An aircraft engine has a gearbox having a casing defining an internal chamber and a scavenge port. A gear is rotatably mounted inside the internal chamber for rotation about a rotation axis. The gear is axially spaced from the scavenge port. A baffle is disposed inside the internal chamber. The baffle includes a perforated plate disposed radially between the gear and the scavenge port relative to the rotation axis. The perforated plate has a first portion extending around a portion of a circumference of the gear, and a second portion extending axially away from the first portion over the scavenge port of the scavenge area.
A battery system is provided that includes a plurality of battery strings, negative path electrical conductors, positive path electrical conductors, and string balancing resistors. The battery strings are disposed in a parallel configuration. Each battery string includes positive and negative terminals. The negative terminal of each battery string is selectively in electrical communication with a load through a respective negative path electrical conductor, and the positive terminal of each battery string is selectively in electrical communication with the load through a respective positive path electrical conductor. The battery system is configurable in a balancing state wherein the battery system is disconnected from the load and the battery system is configured such that the string balancing resistors cause the battery strings to balance in voltage relative to one another.
B64D 27/355 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles à combustible
H01M 10/42 - Procédés ou dispositions pour assurer le fonctionnement ou l'entretien des éléments secondaires ou des demi-éléments secondaires
H01M 50/509 - Interconnecteurs pour connecter les bornes des batteries adjacentesInterconnecteurs pour connecter les cellules en dehors d'un boîtier de batterie caractérisées par le type de connexion, p. ex. connexions mixtes
A drain shield can include a body, a stop, and a plastically deformable portion. The body includes a passage extending through the body and along an axis. The stop extends outward from a first end of the body relative to the axis. The plastically deformable portion is at a second end opposite the first end. A casing module can include a wall separating a high-pressure region from a low-pressure region fluidly connected by an aperture extending through the wall. The drain shield is insertable through the aperture and deformable to enclose the aperture.
A fitting assembly for a high-pressure gaseous fluid includes a female connector disposed about an axis and comprising a first frustoconical sealing surface, a male connector configured to be received in the female connector about the axis, the male connector comprising a second frustoconical sealing surface complimentary with and configured to engage the first frustoconical sealing surface to form a first seal at a first location, an annular sealing element disposed between the female connector and the male connector to form a second seal at a second location, and a threaded coupling nut configured to engage the first and second frustoconical sealing surfaces and secure the female connector to the male connector.
A fuel injector includes a nozzle and a cap for delivering an oxidant-fuel mixture along a nozzle axis. The nozzle includes a fuel passage and a swirler. The fuel passage extends along the nozzle axis. The swirler circumscribes the fuel passage and includes an oxidant passage that converges towards the nozzle axis. The cap includes a peripheral body, an end body, and an effusion passage. The peripheral body circumscribes the swirler. The end body joins to the peripheral body and extends radially towards the nozzle axis. The effusion passage extends through the cap to intersect at least one of the peripheral body and the end body.
A method of manufacturing an annular bearing housing for a gas turbine engine is provided that includes: producing a flange outer structure segment; producing an intermediate structure segment having an outer radial end, an inner radial end, a body that extends between the outer radial end and the inner radial end, and a branch member that extends outwardly from the body; producing a main body segment; attaching the flange outer structure segment to the outer radial end of the intermediate structure segment; and attaching the main body segment to the inner radial end of the intermediate structure segment. At least one of the flange outer structure segment, the intermediate structure segment, or the main body segment is produced using an additive manufacturing process.
A support stand assembly for an aircraft engine is provided. The support stand assembly includes a mounting structure adapted to engage with and receive the aircraft engine and a support stand having a frame. Engine support members extend upwardly from opposed lateral sides of the frame. The engine support members have brackets at upper ends thereof. The brackets face toward one another and have slots defined therein. The mounting structure is displaceable between an engaged position in which the mounting structure is received in the slots to prevent relative movement between the mounting structure and the support stand, and a disengaged position in which the mounting structure is removed from the slots.
A test blade for a blade release test, including: an airfoil extending radially from a platform; and a root portion secured to the platform, the root portion having a first inner opening extending longitudinally from a leading edge of the root portion towards a trailing edge of the root portion, wherein the first inner opening extends from and is located within two second inner openings, one of which extends from the leading edge of the root portion and the other extends from the trailing edge of the root portion.
