temperature of the lubricant.
