Viscous damper apparatus and associated methods to control a response to a resonant vibration frequency are disclosed. An apparatus to support an aircraft engine includes a thrust link including a forward end and an aft end, the forward end of the thrust link coupled to the aircraft engine, and a damper including a piston rod coupled to the aft end of the thrust link, the piston rod including a piston, and a chamber including a fluid, the piston to move within the chamber.
A gas turbine engine is provided having a turbomachine comprising 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 and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:
A gas turbine engine is provided having a turbomachine comprising 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 and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:
(
T
OUT
A
DTExit
)
2
*
EGT
A
HPCExit
*
1
0
-
1
1
;
A gas turbine engine is provided having a turbomachine comprising 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 and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:
(
T
OUT
A
DTExit
)
2
*
EGT
A
HPCExit
*
1
0
-
1
1
;
wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.
F02C 6/06 - 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
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
3.
GEARBOX ASSEMBLY WITH LUBRICANT EXTRACTION VOLUME RATIO
A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
G
B
·
V
G
A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
G
B
·
V
G
is a gutter volume of the gutter and VGB is a gearbox volume. The gas turbine engine includes a lubricant flow control system that includes a variable flow lubricant pump that generates a pump variable flow of lubricant to the gearbox assembly. The gearbox assembly has a variable consumption demand for delivery of lubricant. A lubricant flow controller is configured to generate a pump control command for the variable flow lubricant pump to produce the pump variable flow of lubricant based on the variable consumption demand.
F16H 57/04 - Caractéristiques relatives à la lubrification ou au refroidissement
F16N 7/38 - Installations à huile ou autre lubrifiant non spécifié, à réservoir ou autre source portés par la machine ou l'organe machine à lubrifier avec pompe séparéeInstallations centralisées de lubrification
4.
GEARBOX ASSEMBLY WITH LUBRICANT EXTRACTION VOLUME RATIO
A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
G
B
·
V
G
A gas turbine engine includes a fan, a combustor positioned in a core air flowpath that generates combustion gases, a steam system that extracts water from the combustion gases and generates steam, and a gearbox assembly. The steam system includes water storage devices that store the water therein. The water storage devices include a first state in which a level of the water increases or is maintained and a second state in which the level of the water decreases as the water flows through the water storage devices. The gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by:
V
G
V
G
B
·
V
G
is a gutter volume of the gutter and VGB is a gearbox volume.
Systems and methods for operating systems are provided. For example, a system comprises a heat source for providing a flow of a hot fluid and a fuel flowpath for a flow of a fuel. The fuel flowpath includes a fuel accumulator and a heat exchanger for heat transfer between the hot fluid and fuel. The heat exchanger includes a hot fluid inlet for receipt of the hot fluid at an inlet temperature and a fuel inlet for receipt of the fuel at an inlet temperature. The hot fluid inlet temperature is greater than the fuel inlet temperature such that the fuel is heated through heat transfer with the hot fluid in the heat exchanger. The fuel accumulator accumulates at least a portion of the heated fuel. An exemplary system is selectively operated to heat and circulate the fuel through the fuel flowpath for consumption and/or accumulation in the fuel accumulator.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
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 7/232 - Soupapes pour combustibleSystèmes ou soupapes de drainage
F02C 9/26 - Commande de l'alimentation en combustible
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
6.
METHODS AND APPARATUS FOR LEAK DETECTION AND MITIGATION FOR HYDROGEN FUELED AIRCRAFT
Systems, apparatus, articles of manufacture, and methods for leak detection and mitigation for hydrogen fueled aircraft are disclosed. An example apparatus disclosed herein includes machine readable instructions, and programmable circuitry to at least one of instantiate or execute the machine readable instructions to determine a hydrogen concentration threshold for a location within an undercowl of an engine of an aircraft, based on an engine condition of the aircraft, determine, based on an output of a hydrogen concentration sensor within the undercowl, a hydrogen concentration at the location, compare the hydrogen concentration to the hydrogen concentration threshold, and conduct a mitigation action in the hydrogen fuel distribution system based on the comparison of the hydrogen concentration and the hydrogen concentration threshold.
B64D 37/32 - Mesures de sécurité non prévues ailleurs, p. ex. contre les explosions
B64D 37/30 - Circuits de carburant pour carburants particuliers
G01M 3/16 - Examen de l'étanchéité des structures ou ouvrages vis-à-vis d'un fluide par utilisation d'un fluide ou en faisant le vide par détection de la présence du fluide à l'emplacement de la fuite en utilisant des moyens de détection électrique
7.
SYSTEM AND METHOD FOR GENERATING A TIME-SENSITIVE NETWORK SCHEDULE
Systems, methods, and other embodiments described herein relate to generating a time-sensitive network schedule for a time-sensitive network (TSN). In one embodiment, a method includes receiving a measured value of at least one environmental condition from at least one sensor. The method includes determining an impact of the measured value of the at least one environmental condition on at least one component of the TSN. The method includes programming a switching node based on at least the determined impact. The TSN includes a first end node, a second end node, and the switching node. The switching node is communicatively linked between the first end node and the second end node.
A shroud hanger assembly is provided for hangers and shrouds defining dimensionally incompatible components such as those which are press or frictionally fit to engage one another. The shroud hanger assembly includes a multi-piece hanger and a shroud that is pinned to the hanger assembly by at least one axially extending pin and which locates the shroud relative to the hanger to control motion in one or both of circumferential (tangential) and radial directions relative to the engine.
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
F01D 11/12 - 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 utilisant un élément de friction allongé, p. ex. un élément d'usure, déformable ou contraint de façon élastique
A vehicle having a body extending between a nose section and a tail section. The nose section terminating at a tip. The body having a centerline axis extending from the tip. The nose section having an outer wall terminating at the tip. The outer wall including an outer wall centerline axis, an inner surface and an outer surface.
General Electric Company Polska Sp. z o.o. (Pologne)
Inventeur(s)
Sibbach, Arthur William
Pazinski, Adam Tomasz
Abrégé
An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.
Methods are provided for repairing a defect on a silicon-containing substrate. The method may include applying a powder mixture into the defect of an existing coating on a surface of the silicon-containing substrate, wherein the powder mixture comprises silicon and germanium at a Ge mole fraction of 0.01 to 0.3; and heat treating the powder mixture within the defect at a sintering temperature that is 1150° C. to 1400° C. to form a repaired bondcoat within the defect. Repaired components are also provided that include a repaired bondcoat formed within the defect on the silicon-containing substrate, wherein the repaired bondcoat comprises a silicon-germanium phase comprising a Ge mole fraction of germanium of 0.01 to 0.3 and a Si mole fraction of silicon of 0.7 to 0.99.
Variable diameter thrust link apparatus are disclosed. A thrust link for an aircraft engine includes a forward end coupled to the aircraft engine, the forward end having a first diameter, and an aft end coupled to the aircraft engine or a pylon, the aft end having the first diameter, wherein the thrust link defines a thrust link span that extends from the forward end to the aft end, wherein a second diameter greater than the first diameter is defined between the forward end and the aft end, wherein the second diameter spans a center diameter span that is equivalent to or between 60% and 90% of the thrust link span.
General Electric Deutschland Holding GmbH (Allemagne)
Inventeur(s)
Chaudhari, Pushkar Chandrakant
Osama, Mohamed
Abrégé
A method for operating an electric machine assembly is provided. The electric machine assembly includes an electric machine having a first set of windings and a second set of windings. The method includes: operating the electric machine in a partial phase mode, wherein operating the electric machine in the partial phase mode comprises: powering a first set of windings to provide a net zero current in the first set of windings while maintaining one phase of the first set of windings in a non-conducting condition; and powering a second set of windings to provide a net zero current in the second set of windings while maintaining one phase of the second set of windings in a non-conducting condition.