An aircraft component identification system, has: a camera configured for capturing an image of a part identifier on an aircraft component of an aircraft engine, the part identifier including a series of characters; and a controller operatively connected to the camera, the controller having a processing unit and a computer-readable medium having stored thereon instructions executable by the processing unit to: perform optical character recognition on image data obtained from the image captured by the camera to identify the series of characters, including feeding the image data to a trained model, the trained model having been trained using machine learning and training data, the training data including image data sets associated with part identifier sets; and obtain information about the aircraft component using the series of characters of the part identifier.
A fuel injector includes a mount, a stem, a nozzle, and a heat shield. The stem extends from the mount. The nozzle is at a distal end of the stem opposite the mount and configured to discharge an oxidant-fuel mixture along a nozzle axis. The heat shield extends from the mount towards the nozzle. The heat shield includes a stem segment surrounding the stem and at least a portion of the nozzle and a compliant segment extending from the stem segment to the nozzle, a lip extending from the nozzle, or a cap surrounding the nozzle.
An aircraft engine, has: a turbine; a scroll case connected to a source of combustion gases and to the turbine, and a conduit extending around a central axis; a bearing housing including a support flange; a turbine support case secured to the bearing housing, the turbine support case having spokes extending through the scroll case and radially supported by the bearing housing, a spoke of the spokes having a distal end secured to the support flange via one or more fasteners; and ribs secured to the support flange and distributed around the central axis, a rib of the ribs circumferentially and axially overlapping the distal end of the spoke, the rib having a rib face in radially opposed facing relationship to a spoke face of the spoke, a projection of the rib face in a circumferential direction relative to the central axis intersecting the spoke face.
A hybrid multi-spool gas turbine engine, has: a LP spool and a HP spool rotatable about a central axis, the LP spool having an LP compressor and an LP turbine engaged to the LP compressor via an LP shaft, the LP shaft engaged to a rotatable load at a first end thereof, the HP spool having an HP turbine and an HP compressor engaged to the HP turbine via a HP shaft; an accessory gearbox (AGB) engaged to both of the LP shaft and the HP shaft and located proximate a second end thereof, the AGB having at least one accessory output drivingly engageable to at least one accessory and at least one input; and at least one electric motor drivingly engaged to the at least one input of the AGB, the at least one electric motor drivingly engaged to the rotatable load via the AGB and the LP shaft.
F02C 3/14 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail caractérisés par l'aménagement de la chambre de combustion dans l'ensemble
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64D 35/022 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F01D 15/12 - Combinaisons avec des transmissions mécaniques
F02C 3/073 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur le compresseur ne comprenant que des étages axiaux les étages de la turbine et du compresseur étant concentriques
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
13.
DIFFERENTIAL GEARTRAIN FOR AIRCRAFT PROPULSION SYSTEM
An assembly is provided for an aircraft propulsion system. This assembly includes a geartrain, a propulsor rotor, a compressor rotor, a rotating structure and an electric machine. The geartrain includes a first component, a second component and a third component. The propulsor rotor is coupled to the geartrain through the first component. The compressor rotor is coupled to the geartrain through the second component. The rotating structure is coupled to the geartrain through the third component. The rotating structure includes a turbine rotor. The electric machine is positioned remote from the geartrain. The electric machine includes an electric machine rotor, and the electric machine rotor is coupled to the geartrain through the third component.
F01D 15/12 - Combinaisons avec des transmissions mécaniques
B64D 35/022 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride
14.
ENGINE OIL SYSTEM FOR AN AIRCRAFT PROPULSION SYSTEM AND METHOD FOR CONTROLLING SAME
An engine oil system includes a filter assembly, a sensor assembly, and a controller. The filter assembly includes a filter and a bypass valve. The filter assembly forms a portion of an oil flow path of the engine oil system. The sensor assembly includes an inlet pressure sensor, an outlet pressure sensor, and a temperature sensor. The controller is configured to control a position of the bypass valve in the open position or the closed position with an unfaulted control routine using a temperature (TOIL) measured by the temperature sensor and a differential pressure (ΔPOIL) measured by the inlet pressure sensor and the outlet pressure sensor, execute a bypass valve control algorithm configured to identify a faulted condition and an unfaulted condition of the sensor assembly, and identify the faulted condition or the unfaulted condition using the bypass valve control algorithm. The faulted condition is identified where the temperature (TOIL) or the differential pressure (ΔPOIL) is outside of an expected temperature range or an expected differential pressure range, respectively.