H02P 29/028 - Détection d’un défaut, p. ex. court circuit, rotor bloqué, circuit ouvert ou perte de charge le moteur continuant de fonctionner malgré le défaut, p. ex. élimination, compensation ou résolution du défaut
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
H02P 6/182 - Dispositions de circuits pour détecter la position sans éléments séparés pour détecter la position utilisant la force contre-électromotrice dans les enroulements
H02P 25/22 - Enroulements multiplesEnroulements pour plus de trois phases
H02P 29/024 - Détection d’un défaut, p. ex. court circuit, rotor bloqué, circuit ouvert ou perte de charge
14.
METHODS OF ALTERING A SURFACE OF A NI-BASED ALLOY AND RESULTING COMPONENTS
Method of repairing a Ni-based alloy component are provided, along with the resulting repaired coated component. The method may include: spraying a plurality of particles onto a surface of the Ni-based alloy component to form a coating thereon. The plurality of particles comprises a mixture of Ni-based superalloy particles and Co-based superalloy particles. The particles are sprayed at a spray temperature that is less than a melting point of both the Ni-based superalloy particles and the Co-based superalloy particles. A repaired coated component may include: a Ni-based alloy component having a surface and a coating on the surface of the Ni-based alloy component. The coating comprises a plurality of deformed particles therein, with the plurality of deformed particles comprises 5% by weight to 80% by weight of a Ni-based superalloy and 20% by weight to 95% by weight of a Co-based superalloy.
Fluid-filled thrust link apparatus and associated method are disclosed. A thrust link for an aircraft engine includes a first wall having a forward portion and an aft portion at opposite ends of the thrust link, the forward portion coupled to the aircraft engine, the aft portion coupled to the aircraft engine, a pylon, or an aircraft associated with the aircraft engine, and a second wall within an interior area surrounded by the first wall, the second wall spaced apart from the first wall, a space between the first wall and the second wall defining a channel within the interior area, the channel including a fluid, the fluid pressurized based on a damping ratio to withstand a resonant vibration frequency generated by the aircraft engine.
F15B 21/00 - Caractéristiques communes des systèmes de manœuvre utilisant des fluidesSystèmes de manœuvre à pression ou parties constitutives de ces systèmes, non couverts par l'un quelconque des autres groupes de la présente sous-classe
16.
INSPECTION SYSTEMS AND METHODS FOR DAMAGE MEASUREMENT
A system may include a sensor comprising a sensor configured to capture data from a part of an engine. The system being configured to form a three-dimensional (3D) surface model of the part of the engine based on signals received from the sensor system, determine a nonplanar reference surface based on the 3D surface model of the part, and measure a characteristic of a damaged portion of the part of the engine based on the 3D surface model and the nonplanar reference surface.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
A gearbox assembly includes a gearbox having a gear assembly and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio is defined by
A gearbox assembly includes a gearbox having a gear assembly and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio is defined by
V
G
V
GB
.
A gearbox assembly includes a gearbox having a gear assembly and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio is defined by
V
G
V
GB
.
VG is a gutter volume of the gutter and VGB is a gearbox volume. A gas turbine engine includes the gearbox assembly and a lubrication system. The lubrication system includes a sump that is a primary reservoir having a first lubricant level and a secondary reservoir in the gearbox assembly. The secondary reservoir has a second lubricant level. The lubrication system fills the secondary reservoir with a lubricant between the first lubricant level and the second lubricant level. The gear assembly collects the lubricant in the secondary reservoir to supply the lubricant to the gear assembly.
A power system is provided. The power system includes: an electric machine comprising a first multiphase winding and a second multiphase winding, the first multiphase winding and the second multiphase winding being electrically opposite in phase with respect to one another; and a power converter system having: first switching elements in electric connection with the first multiphase winding and having a DC side and an AC side; second switching elements in electric connection with the second multiphase winding and having a DC side and an AC side; a plurality of capacitors coupled to the first switching elements on the DC side of the first switching elements and to the second switching elements on the DC side of the second switching elements; and a means for reducing common mode currents on the AC side of the first switching elements and on the AC side of the second switching elements.
Methods of protecting a surface of a Ni-based alloy component are provided, along with the wear strip utilized and the repaired Ni-based alloy component. The method may include: spraying a plurality of particles to form a wear strip. The plurality of particles includes a mixture of Ni-based superalloy particles and Co-based superalloy particles. The plurality of particles is sprayed at a spray temperature that is less than a melting point of the Ni-based superalloy particles and less than a melting point of the Co-based superalloy particles. The wear strip may be attached onto a surface of the Ni-based alloy component, either during the spraying of the particles (when wear strip formed on the surface of the Ni-based alloy component) or after a standalone wear strip is formed.
C22C 19/05 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de nickel avec du chrome
C22C 19/07 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de cobalt
C22F 1/10 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid du nickel ou du cobalt ou de leurs alliages
20.
System and method for neuroactivity detection in infants
A neuroactivity monitoring system includes a camera configured to acquire image data of a patient positioned on the patient support and a monitoring device in communication with the camera. The monitoring device uses the acquired image data of the camera to identify and track patient landmarks, such as facial and/or posture landmarks, and, based on the tracked movement, characterize patient neuroactivity.
G06K 9/00 - Méthodes ou dispositions pour la lecture ou la reconnaissance de caractères imprimés ou écrits ou pour la reconnaissance de formes, p.ex. d'empreintes digitales
A61B 5/00 - Mesure servant à établir un diagnostic Identification des individus
A61B 5/11 - Mesure du mouvement du corps entier ou de parties de celui-ci, p. ex. tremblement de la tête ou des mains ou mobilité d'un membre
A radiation hardened semiconductor device including a heavily doped substrate of a semiconductor device, a drift layer having a substantially uniform doping concentration and a thickness is provided. The doping concentration and the thickness of the drift layer are such that when the semiconductor device is operating at a maximum voltage rating, an electrical field profile in the drift layer extends less than 80% of the thickness of the drift layer, providing the radiation hardened nature of the device.
An electrodialysis cell includes a housing defining an internal chamber, a core positioned within the internal chamber, a first electrode positioned in the internal chamber adjacent the housing, a second electrode coupled to the core and spaced from the first electrode, and a membrane assembly positioned between the first and second electrodes in a spiral wound configuration. The housing includes an inlet end for receiving a feed fluid and an outlet end in fluid communication with the inlet end. The membrane assembly includes a plurality of ion exchange membranes spaced from each other to define a plurality of fluid channels between the inlet and outlet ends.
B01D 63/10 - Modules à membranes enroulées en spirale
C02F 1/469 - Traitement de l'eau, des eaux résiduaires ou des eaux d'égout par des procédés électrochimiques par séparation électrochimique, p. ex. par électro-osmose, électrodialyse, électrophorèse
C02F 103/08 - Eau de mer, p. ex. pour le dessalement
23.
TURBINE ENGINE INCLUDING A GAS PATH COMPONENT HAVING A HYDROPHOBIC COATING
A turbine engine for an aircraft. The turbine engine includes a core turbine engine and a steam system. The steam system extracts water from combustion gases, vaporizes the water to generate steam, and injects the steam into a core air flow path of the core engine to add mass flow to core air. A core hot gas path component may be fluidly connected to a hot gas path that routes combustion gasses from the combustor. The core turbine engine may also include a turbine fluidly connected to the hot gas path to receive the combustion gases. The turbine may include a turbine airfoil. Each of the core hot gas path component and the turbine airfoil includes a combustion-gas-facing surface facing the hot gas path. A hydrophobic coating is formed on the combustion-gas-facing surface, reducing wetting of water vapor within the combustion gases on the core hot gas path component.
GE Marmara Technology Center Muhendislik Hizmetleri Ltd (Turquie)
GENERAL ELECTRIC COMPANY (USA)
Inventeur(s)
Ataman, Volkan
Yasar, Fatih
Deniz, Emrah
Bucaro, Michael T.
Abrégé
A gas turbine engine comprising a combustion section enshrouded by a casing having at least one through passage, the combustion section comprising a dome wall and a liner at least partially defining a combustion chamber; a fuel nozzle connected to the dome wall and having a nozzle tube; and a coupling securing the fuel nozzle with casing and disposed at least partially in the at least one through passage.
GE Marmara Technology Center Muhendislik Hizmetleri Ltd (Turquie)
GE Aerospace Poland sp. z o.o. (Pologne)
Inventeur(s)
Deniz, Emrah
Yasar, Fatih
Mikolajczyk, Katarzyna Anna
Bucaro, Michael T.
Kacar, Ahmet
Wang, Anquan
Li, Hejie
Gonyou, Craig Alan
Brady, Aaron C.
Abrégé
A gas turbine engine comprises a compressor section, a combustion section, and a turbine section in a serial flow arrangement and enshrouded by a casing, the combustion section comprising: an annular combustion chamber defined by at least a dome wall and an annular liner; a plurality of combustor cups circumferentially arranged on the dome wall; a fuel supply connected to the casing; and a fuel nozzle passing through the casing and having a nozzle tube and a distribution manifold fluidly coupling the nozzle tube to at least two combustor cups of the plurality of combustor cups; wherein the fuel nozzle is fixed to the at least two combustor cups.
A method for predicting distortion of a part during sintering in an additive process includes receiving a computerized representation of a complex geometric part, discretizing the computerized representation of the complex geometric part into a plurality of elements, processing the plurality of elements of the computerized representation of the complex geometric part with a machine-learning model configured to predict a distorted geometry of the complex geometric part in response to a sintering process, wherein the machine-learning model is trained to predict distortion of a set of primitive geometric coupons represented by image data fed into the machine-learning model during training, the set of primitive geometric coupons having fewer geometries than the complex geometric part, the complex geometric part comprises a plurality of geometries corresponding to geometries associated with the set of primitive geometric coupons, and generating a computerized representation of the predicted distorted geometry of the complex geometric part.
A method for forming an object includes moving a recoat assembly over a build material, where the recoat assembly includes a first roller and a second roller that is spaced apart from the first roller, moving the second roller above the first roller in a vertical direction, rotating the first roller of the recoat assembly in a counter-rotation direction, such that a bottom of the first roller moves in a coating direction, contacting the build material with the first roller of the recoat assembly, thereby fluidizing at least a portion of the build material, while the second roller is spaced apart from the build material in the vertical direction, and moving the fluidized build material with the first roller, thereby depositing a second layer of the build material over an initial layer of build material positioned in a build area.
B29C 64/153 - Procédés de fabrication additive n’utilisant que des matériaux solides utilisant des couches de poudre avec jonction sélective, p. ex. par frittage ou fusion laser sélectif
B22F 10/14 - Formation d’un corps vert par projection de liant sur un lit de poudre
B22F 10/28 - Fusion sur lit de poudre, p. ex. fusion sélective par laser [FSL] ou fusion par faisceau d’électrons [EBM]
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
A cleaning system for an additively manufactured component includes a tank storing a cleaning fluid. A fluid circuit is operably coupled with the tank. A pump is coupled with the fluid circuit. A manifold is configured to receive fluid from the fluid circuit through the pump. At least one of a coupler defined by the manifold or a hose is coupled with the manifold. The at least one of the coupler defined by the manifold or the hose is further configured to couple with said additively manufactured component.
B08B 3/02 - Nettoyage par la force de jets ou de pulvérisations
B08B 3/10 - Nettoyage impliquant le contact avec un liquide avec traitement supplémentaire du liquide ou de l'objet en cours de nettoyage, p. ex. par la chaleur, par l'électricité ou par des vibrations
B08B 3/12 - Nettoyage impliquant le contact avec un liquide avec traitement supplémentaire du liquide ou de l'objet en cours de nettoyage, p. ex. par la chaleur, par l'électricité ou par des vibrations par des vibrations soniques ou ultrasoniques
B08B 5/02 - Nettoyage par la force de jets, p. ex. le soufflage de cavités
B08B 5/04 - Nettoyage par aspiration, avec ou sans action auxiliaire
B08B 9/032 - Nettoyage des surfaces intérieuresÉlimination des bouchons par l'action mécanique d'un fluide en mouvement, p. ex. par effet de chasse d'eau
An aircraft comprising a fuselage and an unducted turbine engine. The fuselage having a divot with an upstream edge and a downstream edge. The divot is defined by a straight reference line having a length (L) and a maximum depth (h) relative to the straight reference line. The unducted turbine engine having an engine core, a nacelle, and a set of blades. A first flow ratio (FR1) is equal to:
An aircraft comprising a fuselage and an unducted turbine engine. The fuselage having a divot with an upstream edge and a downstream edge. The divot is defined by a straight reference line having a length (L) and a maximum depth (h) relative to the straight reference line. The unducted turbine engine having an engine core, a nacelle, and a set of blades. A first flow ratio (FR1) is equal to:
h
L
.
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
B64C 1/26 - Fixation de la voilure, des empennages ou des surfaces stabilisatrices
B64D 27/40 - Aménagements pour le montage de groupes moteurs sur aéronefs
B64D 29/06 - Fixation des nacelles, carénages ou capotages
B64D 35/02 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques
30.
TURBINE ENGINE WITH A BLADE HAVING WOVEN CORE AND TOUGHENED REGION
A turbine engine comprising a compressor section, combustor section, and turbine section in serial flow arrangement, and defining an engine longitudinal axis. The turbine engine includes an airfoil with an outer wall having a pressure side and a suction side, extending between a root and a tip in a span-wise direction, and extending between a leading edge and a trailing edge in a chord-wise direction. The airfoil includes a woven core, a first bonding layer including a toughened region, and a laminate skin provided exterior of the first bonding layer and bonded to the woven core by the first bonding layer.
A turbine engine for an aircraft. The turbine engine includes a combustor and a steam system. Fuel and steam are injected into the combustor to mix with compressed air to generate a fuel and air mixture. The fuel and air mixture is combusted in the combustor to generate combustion gases. The steam system is fluidly coupled to the combustor as the steam source to provide steam to the combustor. The steam system includes a hot gas path and a steam hot gas path component. The hot gas path is fluidly coupled to the combustor to receive the combustion gases and to route the combustion gases through the steam system. The steam hot gas path component includes a wall having a combustion-gas-facing surface facing the hot gas path and a hydrophobic coating formed on the combustion-gas-facing surface.
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
F02C 3/30 - Addition d'eau, de vapeur ou d'autres fluides aux composants combustibles ou au fluide de travail avant l'échappement de la turbine
A turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order and together defining a working gas flow path; a thermal system operable with the turbomachine; a heat exchanger in fluid communication with the thermal system; and a plurality of elongated delivery devices in fluid communication with the compressor section of the turbomachine. The plurality of elongated delivery devices are configured to deliver a fluid from the compressor section to a surface of the heat exchanger.
B64D 15/00 - Dégivrage ou antigivre des surfaces externes des aéronefs
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
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p. ex. ensembles fonctionnels de chauffage à turbine à gaz
F02C 7/143 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel du fluide de travail avant ou entre les étages du compresseur
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
F02K 3/10 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant avec réchauffage supplémentaire du fluide de travailLeur commande par postcombustion
33.
PROBABILISTIC FATIGUE AND BLEND LIMIT ASSESSMENT AND VISUALIZATION METHODS FOR AIRFOILS
A method of analyzing a blended airfoil that includes generating a plurality of simulated blended airfoil designs each including one of a plurality of blend geometries, training surrogate models representing the plurality of simulated blended airfoil designs based on natural frequency, modal force, and Goodman scale factors, determining a likelihood of operation failure of each of the plurality of blended airfoil designs in response to one or more vibratory modes, determining which of the plurality of simulated blended airfoil designs violate at least one aeromechanical constraint and generating, a blend parameter visualization including a blend design space, where the blend design space includes one or more restricted regions indicating blended airfoil designs where at least one aeromechanical constraint is violates and one or more permitted regions indicating blended airfoil designs where no aeromechanical constraints are violated.
A turbine engine includes a high-pressure (HP) compressor and a bleed system. The HP compressor includes an HP compressor flowpath and a plurality of stages of HP compressor rotor blades and HP compressor stator vanes. The bleed system includes a plurality of bleed flowpaths including at least three bleed flowpaths in fluid communication with the HP compressor flowpath. The plurality of bleed flowpaths direct compressed air from the HP compressor flowpath. At least two of the bleed flowpaths are at successive stages of the plurality of stages.