A plenum assembly for a compressor has a plenum housing defining a radially inward entry to a compressor and an outer circumference, and at least two open areas in the outer circumferences which are defined by at least two plenum flanges; an inlet duct connected to a first plenum flange of the at least two plenum flanges; and a panel having a flow control feature, wherein the panel is connected to a second plenum flange of the at least two plenum flanges.
An engine oil system includes a filter assembly. The filter assembly includes a filter manifold and at least one bypass valve. The filter manifold includes a manifold body. The manifold body forms an inlet passage and an outlet passage. The manifold body includes a wall portion. The wall portion forms the inlet passage and the outlet passage. The wall portion forms a bypass channel extending between and to the inlet passage and the outlet passage. The at least one bypass valve includes a housing and a valve member. The housing is disposed on the manifold body. The housing extends circumferentially about a valve axis. The valve member includes a valve body forming a valve plug end. The valve member is positionable along the valve axis in a closed position and an open position. In the closed position the valve plug end is seated on the wall portion obstructing the bypass channel. In the open position the valve plug end is separated from the wall portion.
A de-aerator has: a housing defining an air-oil inlet, an oil outlet, and an air outlet; a rotor within the housing and rotatable about a rotation axis and having: inner and outer rims and blades, passages between the blades and communicating with the air-oil inlet; a conduit extending around the rotation axis and defining a central passage extending axially and communicating with the air outlet, the conduit being radially spaced apart from the outer rim by radial passages communicating with the oil outlet; and openings located inwardly of the blades and communicating with the central passage, wherein one or more of: a cross-sectional area of the central passage taken on a plane transverse to the rotation axis increasing in an axial direction away from the air-oil inlet; and the openings located at least partially radially between an inner face of the conduit and an outer face of the inner rim.
An electric aircraft engine includes an electric motor driving a propulsor. A power source powers the electric motor. An electric compressor supplies compressed air. A nacelle surrounds the electric motor, the power source and the electric compressor. An aircraft and a method are also disclosed.
B64D 13/06 - Aménagements ou adaptations des appareils de conditionnement d'air pour équipages d'aéronefs, passagers ou pour emplacements réservés au fret l'air étant climatisé
B64D 15/00 - Dégivrage ou antigivre des surfaces externes des aéronefs
B64D 27/31 - Aéronefs caractérisés par des groupes moteurs électriques à l'intérieur des ailes ou fixés à celles-ci
B64D 35/026 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques le groupe moteur électrique faisant corps avec l’hélice ou le rotor
A filter assembly for an aircraft engine includes a filter housing defining an inner cavity. The filter housing has a fluid inlet and a fluid outlet for flowing a triboelectric fluid into and out of the inner cavity. A filter bowl is removably attached to the filter housing to seal the inner cavity. A filter is disposed in the inner cavity, the filter electrically coupled to the filter bowl at a first electrical connection and the filter electrically coupled to the filter housing at a second electrical connection distinct from the first electrical connection to thereby indirectly electrically couple the filter bowl to the filter housing via the filter. One or both of the first electrical connection and the second electrical connection provides a degree of freedom for the filter to displace relative to the filter housing or the filter bowl upon engagement with the filter.
B01D 35/31 - Structure du carter de filtre comprenant des dispositions pour la protection de l'environnement, p. ex. dispositifs résistant à la pression
F01M 1/10 - Systèmes de lubrification caractérisés par des dispositifs de ventilation ou d'épuration du lubrifiant, p. ex. des filtres
F02M 37/42 - Installation ou enlèvement des filtres
H01R 13/652 - Dispositions de protection par mise à la terre ou par blindage sur les dispositifs de couplage avec broche, lame ou alvéole de mise à la terre
An assembly for an aircraft includes a wing assembly, a plurality of underwing heat exchangers, and a mounting assembly. The wing assembly includes a wing body. The plurality of underwing heat exchangers are arranged along the wing body in a longitudinal direction. The plurality of underwing heat exchangers includes a first underwing heat exchanger and a second underwing heat exchanger. The first underwing heat exchanger is longitudinally adjacent the second underwing heat exchanger. The mounting assembly includes a first heat exchanger bracket, a second heat exchanger bracket, and a first support bracket. The first heat exchanger bracket is fixedly mounted to the first underwing heat exchanger. The second heat exchanger bracket is fixedly mounted to the second underwing heat exchanger. The first heat exchanger bracket and the second heat exchanger bracket are pivotably mounted to the first support bracket. The first support bracket is fixedly mounted to the wing body.
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
B64C 3/26 - Construction, forme ou fixation des revêtements distincts, p. ex. panneaux
B64C 7/00 - Structures ou carénages non prévus ailleurs
F28F 9/007 - Supports auxiliaires pour les éléments
The process can include: scanning the vane ring, including creating a tri-dimensional digital model of the vane ring in an initial configuration; performing a computerized simulation of airflow across the vane ring based on the tri-dimensional digital model; identifying one or more flow discrepancy based on the computerized airflow simulation; associating a subset of one or more vanes of the set of vanes to the one or more flow discrepancy; and plastically deforming at least one vane of the subset.
G06F 30/28 - Optimisation, vérification ou simulation de l’objet conçu utilisant la dynamique des fluides, p. ex. les équations de Navier-Stokes ou la dynamique des fluides numérique [DFN]
A stator vane is provided that includes an airfoil and a structural member. The structural member is attached to the airfoil. The structural member has a forward surface, an internal passage in fluid communication with the airfoil, and a multi-access cooling air feed passage. The multi-access cooling air feed passage includes a primary aperture in fluid communication with the internal passage and a plurality of secondary passages each configured to provide a fluid passage into the primary aperture. The primary aperture extends along a primary aperture centerline, and each secondary passage of the plurality of secondary passages extends in a direction that diverges from the primary aperture centerline.
A method is disclosed for providing a component. During this method, braze powder is deposited with a substrate. The braze powder is sintered together during the depositing of the braze powder to provide the substrate with sintered braze material. The sintered braze material is heated to melt the sintered braze material and to diffusion bond the sintered braze material to the substrate to provide braze filler material. A first object is scanned using computed tomography to provide first object scan data. The first object includes the substrate and the braze filler material diffusion bonded to the substrate. The first object scan data is compared to first object reference data to provide machining data. The first object is machined using the machining data to provide a second object.
Systems and methods for measuring an axial position of a phonic wheel or other rotating component are provided. The system includes a phonic wheel rotatable about a rotation axis and translatable along the rotation axis, a first sensor, a second sensor and a computer. The phonic wheel includes an inclined tooth having an axially non-uniform radial height and a reference tooth having an axially uniform radial height. The first sensor generates a positioning signal indicative of a gap between the inclined tooth and the first sensor. The second sensor generates a reference signal indicative of a gap between the reference tooth and the reference sensor. The first and second sensors have different orientations. The computer generates an output indicative of the axial position of the phonic wheel based on the positioning signal and the reference signal.
F01D 17/06 - Aménagement des éléments sensibles sensibles à la vitesse
F01D 17/24 - Dispositifs utilisant des éléments sensibles ou des organes de commande terminaux ou les organes de liaison entre les deux, p. ex. commande assistée l'énergie de fonctionnement ou de puissance assistée étant essentiellement non mécanique électrique
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F01D 21/14 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à d'autres conditions spécifiques
G01B 7/02 - Dispositions pour la mesure caractérisées par l'utilisation de techniques électriques ou magnétiques pour mesurer la longueur, la largeur ou l'épaisseur
G01B 7/30 - Dispositions pour la mesure caractérisées par l'utilisation de techniques électriques ou magnétiques pour mesurer des angles ou des cônesDispositions pour la mesure caractérisées par l'utilisation de techniques électriques ou magnétiques pour tester l'alignement des axes
An assembly is provided for an aircraft motor. The assembly includes a bearing, a stationary structure and a flexible support. The bearing extends circumferentially around an axis. The stationary structure circumscribes the bearing. The flexible support is arranged radially between the bearing and the stationary structure. The flexible support includes an inner ring, an outer ring, a bridge, an open first channel and an open second channel. The inner ring radially engages the bearing. The outer ring radially engages the stationary structure. The bridge projects radially out from the inner ring to the outer ring. The bridge extends circumferentially about the inner ring. The open first channel extends axially into the flexible support from a first side of the flexible support to the bridge. The open second channel extends axially into the flexible support from a second side of the flexible support to the bridge.