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 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
35.
SYSTEM AND METHOD FOR DETECTING AND REDUCING TORQUE IN A TURBINE ENGINE
A gas turbine engine includes a fan section having a fan rotatable with a fan shaft, a turbomachinery section having a turbine and a turbomachine shaft rotatable with the turbine, a power gearbox mechanically coupled to the fan shaft and the turbomachine shaft such that the fan shaft is rotatable by the turbomachine shaft across the power gearbox, a grounded structure coupled to and supporting the power gearbox, and a torque monitoring system. The torque monitoring system includes a gearbox sensor. The gearbox sensor is coupled to the grounded structure and the torque monitoring system configured to determine a torque across the power gearbox using the gearbox sensor.
F02C 9/58 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel avec la commande de la transmission de puissance avec la commande d'une hélice à pas variable
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
36.
ADDITIVE MANUFACTURING USING SOLID STATE OPTICAL DEFLECTORS
An additive manufacturing apparatus comprises a laser beam source emitting a laser beam, a build platform, a powder source depositing a layer of powder onto the build platform, and a scanning assembly disposed along an optical path between the laser beam source and the build platform. The scanning assembly comprises at least one solid state optical deflector that modifies at least one of a size or an impingement location of the laser beam on the layer of powder at a scanning position of the laser beam. The at least one solid state optical deflector may be used to heat treat the layer of powder either before or after the powder is melted.
B29C 64/273 - Agencements pour irradiation par faisceaux laserAgencements pour irradiation par faisceaux d’électrons [FE] à pulsationsAgencements pour irradiation par faisceaux laserAgencements pour irradiation par faisceaux d’électrons [FE] à modulation de fréquence
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
A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of the combustor to receive the combustion gases and to condense water. A fuel heat exchanger is thermally coupled to the condenser by at least one heat transfer loop to receive heat from the condenser.
F02C 7/141 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel du fluide de travail
F01K 23/10 - Ensembles fonctionnels caractérisés par plus d'une machine motrice fournissant de l'énergie à l'extérieur de l'ensemble, ces machines motrices étant entraînées par des fluides différents les cycles de ces machines motrices étant couplés thermiquement la chaleur de combustion provenant de l'un des cycles chauffant le fluide dans un autre cycle le fluide à la sortie de l'un des cycles chauffant le fluide dans un autre cycle
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of a turbine to receive the combustion gases and to condense water. The fuel heat exchanger is thermally coupled to the condenser to receive heat from the water condensed by the condenser. The fuel heat exchanger is located in the fuel delivery assembly to receive the fuel and to transfer the heat received from the water to the fuel. The boiler is located downstream of the fuel heat exchanger. The boiler receives the water and is fluidly connected to the combustor to receive the combustion gases and to boil the water to generate steam.
F02C 6/18 - Utilisation de la chaleur perdue dans les ensembles fonctionnels de turbines à gaz à l'extérieur des ensembles eux-mêmes, p. ex. ensembles fonctionnels de chauffage à turbine à gaz
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
F01K 23/10 - Ensembles fonctionnels caractérisés par plus d'une machine motrice fournissant de l'énergie à l'extérieur de l'ensemble, ces machines motrices étant entraînées par des fluides différents les cycles de ces machines motrices étant couplés thermiquement la chaleur de combustion provenant de l'un des cycles chauffant le fluide dans un autre cycle le fluide à la sortie de l'un des cycles chauffant le fluide dans un autre cycle
39.
METHOD OF OPERATING A TURBINE ENGINE HAVING A BLEED SYSTEM
A method of operating a turbine engine. The turbine engine includes a high-pressure compressor including a high-pressure compressor flowpath and a plurality of stages, and a bleed system. The bleed system includes a plurality of bleed flowpaths including a first bleed flowpath from one stage of the plurality of stages and a second bleed flowpath from another stage of the plurality of stages. The method includes directing compressed air through the high-pressure compressor flowpath, directing a first portion of the compressed air through the first bleed flowpath, the first portion of the compressed air having a first mass flow, directing a second portion of the compressed air through the second bleed flowpath, determining an altitude of the turbine engine, and changing the first mass flow of the first portion of the compressed air through the first bleed flowpath based on the altitude of the turbine engine.
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
40.
TURBINE ENGINE HAVING A VARIABLE PITCH AIRFOIL ASSEMBLY
A turbine engine comprising a fan section, a compressor section, a combustion section, and a turbine section in serial flow arrangement and defining a stator portion and a rotor portion, which rotates about an engine centerline. The rotor portion comprising a variable pitch airfoil assembly. The variable pitch airfoil assembly having an airfoil and a balancing insert.
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
41.
MACHINE LEARNING MODEL TRAINING CORPUS APPARATUS AND METHOD
A control circuit accesses three-dimensional image information for a given three-dimensional object. The control circuit accesses a selection corresponding to a feature of the three-dimensional object, and then automatically generates a plurality of synthetic images of the three-dimensional object as a function of the three-dimensional and the selection of the aforementioned feature. By one approach, these synthetic images include supplemental visual emphasis corresponding to the aforementioned feature. The generated plurality of synthetic images can then be used as a training corpus when training a machine learning model.
G06V 10/774 - Génération d'ensembles de motifs de formationTraitement des caractéristiques d’images ou de vidéos dans les espaces de caractéristiquesDispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant l’intégration et la réduction de données, p. ex. analyse en composantes principales [PCA] ou analyse en composantes indépendantes [ ICA] ou cartes auto-organisatrices [SOM]Séparation aveugle de source méthodes de Bootstrap, p. ex. "bagging” ou “boosting”
On-cable pressure barrier device (6) fitted on a cable comprising an insulating sheath (3 d) around a plurality of wires (3 e) individually insulated, a part of said cable within the on-cable pressure barrier device is stripped of all insulation in order to expose the conducting core (6d1,6d2) of each wire -(6c1,6c2) of the plurality of wires, the on-cable pressure barrier device (6) comprising a resin part (6f) extending over, a first part (6d1) of the exposed conducting core of each wire connected to a first part (3c1) of the cable through a first part (6d1) of said wire, a second part (6d2) of the exposed conducting core of each wire connected to a second part (3c2) of the cable through a second part (6c2) of said wire extends outside said resin part (6f), the exposed conducting core of each wire and the resin part (6f) being able to stop the diffusion of gas incoming from the first part of the cable.
09 - Appareils et instruments scientifiques et électriques
10 - Appareils et instruments médicaux
42 - Services scientifiques, technologiques et industriels, recherche et conception
Produits et services
Computer software for collecting and analyzing data in the field of healthcare; hardware, computer software and firmware for automating control of medical equipment; software for information management, data collection and data analysis in the fields of asset optimization, machine diagnostics, and optimization of healthcare and hospital processes; computer hardware and software for medical imaging and for analysis of data gathered from a medical diagnostic apparatus to enhance clinical decision making; computer hardware and software for use with medical patient monitoring equipment, for receiving, processing, transmitting and displaying data; computer software for analyzing medical diagnostic information; computer software for controlling and managing patient medical information; computer software for electrical signaling in a patient to diagnose and/or treat medical conditions; computer software for automating and analyzing the administrative, financial, billing, and clinical records of healthcare organizations, providing healthcare regulatory compliance information, for capturing, distributing, managing and viewing electronic documents, for electronic data interchange (EDI) of healthcare transaction information, and for managing workflow in the delivery of healthcare; Computer software used to capture, store, track, report and monitor radiation levels delivered by medical devices and instruments. Medical imaging apparatus for screening and diagnostic use and for use in planning intervention and surgery; medical devices, namely, bone densitometer machines; patient monitoring equipment, namely patient monitors for monitoring patient physiological data; fetal and neonatal monitoring systems comprised of instruments for monitoring and measuring fetal and human body reactions; medical equipment namely wireless patient monitoring platform and wearable sensors for gathering data from patients to be viewed remotely on monitors and mobile devices, namely, smart phones and tablets; medical ventilators; anesthesia machines for use in patient care; electrocardiographs; apparatus for monitoring and recording electrocardiographic data from patients; incubators for babies; warming device for stabilizing infant body temperature for medical purposes; phototherapeutic apparatus for medical purposes for infant care; replacement parts for medical diagnostic and medical imaging equipment. Software as a service SaaS services featuring software for data collection and data analytics, for use in the field of healthcare; software as a service SaaS services featuring software for use in asset optimization, machine diagnostics, and optimization of healthcare processes; technical support services, namely, troubleshooting of computer software problems; remote diagnosis of medical and clinical equipment for determining the need for repair; providing technical consulting services to start-ups in the field of Healthcare IT and development of software applications for use in the medical field; providing real time monitoring of networked medical and Internet of Things (IOT) devices installed in a hospital or other healthcare facilities for detecting unauthorized access or data breach.