A casing for a gas turbine engine, including: an inner surface portion; a pair of wall portions extending radially outward from the inner surface portion; and a containment ring located between the pair of wall portions, the containment ring including an inner ring, an intermediate ring, and an outer ring, the intermediate ring being located between the inner ring and the outer ring, wherein the containment outer ring is mechanically decoupled from the pair of wall portions.
F01D 21/04 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à une position incorrecte du rotor par rapport au stator, p. ex. indiquant cette position
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
A casing for a gas turbine engine, including: an inner ring portion; an intermediate layer of honeycomb and/or foam secured to a radially outward surface of the inner ring portion; an outer ring portion secured to a radially outward surface of the layer of honeycomb and/or foam; and a composite containment belt secured to a radially outward surface of the outer ring portion, the composite containment belt having a central portion integrally formed with a forward ring portion and an aft ring portion, the composite containment belt including at least one single woven belt wound about the outer ring portion, the at least one single woven belt including a plurality of continuous fibers, wherein a first portion of the plurality of continuous fibers are at arranged at an angle of 45/−45 degrees or greater with respect to a circumferential direction of the composite containment belt in the central portion and a second portion of the plurality of continuous fibers are at arranged at an angle of less than 45/−45 degrees with respect to the circumferential direction of the composite containment belt in the forward ring portion and the aft ring portion.
F01D 21/04 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à une position incorrecte du rotor par rapport au stator, p. ex. indiquant cette position
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
A non-rotating spacer for use in a mating part of a gas turbine engine, including: a ring of material having a first contact surface and a second contact surface, the first contact surface being opposite to the second contact surface, the ring of material also having an inner ring surface and an outer ring surface; a pair of slots located in the first contact surface, the pair of slots extending from the inner ring surface to the outer ring surface, the pair of slots do not extend though to the second contact surface such that the pair of slots are defined by three surfaces of the ring of material; and at least one lug protruding radially outward from the outer ring surface, the at least one lug being located adjacent to one of the pair of slots.
An injector includes a mount, a stem, a nozzle, and a heat shield. The heat shield extends from the mount to surround the stem and at least a portion of the nozzle. The nozzle includes an exterior annular body. The distal end of the heat shield is spaced from a lip extending from the exterior annular body or a cap surrounding the nozzle to form an axial gap.
An aircraft powerplant apparatus includes a rotating structure configured to rotate about an axis. The rotating structure includes a bladed rotor, a first component, a second component, a first balancing mass, a second balancing mass and a spring element. The first balancing mass and the second balancing mass are configured to rotationally balance the rotating structure about the axis. The first component include a plurality of first component teeth. The second component includes a plurality of second component teeth. The first balancing mass includes a first mass ring and a plurality of first mass teeth. The first mass teeth are meshed with the first component teeth. The second balancing mass includes a second mass ring and a plurality of second mass teeth. The second mass teeth are meshed with the second component teeth. The spring element is engaged with the first balancing mass and the second balancing mass.
A compressor assembly has: an impeller rotatable about a central axis, the impeller mounted on a shaft; a bearing housing extending around the central axis and radially supporting the shaft; and a diffuser downstream of the impeller, the diffuser having: an annular member extending circumferentially around the central axis, vanes circumferentially distributed around the central axis and protruding from one of the annular member and the bearing housing, the vanes extending from bases at the one of the annular member and the bearing housing to tips at the other of the annular member and the bearing housing, and lug and slot connections defined between the annular member and the bearing housing to secure the annular member to the bearing housing.
A turbine exhaust case has an inner ring structurally connected to an outer ring via a plurality of struts. An outer flange projects outwardly from the outer case. The outer flange is made of a material different from that of the outer case and includes a ring portion having an inner diameter surface joined to an outer diameter surface of the outer case at a first axial location corresponding to a junction of a leading edge of the plurality of struts and the outer case. The outer flange further includes a tubular portion projecting axially from the ring portion to a second axial location disposed upstream of the struts, and an annular mounting portion projecting outwardly from an outer surface of the tubular portion at the second axial location.
A method is provided for operating a powerplant. During this method, a hydrocarbon fuel is delivered to a turbine engine for combustion within a combustion chamber of the turbine engine during initial startup operation of the turbine engine. Ammonia is at least partially cracked into hydrogen gas and nitrogen gas. A non-hydrocarbon fuel is delivered to the turbine engine for combustion within the combustion chamber of the turbine engine during post-startup operation of the turbine engine. The non-hydrocarbon fuel is or otherwise includes the hydrogen gas or a combination of the hydrogen gas and the nitrogen gas.