A turbine engine having an inner cowl that circumscribes at least a portion of an engine core, where the inner cowl is radially spaced from the engine core to define an inner cowl space. An outer cowl circumscribes at least a portion of the inner cowl where the outer cowl includes a radially outer surface spaced from a radially inner surface to define an outer cowl space. A fairing extends radially between the inner cowl and the outer cowl. A first accessory gearbox is located in the inner cowl space and the fairing. A second accessory gearbox is located in the outer cowl space and operably couples to the first accessory gearbox.
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
F02C 3/06 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux
A system may include an image sensor. A system may include an actuator configured to cause a controlled movement of the image sensor relative to a target element, the controlled movement being based on an operating velocity of the target element relative to an initial position of the image sensor. A system may include a controller communicatively coupled to the image sensor, the controller configured to: identify the operating velocity, determine an activation time to activate the image sensor for a designated exposure time based on the operating velocity and the controlled movement; and activate the image sensor at the activation time.
H04N 23/695 - Commande de la direction de la caméra pour modifier le champ de vision, p. ex. par un panoramique, une inclinaison ou en fonction du suivi des objets
A combustor having a main chamber and a trapped vortex cavity. The main chamber includes an outer liner and an inner liner. The trapped vortex cavity extends from at least one of the outer liner or the inner liner. A plurality of mixing assemblies operably injects a fuel-air mixture into the trapped vortex cavity to produce combustion gases. The trapped vortex cavity injects the combustion gases into the main chamber. A steam system is in fluid communication with the main chamber. The steam system operably injecting steam into the main chamber such that the steam flows downstream of the trapped vortex cavity.
F23R 3/16 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la configuration du flux d'air ou du flux de gaz avec des dispositifs à l'intérieur du tube à flamme ou de la chambre de combustion pour influer sur le flux d'air ou de gaz
F23L 7/00 - Alimentation du foyer en liquides ou gaz non combustibles autres que l'air, p. ex. oxygène, vapeur d'eau
50.
COMBUSTION SECTION WITH A PRIMARY COMBUSTOR AND A SET OF SECONDARY COMBUSTORS
A turbine engine with a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. A combustion section for the turbine engine, having a primary combustor liner including an inner liner and an outer liner annular about an engine centerline. A dome wall extending between the inner liner and the outer liner. A set of primary dome inlets located in the dome wall and circumferentially arranged about the engine centerline. A set of secondary combustors fluidly coupled to a primary combustion chamber, the set of secondary combustors including a first mini combustor and a second mini combustor.
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
F23R 3/42 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la disposition ou la forme des tubes à flamme ou des chambres de combustion
51.
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 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
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 rotor blade comprises an elongated body that has a pressure side and a suction side. The pressure side and suction side intersect at a leading edge and a trailing edge. The elongated body extends outward from a rotor hub. A span is described by a first straight line distance extending outward from the rotor hub along the elongated body. A chord of the elongated body is defined by a second straight-line distance extending between the leading edge and the trailing edge. At least one passageway extends through the elongated body between the pressure side and the suction side allowing a corrective flow of air to move from the pressure side to the suction side. The corrective flow of air interacts with and energizes low momentum airflow occurring along the suction side.
A trunnion-to-disk connection for use on an open fan configuration of a gas turbine engine may include an integral trunnion and blade spar inserted through a trunnion aperture of a fan disk and supported by top bearing and a bottom bearing. A cavity can be provided between a trunnion of the integral trunnion and blade spar and the fan disk, as well as between the top bearing and bottom bearing. Pressurized hydraulic fluid can be supplied to the cavity to urge the integral trunnion and blade spar in a direction to preload the bearings. Prior to pressurization, and prior to installation of the bottom bearing, the trunnion can be inserted into a trunnion aperture of the fan disk such that an end of the trunnion extends past the fan disk to provide sufficient space to insert the bottom bearing from within the open interior of the fan disk.
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 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 3/077 - 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 rotor system for a turbine engine. The rotor system includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades that rotate. The stator assembly includes a plurality of stator vanes arranged circumferentially about the stator assembly and includes at least one pair of non-uniform gaps between adjacent stator vanes. The plurality of stator vanes includes a first group of stator vanes having a first non-uniform gap between adjacent stator vanes, a second group of stator vanes having a second non-uniform gap between adjacent stator vanes, and a third group of stator vanes having a uniform spacing between adjacent stator vanes. The first non-uniform gap is positioned 180° from the second non-uniform gap. The plurality of rotor blades directs air through the plurality of stator vanes.
A method of operating a gas turbine engine is provided. The method includes receiving sensor data from one or more sensors. The method further includes receiving additional data associated with an engine event. The method further includes generating a current clearance based on at least one of the sensor data and the additional data associated with the engine event. The method further includes generating a target clearance based on at least one of the sensor data and the additional data associated with the engine event and comparing the target clearance to the current clearance. The method further includes causing the clearance adjustment system to adjust the clearance based on the comparison between the target clearance and the actual clearance by actuating the piezoelectric actuator.
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
58.
SYSTEM AND METHOD OF REPAIRING A MULTI-LAYER COMPONENT OF AN ENGINE
A method of repairing a multi-layer component of an engine in-module includes identifying a damage location on the multi-layer component. The damage location extends at least partially into a ceramic-based top layer of the multi-layer component. The method further includes depositing a material composition onto the damage location at a first temperature range so as to cover the damage location. The material composition includes one or more sintering additives. Further, the method includes applying localized curing to the material composition deposited at the damage location at a second temperature range, the second temperature range being higher than the first temperature range.
Aero-acoustically damped bleed valves are disclosed. An example variable bleed valve apparatus comprises a variable bleed valve door to actuate the variable bleed valve apparatus, and a variable bleed valve port including an upstream edge and a downstream edge, the VBV port to define a secondary flowpath, the VBV door to cover the VBV port in a closed position, and a vortex device at the upstream edge of the variable bleed valve port, the vortex device including a vorticity generating feature along the upstream edge of the variable bleed valve port.
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
F02C 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
An apparatus includes a first acoustic sensing resonator formed from a silicon substrate and has a first microelectromechanical system. The apparatus also includes a second acoustic sensing resonator formed from the silicon substrate and has a second microelectromechanical system. The second acoustic sensing resonator is arranged on the silicon substrate at a ninety degree (90°) angle with respect to the first acoustic sensing resonator and together the first acoustic sensing resonator and second acoustic sensing resonator form a torque sensor. A high temperature bonding surface is connected to the torque sensor for directly connecting the torque sensor to a metal object.
G01L 5/173 - Appareils ou procédés pour la mesure des forces, du travail, de la puissance mécanique ou du couple, spécialement adaptés à des fins spécifiques pour la mesure de plusieurs composantes de la force en utilisant des moyens acoustiques
An aircraft gas turbine includes a combustor that combusts compressed air and at least one fuel flow from at least one fuel source to generate combustion gases. A fuel supply system is arranged to provide at least one of a first flow of fuel to the combustor via a first fuel supply line and a second flow of a heated fuel to the combustor via a second fuel supply line. A fuel heat exchanger is arranged within the fuel supply system to generate the heated fuel that is heated above an autoignition temperature of the fuel, and a heat source communicates with the fuel heat exchanger for generating the heated fuel.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p. ex. à la température, à la pression, à la vitesse du rotor
62.
COMPOSITE COMPONENT HAVING AN ADDITIVELY PRINTED INNER PORTION
A composite component for a gas turbine engine is provided, along with its methods of formation. The composite component includes: an additively printed inner portion defining at least one flowpath feature, and a ceramic matrix composite (CMC) outer portion formed on the additively printed inner portion such that the CMC outer portion substantially surrounds the additively printed inner portion. The additively printed inner portion includes SiC; and the CMC outer portion includes a fiber reinforced ceramic matrix (e.g., including SiC) and defines at least one cooling cavity fluidly coupled to the at least one flowpath feature of the additively printed inner portion.
A combustor includes a combustion chamber and an annular dome. The combustion chamber includes an outer liner and an inner liner and has a combustion zone. The annular dome is coupled to the outer liner and the inner liner. A plurality of mixing assemblies operably injects a fuel-air mixture into the combustion zone of the combustion chamber to produce combustion gases. A combustor steam system is in fluid communication with the combustion chamber. The combustor steam system includes a steam path defined by at least one of the outer liner, the inner liner, or the annular dome. The combustor steam system operably directs steam through the steam path from the at least one of the outer liner, the inner liner, or the annular dome and into the combustion chamber.
An additive manufacturing machine includes an energy beam system configured to emit an energy beam utilized in an additive manufacturing process, and one or more optical elements utilized by, or defining a portion of, the energy beam system and/or an imaging system of the additive manufacturing machine. The imaging system monitors one or more operating parameters of the additive manufacturing process. A light source is configured to emit an assessment beam that follows an optical path incident upon the one or more optical elements. One or more light sensors detect a reflected beam comprising at least a portion of the assessment beam at a perimeter edge of the one or more optical elements. A control system determines, based at least in part on assessment data comprising data from the one or more light sensors, whether at least one of the one or more optical elements exhibits an optical anomaly.
An additive manufacturing machine includes an energy beam system configured to emit an energy beam utilized in an additive manufacturing process, and one or more optical elements utilized by, or defining a portion of, the energy beam system and/or an imaging system of the additive manufacturing machine. The imaging system monitors one or more operating parameters of the additive manufacturing process. A light source is configured to emit an assessment beam that follows an optical path incident upon the one or more optical elements. One or more light sensors detect a reflected beam comprising at least a portion of the assessment beam at a perimeter edge of the one or more optical elements. A control system determines, based at least in part on assessment data comprising data from the one or more light sensors, whether at least one of the one or more optical elements exhibits an optical anomaly.
A light source is configured to emit an assessment beam that follows an optical path incident upon the first and second optical elements. One or more light sensors detect a reflected beam that is either internally reflected by the first optical element or reflectively propagated between the first and second optical elements. A control system determines, based at least in part on assessment data comprising data from the one or more light sensors, whether at least one of the first and second optical elements exhibits an optical anomaly.
A seal assembly for rotary machines having a wear protection assembly with an abradable covering. The wear protection assembly has one or more wear protection features that enable in-operation opening of the seal assembly by increasing a clearance between a seal rotor and a seal slider of the seal assembly. In particular, the wear protection assembly is configured to open the seal assembly and increase the clearance of the seal rotor and the seal slider during slider air-bearing wear conditions, slider counterbore wear conditions, rotor air-bearing wear conditions, or extreme vibration wear conditions. Under such or similar wear conditions (or under conditions preceding or leading up to a wear condition), the wear protection assembly yields a pressure gradient and/or a fluid flow that oppose the forces acting to decrease the clearance between the seal rotor and the seal slider.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages
F16J 15/3284 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre par joints élastiques, p. ex. joints toriques caractérisés par leur structureEmploi des matériaux
A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.
B64D 33/02 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des entrées d'air de combustion
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
67.
COMBUSTOR INCLUDING A STEAM INJECTOR OPERABLY INJECTING STEAM INOT A TRAPPED VORTEX CAVITY
A combustor includes a combustion chamber having an outer liner and an inner liner and defining a first combustion zone, an annular dome coupled to the outer liner and the inner liner, and a trapped vortex cavity extending from at least one of the outer liner or the inner liner and defining a second combustion zone. A plurality of first mixing assemblies are disposed through the annular dome, and operably inject a first fuel-air mixture into the first combustion zone. A plurality of second mixing assemblies are disposed at the trapped vortex cavity, and operably inject a second fuel-air mixture into the second combustion zone defined in the trapped vortex cavity to produce combustion gases. A steam system includes a steam injector in fluid communication with the trapped vortex cavity. The steam injector operably injects steam into the trapped vortex cavity and the steam mixes with the combustion gases.
A turbo engine for an aircraft includes a gas turbine engine having a fuel system to provide fuel to a combustion section. The fuel system includes a variable displacement pump driven by an accessory gearbox that is powered by a shaft of the gas turbine engine. The variable displacement pump has an inlet side and an outlet side. The variable displacement pump includes an actuator configured to adjust a discharge flow rate of the variable displacement pump. The fuel system also includes a fuel metering valve configured to receive pressurized fuel from the variable displacement pump and control a flow rate of the pressurized fuel to the combustion section. The fuel system further includes a bypass valve fluidly coupling the outlet side and the inlet side of the variable displacement pump to direct an excess portion of the pressurized fuel at the outlet side back to the inlet side.
A microneedle actuator for monitoring ISF of a user includes a rigid support defining a slot, a geared cam positioned within the rigid support and including an axle extending through the slot, and a spring positioned within the rigid support and configured to rotate the geared cam about the axle. The microneedle actuator further includes an ISF sensing module and an actuation device. The ISF sensing module is attached to the geared cam and including a plurality of microneedles, each microneedle of the plurality of microneedles including at least one sensor for detecting the ISF of the user. The actuation device is attached to the rigid support and the geared cam. Actuation of the actuation device releases the spring, rotating the geared cam to insert the plurality of microneedles into the user's skin, and wherein the at least one sensor detects constituents of the ISF of the user.
A wearable interstitial fluid (ISF) sensing includes a first patch layer and an electronics module component layer. The electronics module component layer is laminated to the first patch layer to form a patch component and a plurality of microneedles extending from the patch component. The electronics module component layer includes a substrate and a sensor positioned on the substrate. Each microneedle of the plurality of microneedles includes a microneedle base extending from the patch component, a body extending from the microneedle base, and a tip extending from the body. The body defines a slot extending through the first patch layer of the body. The slot defines a first opening and a second opening. The at least one sensor is positioned over the second opening. The first opening, the slot, and the sensor define a channel exposed to the environment. The channel is configured to channel ISF to the sensor.
A heat shield for a combustor of a gas turbine engine. The heat shield includes an annular ring having an axial direction, a radial direction, and a circumferential direction. The annular ring includes a plurality of circumferential segments. Each circumferential segment of the plurality of circumferential segments is disconnected from an adjacent circumferential segment of the plurality of circumferential segments to allow for thermal growth of each circumferential segment during operation of the combustor.
A blade assembly for a gas turbine engine having an engine casing, with the blade assembly being configured to rotate about a rotational axis. The blade assembly having a blade, and at least one fin. The blade extending between a root and a tip, with the tip being spaced radially from the engine casing to define a space therebetween. The at least one fin extending radially with respect to the tip and into the space.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages par obturation non contact, p. ex. du type labyrinthe
A method of forming a ceramic matrix composite (CMC) fastener, the method comprising: shaping a preform to form a green fastener; applying one or more layers to an outer surface of the green fastener to form a layered green fastener; and finishing the layered green fastener to form the CMC fastener.
A combustor having a casing, an inner liner, an outer liner, a compressed air passageway, a combustion chamber and a set of dilution passages. The compressed air passageway being formed between the casing, the inner liner and the outer liner. The combustion chamber being at least partially defined by the inner liner and the outer liner. The set of dilution passages extending through at least one of the inner liner or the outer liner. Each dilution passage of the set of dilution passages extending between an inlet fluidly coupled to the compressed air passageway and a dilution hole fluidly coupled to the combustion chamber.