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
F23R 3/36 - Alimentation en combustibles différents
A method includes detecting an asymmetric thrust condition involving first and second engines of an aircraft. The asymmetric thrust condition is associated with one of the engines providing more thrust than another of the engines. The method also includes, in response to detecting the asymmetric thrust condition, determining a modifier for one of the engines and applying the modifier to an acceleration schedule for that engine. The modifier alters the acceleration schedule for that engine to reduce asymmetric thrust between the engines. In some cases, the first engine is associated with a first thrust controller, the second engine is associated with a second thrust controller, the first thrust controller is configured to modify an acceleration schedule of the first engine in response to the asymmetric thrust condition, and the second thrust controller is configured to modify an acceleration schedule of the second engine in response to the asymmetric thrust condition.
B64D 31/10 - Dispositifs amorçant la mise en œuvre actionnés automatiquement en réponse à une défaillance des groupes moteurs pour empêcher une poussée asymétrique
35.
PARTIAL-ADMISSION TURBINE ASSEMBLY FOR AN AIRCRAFT AND METHOD FOR CONTROLLING SAME
An assembly for an aircraft includes a fluid source, a first heat exchanger, a partial-admission turbine, a mechanical load, and a control assembly. The partial-admission turbine includes a rotational assembly and a plurality of partial-admission turbine stages. The rotational assembly includes a bladed turbine rotor. The bladed turbine rotor includes a plurality of rotor blade stages. Each of the plurality of partial-admission turbine stages includes a respective rotor blade stage of the plurality of rotor blade stages. The fluid source, the first heat exchanger, and the partial-admission turbine sequentially form a portion of a fluid flow path through the assembly. The mechanical load is coupled to the rotational assembly. The control assembly including a controller configured to determine a corrected turbine inlet flow and a corrected rotation speed for the partial-admission turbine, maintain the corrected turbine inlet flow within an inlet corrected flow threshold range, and maintain the corrected rotation speed within a corrected rotation speed threshold range.
F02C 9/48 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
A casing for a gas turbine engine, including: a plurality of layers of a structural material, wherein each layer of the plurality of layers of the structural material are separated from each other by at least one layer of an impact absorbing material, the at least one layer of impact absorbing material being different from the structural material and the impact absorbing material having a radial thickness that is greater than a radial thickness of the structural material, wherein the plurality of layers of the structural material comprises at least three layers.
A casing for a gas turbine engine, including: an inner ring portion; an outer ring portion; and a plurality of blunting plates arranged circumferentially between the inner ring portion and the outer ring portion, wherein the plurality of blunting plates are positioned to overlap each other.
Systems and methods for performing dimensional inspection of components are provided. The systems and methods correct errors associated with non-contact scanning of components based on surface quality. A method of inspecting a component includes using non-contact scanning of the component to acquire a point cloud indicative of a geometry of the component, acquiring a surface quality of the component, and selecting a geometric transformation associated with the surface quality of the component from a plurality of geometric transformations respectively associated with different surface qualities. The point cloud is transformed using a selected geometric transformation associated with the surface quality of the component.
A cooling fan housing is provided. The cooling fan housing includes an inner housing and an outer housing. The inner housing includes inner forward and inner rearward flanges on either side of an inner annulus. The inner rearward flange includes a recessed portion proximate to the inner annulus and a non-recessed portion remote from the inner annulus. The outer housing includes outer forward and outer rearward flanges on either side of an outer annulus. The outer and inner forward flanges are attachable in a tight-fitting condition. The outer rearward flange and the non-recessed portion of the inner rearward flange are attachable in a tight-fitting condition with the recessed portion of the inner rearward flange forming a gap with the outer rearward flange.
A heat management system for a thermal engine includes a source of hydrogen, a first hydrogen conduit flowing a first flow of hydrogen therethrough, the first hydrogen conduit fluidly coupling the source of hydrogen to a fuel system of the thermal engine, and a second hydrogen conduit flowing a second flow of hydrogen therethrough. The second hydrogen conduit is in thermal communication with an engine exhaust conduit to transfer heat from the engine exhaust conduit to the second hydrogen conduit. The second hydrogen conduit is also in thermal communication with the first hydrogen conduit to transfer heat from the second hydrogen conduit to the first hydrogen conduit.
F02C 7/224 - Chauffage du combustible avant son arrivée au brûleur
F02C 3/22 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion le combustible ou l'oxydant étant gazeux aux température et pression normales
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
41.
NON-LINEAR STIFFNESS BRACKET OF GAS TURBINE ENGINE
A mounting bracket of a gas turbine engine includes a first mounting flange configured to connect the mounting bracket to a casing of the gas turbine engine, a second mounting flange configured to connect a component to the mounting bracket, and a bracket leg extending between the first mounting flange and the second mounting flange. The mounting bracket has a first stiffness when a load is applied in a first force direction and has a second stiffness different from the first stiffness when the load is applied in a second force direction opposite the first force direction.
F01D 25/28 - Dispositions pour le support ou le montage, p. ex. pour les carters de turbines
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
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
42.
HEAT EXCHANGER HAVING A MIXING CHAMBER WITH CONVERGENT SECTION
A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; and a mixing chamber intersecting two or more of the first conduits and separating the first conduits into upstream sections and downstream sections, the mixing chamber having a peripheral wall extending around a mixing volume, the peripheral wall defining a convergent section in which a flow circulating area of the mixing volume decreases in a downstream direction relative to the flow through the first conduits.
F28F 13/12 - Dispositions pour modifier le transfert de chaleur, p. ex. accroissement, diminution en affectant le mode d'écoulement des sources de potentiel calorifique en créant une turbulence, p. ex. par brassage, par augmentation de la force de circulation
F28D 21/00 - Appareils échangeurs de chaleur non couverts par l'un des groupes
F28F 1/40 - Éléments tubulaires ou leurs ensembles avec moyens pour augmenter la surface de transfert de chaleur, p. ex. avec des ailettes, avec des saillies, avec des évidements les moyens étant uniquement à l'intérieur de l'élément tubulaire
A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and fins protruding from the peripheral wall into the mixing volume, the fins extending in a direction having an axial component relative to the central axis, the fins defining flow passages interspaced between the fins.
F28F 13/12 - Dispositions pour modifier le transfert de chaleur, p. ex. accroissement, diminution en affectant le mode d'écoulement des sources de potentiel calorifique en créant une turbulence, p. ex. par brassage, par augmentation de la force de circulation
F28D 21/00 - Appareils échangeurs de chaleur non couverts par l'un des groupes
F28F 1/40 - Éléments tubulaires ou leurs ensembles avec moyens pour augmenter la surface de transfert de chaleur, p. ex. avec des ailettes, avec des saillies, avec des évidements les moyens étant uniquement à l'intérieur de l'élément tubulaire
44.
Heat exchanger having a mixing chamber and protrusions
A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and protrusions extending from the peripheral wall into the mixing volume, the protrusions extending around the central axis and across a direction of the flow through the mixing chamber.
F28F 13/12 - Dispositions pour modifier le transfert de chaleur, p. ex. accroissement, diminution en affectant le mode d'écoulement des sources de potentiel calorifique en créant une turbulence, p. ex. par brassage, par augmentation de la force de circulation
A propulsion system for an aircraft includes a first hybrid power plant and a second hybrid power plant, each including a thermal engine and an electrical motor. Acoustic sensors measure an initial combined noise signature, produced by the propulsion system, including first and second noise signatures respectively generated by the first and second hybrid power plants. A controller receives a signal from the acoustic sensors and determines when an initial amplitude variation of a periodically fluctuating amplitude of the initial combined noise signature is greater than an amplitude variation threshold indicative that the initial combined noise signature generates beats. A thrust produced by the second hybrid power plant is modulated to produce a modulated combined noise signature having a modulated amplitude variation less than the initial amplitude variation. A difference in thrusts generated by the first and second hybrid power plants is compensated for.
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64D 35/025 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride du type en parallèle
A seal arrangement of a gas turbine engine includes two seal assemblies arranged in a serial relationship relative to an engine central longitudinal axis. Each seal assembly includes a seal carrier, and a primary seal element positioned in the seal carrier and configured to seal to a rotating component of the gas turbine engine abutting the primary seal element. A seal sleeve is positioned radially outboard of the seal carriers of the two seal assemblies. The seal sleeve is configured to retain the two seal assemblies in position abutting the rotating component.