An airfoil assembly for a turbine engine, the airfoil assembly including a platform defining an inner surface and an outer surface, a variable pitch airfoil extending radially from the outer surface of the platform from a root to a tip to define a span length and a mounting structure connected to the platform.
A system is provided for performing an operation on a component of an engine. The component includes a first side positioned within an interior of the engine. The system includes a first robotic arm defining a first distal end and including a first utility member positioned at the first distal end, the first robotic arm moveable to the interior of the engine to a location operably adjacent to the first side of the component; and a second robotic arm defining a second distal end and including a second utility member positioned at the second distal end, the second robotic arm also moveable to the interior of the engine to facilitate the first and second utility members performing the operation on the component of the engine.
B23K 37/00 - Dispositifs ou procédés auxiliaires non spécialement adaptés à un procédé couvert par un seul des autres groupes principaux de la présente sous-classe
B23P 6/04 - Réparation de pièces ou de produits métalliques brisés ou fissurés, p. ex. de pièces de fonderie
B25J 19/00 - Accessoires adaptés aux manipulateurs, p. ex. pour contrôler, pour observerDispositifs de sécurité combinés avec les manipulateurs ou spécialement conçus pour être utilisés en association avec ces manipulateurs
77.
POWER OVERLAY STRUCTURE FOR A MULTI-CHIP SEMICONDUCTOR PACKAGE
A multi-chip semiconductor package includes a dielectric interconnect layer having an upper surface and a bottom surface, at least one common source pad disposed on the upper surface of the interconnect layer, at least one common gate pad disposed on the upper surface of the interconnect layer, and a plurality of semiconductor devices each including a gate pad and at least one source pad adhered onto the interconnect layer, wherein the source pads of the plurality of semiconductor devices are electrically connected to the at least one common source pad, and wherein the source pads of the plurality of semiconductor devices are electrically connected in parallel with one another, and wherein the gate pads of the plurality of semiconductor devices are electrically connected to the common gate pad, and wherein the gate pads of the plurality of semiconductor devices are electrically connected in parallel with one another.
H01L 25/065 - Ensembles consistant en une pluralité de dispositifs à semi-conducteurs ou d'autres dispositifs à l'état solide les dispositifs étant tous d'un type prévu dans une seule des sous-classes , , , , ou , p. ex. ensembles de diodes redresseuses les dispositifs n'ayant pas de conteneurs séparés les dispositifs étant d'un type prévu dans le groupe
H01L 21/48 - Fabrication ou traitement de parties, p. ex. de conteneurs, avant l'assemblage des dispositifs, en utilisant des procédés non couverts par l'un uniquement des groupes ou
H01L 23/00 - Détails de dispositifs à semi-conducteurs ou d'autres dispositifs à l'état solide
H01L 23/48 - Dispositions pour conduire le courant électrique vers le ou hors du corps à l'état solide pendant son fonctionnement, p. ex. fils de connexion ou bornes
78.
SYSTEM AND METHOD FOR CONTOURING EDGES OF AIRFOILS
A tool for contouring an airfoil comprises a main body that includes a locator portion and an upper portion. The locator portion includes an engagement surface and the engagement surface is sized and configured to engage an airfoil. The upper portion includes an edge slot sized and configured for placement of an edge of the airfoil into the slot. An abrasive material is applied to the sides of the edge slot. The upper portion and the locator portion are together sized and shaped so that when the edge of the airfoil is disposed in the edge slot and the engagement surface of the locator portion is simultaneously pressed against the airfoil, engagement of the abrasive material with the edge of the airfoil is effective to contour the edge of the airfoil to a preselected and desired shape.
A rotor assembly includes a rotor core having a rotatable shaft and defining at least one rotor post, and a winding wound around the post that defines a set of rotor winding end turns. A support assembly for the rotor winding end turns is rotatably coupled to the rotatable shaft and defines a cavity in fluid communication with a fluid coolant flow. The rotor winding end turns extend into the cavity.
H02K 7/18 - Association structurelle de génératrices électriques à des moteurs mécaniques d'entraînement, p. ex. à des turbines
H02K 9/19 - Dispositions de refroidissement ou de ventilation pour machines avec enveloppe fermée et circuit fermé de refroidissement utilisant un agent de refroidissement liquide, p. ex. de l'huile
80.
SYSTEM AND METHOD FOR FLUID CAPTURE USING A CROSS-LINKED BINDER
In some embodiments, the present disclosure relates to a system. The system includes a substrate and a fluid capture material formed on one or more surfaces of the substrate. The fluid capture material includes a sorbent material that binds one or more fluids, the one or more fluids comprising water, carbon dioxide, sulfur oxides, or a combination thereof. The fluid capture material also includes one or more binder materials, wherein the binder material is at least partially cross-linked.
81.
SYSTEM AND METHOD FOR REDUCING A CLEARANCE GAP IN AN ENGINE
A method for reducing a clearance gap between a plurality of rotor blades and a shroud assembly of an engine includes determining, with a flight control system, that an airplane is in a first flight condition. The method also includes adjusting the clearance gap to a first clearance gap distance associated with the first flight condition. Further, the method includes receiving, with the flight control system, a demand for a second flight condition. During the second flight condition, the method includes adjusting at least two independently controllable parameters, the at least two independently controllable parameters comprising, at least, a first parameter for optimizing the clearance gap and a second parameter for satisfying a thrust demand of the engine, the first parameter having a first impact on the clearance gap, the second parameter having a second impact on the clearance gap, the first impact being greater than the second impact.
Methods for manufacturing a ceramic matrix composite (CMC) component are provided. The method may include building a plurality of successive layers of the ceramic component through a selective solidification of a composite feedstock comprising a plurality of composite granules and densifying the ceramic component through infiltration with silicon. Each of the plurality of composite granules comprises one or more silicon carbide particles and one or more carbon particles. Composite feedstocks are also provided, which may include a plurality of composite granules with each of the plurality of composite granules comprises one or more silicon carbide particles and one or more carbon particles.
A turbine engine for an aircraft. The turbine engine includes a combustor, a core shaft, a turbine, and a steam system providing steam to a core air flow path. The steam system includes a boiler that receives combustion gases to boil water to generate steam. A steam turbine is fluidly coupled to the boiler to receive the steam from the boiler and to cause the steam turbine to rotate. The steam turbine is coupled to the core shaft to rotate the core shaft when the steam turbine rotates. The steam system may include a bypass flow path selectively operable to redirect at least one of the steam or water, and to bypass the core air flow path. The turbine engine may also include at least one steam control valve located downstream of the boiler and upstream of the core air flow path to control the flow of steam.
F02C 6/00 - Ensembles fonctionnels multiples de turbines à gazCombinaisons d'ensembles fonctionnels de turbines à gaz avec d'autres appareilsAdaptations d'ensembles fonctionnels de turbines à gaz à des applications particulières
F01K 11/02 - Ensembles fonctionnels de machines à vapeur caractérisés par des machines motrices faisant corps avec les chaudières ou les condenseurs les machines motrices étant des turbines
F01K 15/02 - Adaptations des ensembles fonctionnels de machines à vapeur à des usages particuliers pour véhicules de traction, p. ex. locomotives
84.
GAS TURBINE ENGINE INCLUDING A COMPRESSOR BLEED AIR SYSTEM
A gas turbine includes a compressor bleed air system that includes a compressor shroud, a compressor bleed air plenum, and a plurality of compressor bleed air passages arranged through the compressor shroud that provide fluid communication between the compressor airflow passage and the compressor bleed air plenum. The plurality of compressor bleed air passages are arranged circumferentially spaced apart from one another, where a first compressor bleed air passage and a second compressor bleed air passage circumferentially adjacent to one another are spaced apart a first angular spacing, and the respective inlets of at least a portion of a remainder of the plurality of compressor bleed air passages are circumferentially spaced apart from one another a second angular spacing, the second angular spacing being less than the first angular spacing.
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
F04D 27/00 - Commande, p. ex. régulation, des pompes, des installations ou des systèmes de pompage spécialement adaptés aux fluides compressibles
F04D 29/52 - Carters d'enveloppeTubulures pour le fluide énergétique pour pompes axiales
85.
TURBINE ENGINE AND A COMPOSITE AIRFOIL ASSEMBLY THEREOF
A turbine engine including a fan section, a compressor section, a combustion section, and a turbine section in serial flow arrangement. The turbine engine further including a composite airfoil assembly. The composite airfoil assembly having an outer wall, a composite body at least partially defining the outer wall, and a tip cap at least partially defining the outer wall.
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
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
86.
Guide vane assembly with fixed and variable pitch inlet guide vanes
A guide vane assembly for a nacelle of a gas turbine engine includes a forward vane and an aft vane. The aft vane is located aft of the forward vane and forward of a plurality of fan blades. The forward vane defines a fixed pitch angle and the aft vane is movable between a first pitch angle and a second pitch angle.
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
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
87.
Combustion section with a primary combustion chamber and a secondary combustion chamber
A combustion section of a turbine engine having a primary combustion chamber and a secondary combustion chamber. The primary combustion chamber is defined, at least in part, by a combustor liner and a dome wall. A primary fuel nozzle having an outlet is located at the dome wall, wherein the outlet of the primary fuel nozzle is fluidly coupled with the primary combustion chamber. At least one opening located in an outer liner axially aft of the dome wall fluidly couples the secondary combustion chamber and the primary combustion chamber. A set of secondary fuel nozzles are fluidly coupled with the secondary combustion chamber.
A turbofan engine and an outlet guide vane structure for the turbofan engine is provided. The outlet guide vane structure includes a non-uniform plurality of outlet guide vanes utilized in a bypass duct and between a radially inward wall and a radially outward wall. The non-uniform plurality of outlet guide vanes have varying chord lengths at the base of the outlet guide vanes along the radially inward wall. The non-uniform plurality of outlet guide vanes may include structural outlet guide vanes, non-structural outlet guide vanes, and cut-out outlet guide vanes. The non-uniform plurality of outlet guide vanes include at least one outlet guide vane integrated with a pylon and/or a bifurcation of the turbofan engine. The cut-out outlet guide vane may be integrated with the pylon and/or bifurcation. The cut-out may include a leading edge and a trailing edge disposed in front of a leading edge of the pylon.
A turbine engine for an aircraft. The turbine engine includes a combustor, a turbine, a boiler, a steam turbine, and a reheat boiler. The combustor generates combustion gases and the turbine is positioned downstream of the combustor to receive the combustion gases and to rotate the turbine. The boiler is positioned downstream of the combustor to receive the combustion gases and to boil water to generate steam. The steam turbine is fluidly coupled to the boiler to receive the steam from the boiler and to rotate the steam turbine. Each of the turbine and the steam turbine is drivingly coupled to a core shaft to rotate the core shaft. The reheat boiler is fluidly coupled to the steam turbine to receive the steam from the steam turbine and to reheat the steam. The combustor is fluidly coupled to the reheat boiler to receive the reheated steam.
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
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 with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.
General Electric Company Polska Sp. z o.o. (Pologne)
Inventeur(s)
Sibbach, Arthur William
Pazinski, Adam Tomasz
Abrégé
An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.
A turbine engine for an aircraft includes a fan that rotates to generate a volume of air, a core turbine engine, a nacelle, and a steam system. The core turbine engine includes a combustor that generates combustion gases, and a turbine including a shaft. The combustor and the turbine define a core air flowpath. The fan is drivingly coupled to the shaft. The nacelle circumferentially surrounds the fan and defines a bypass airflow passage between the nacelle and the core turbine engine. The volume of air flows into the bypass airflow passage as bypass air and flows into the core air flowpath as core air. The steam system extracts water from the combustion gases, vaporizes the water to generate steam, and injects the steam into the core air flowpath to add mass flow to the core air. A bypass ratio of the turbine engine is greater than 18:1.
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
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
93.
FAN ASSEMBLY FOR AN ENGINE HAVING REDUNDANT TRUNNION RETENTION
A fan assembly for an engine has a fan blade, a disk having a disk segment, a trunnion mechanism having a trunnion extending at least partially through the disk segment, and first and second attachment systems. The trunnion has a platform and a shaft. The fan blade is adjacent to the platform of the trunnion. The trunnion is attached to the disk segment using the first attachment system along a preload path. The first and second attachment systems further include first and second retention features, respectively, separate from the preload path. During normal operation of the fan assembly, the first attachment system retains the trunnion within the disk along a first load path. Upon a failure in the trunnion, the first and second retention features together retain the trunnion within the disk segment along a second load path, the first load path being different than the second load path.
Systems and methods for engine imaging are provided. An imaging system includes an optical sensor coupled to an insertion tool and a light source configured to output alight projection onto a component of the engine assembly. The system includes a processor configured to: receive data comprising a plurality of frames from the optical sensor captured while the component of the engine assembly is rotating, determine an angular displacement of the component between frames, and form a 3D point cloud of the component based on combining the data in the frames based on the angular displacement of the component between the frames.
A gas turbine engine includes a fan section, a compressor section, combustion section, and turbine section in serial flow arrangement, that defining an engine centerline extending between a forward direction and an aft direction. A composite component, such as an airfoil, extends between a root and a tip, defining a span-wise direction therebetween, and between a leading edge and a trailing edge, defining a chord-wise direction therebetween. The composite component includes a preform core including a first set of fibers as a first set of warp fibers and a first set of weft fibers.
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
F02C 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
A rotor blade system for a turbine engine having a longitudinal centerline axis. The rotor blade system includes a rotor configured to rotate about the longitudinal centerline axis and a plurality of rotor blades coupled circumferentially around the rotor, each rotor blade having a leading edge, a trailing edge, a suction side, and a pressure side. At least one rotor blade of the plurality of rotor blades has an axial sweep of both the leading edge and the trailing edge as compared to a base blade of the plurality of rotor blades.
General Electric Company Polska Sp. z o.o. (Pologne)
Inventeur(s)
Sibbach, Arthur William
Pazinski, Adam Tomasz
Abrégé
An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.
A power converter apparatus comprises a set of switching elements communicatively coupled with a set of gate drive circuits. Each gate drive circuit is configured to provide a respective drive signal to a corresponding switching element, each switching element being switchably responsive to the respective drive signal. The apparatus includes a controller module configured to control an output state of the power converter, and selectively change one of a respective gate resistance and a respective gate current of a corresponding subset of the gate drive circuits based on the output state of the power converter.
H02M 1/08 - Circuits spécialement adaptés à la production d'une tension de commande pour les dispositifs à semi-conducteurs incorporés dans des convertisseurs statiques
H02M 7/5395 - 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
99.
METHODS AND APPARATUS FOR RECOATING PARAMETER CONTROL
Methods and apparatus for recoating parameter control are disclosed. An example apparatus disclosed herein includes a blade holder, a blade, and a control element disposed within the blade holder, the control element to move the blade between a first position and a second position, the apparatus having a first stiffness when the blade is in the first position, the apparatus having a second stiffness when the blade is in the second position, the first stiffness greater than the second stiffness.
B22F 10/28 - Fusion sur lit de poudre, p. ex. fusion sélective par laser [FSL] ou fusion par faisceau d’électrons [EBM]
B22F 10/85 - Acquisition ou traitement des données pour la commande ou la régulation de procédés de fabrication additive
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 12/41 - Moyens de rayonnement caractérisés par le type, p. ex. laser ou faisceau d’électrons
B33Y 30/00 - Appareils pour la fabrication additiveLeurs parties constitutives ou accessoires à cet effet
B33Y 50/02 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
The present disclosure is generally related to aircraft having one or more unducted fan propulsors at locations within specific regions relative to an airfoil, such as a wing or horizontal stabilizer. More specifically, the specific regions are located where there is a relatively higher pressure air flow beneath the wings or above a horizontal stabilizer. That higher pressure air flow can be utilized to provide increased thrust from the unducted fan propulsor.