A system for an aircraft engine, has: an aircraft component; and a heat exchanger having: a housing defining a first inlet, a first outlet, a second inlet, and a second outlet; first conduits within the housing, the first conduits fluidly connecting the first inlet to the first outlet; one or more second conduit within the housing, the one or more second conduits fluidly connecting the second inlet to the second outlet, the one or more second conduit in heat exchange relationship with the first conduits; a mixing chamber intersecting two or more of the first conduits, the mixing chamber having a peripheral wall extending around a mixing volume and a central axis; and louvers mounted to the peripheral wall and extending transversally to the central axis, the louvers movable between a collapsed configuration and a deployed configuration and extending across the mixing volume in the deployed configuration.
F28F 1/32 - Éléments tubulaires ou leurs ensembles avec moyens pour augmenter la surface de transfert de chaleur, p. ex. avec des ailettes, avec des saillies, avec des évidements ces moyens étant uniquement à l'extérieur de l'élément tubulaire et s'étendant transversalement les moyens ayant des parties engageant d'autres éléments tubulaires
F28D 21/00 - Appareils échangeurs de chaleur non couverts par l'un des groupes
F28F 13/12 - Dispositions pour modifier le transfert de chaleur, p. ex. accroissement, diminution en affectant le mode d'écoulement des sources de potentiel calorifique en créant une turbulence, p. ex. par brassage, par augmentation de la force de circulation
48.
Exhaust nozzle assembly for an aircraft propulsion system
An aircraft propulsion system includes a gas turbine engine and an exhaust nozzle assembly. The gas turbine engine includes a compressor section, a combustor section, a turbine section, and an exhaust section. The compressor section, the combustor section, the turbine section, and the exhaust section form a core flow path through the gas turbine engine. The exhaust nozzle assembly is disposed at the exhaust section. The exhaust nozzle assembly includes a nozzle bypass system and an exhaust treatment system. The nozzle bypass system is selectively configurable in a bypass mode or an exhaust treatment mode. The nozzle bypass system in the bypass mode directs a core combustion gas from the core flow path through the exhaust nozzle assembly bypassing the exhaust treatment system. The nozzle bypass system in the exhaust treatment mode directs the core combustion gas through the exhaust nozzle assembly and exhaust treatment system.
F02K 1/00 - Ensembles fonctionnels caractérisés par la forme ou la disposition de la tubulure de jet ou de la tuyèreTubulures de jet ou tuyères particulières à cet effet
F02C 3/20 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail utilisant un combustible, un oxydant ou un fluide de dilution particulier pour produire les produits de combustion
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
F02C 9/40 - Commande de l'alimentation en combustible spécialement adaptée à l'utilisation d'un combustible particulier ou de plusieurs combustibles
F02C 9/52 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel avec la commande du flux du fluide de travail par prélèvement ou bipasse du fluide de travail
F02K 1/82 - Parois des tubulures de jet, p. ex. chemises
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
49.
MODEL-BASED DEFINITION MANUFACTURING AND INSPECTION SYSTEM INCLUDING DATA MODEL
A computer system includes a controller having a computer readable storage medium having program instructions embodied therewith, the program instructions executable by controller to perform operations including generating a three-dimensional (3D) model of a product, extracting data indicative of one or more elements of the product, and assigning a FAE address to the one or more elements. The method further includes generating a grouping tree to organize the extracted elements, each extracted element organized according to the FAE address, and adding product and manufacturing information (PMI) data to one or more of the addressed FAEs in the grouping tree.
G06F 30/12 - CAO géométrique caractérisée par des moyens d’entrée spécialement adaptés à la CAO, p. ex. interfaces utilisateur graphiques [UIG] spécialement adaptées à la CAO
G06F 30/20 - Optimisation, vérification ou simulation de l’objet conçu
G06F 119/18 - Analyse de fabricabilité ou optimisation de fabricabilité
50.
Hybrid electric propulsion system with regeneration and method for operating the same
A method of regenerating an electric power storage unit for an aircraft hybrid electric propulsion (HEP) system is provided. The HEP system includes a thermal engine, an electric motor, an electric power storage unit, a primary gearbox, and a propulsion unit. The electric motor is configured to selectively provide rotational drive to the propulsion unit. The method includes: controlling the electric motor to operate as the generator; controlling the primary gearbox to be in an output disengaged configuration, wherein in the output disengaged configuration the primary gearbox is disengaged from driving the propulsion unit and the primary gearbox is transfers rotational drive produced by the thermal engine to the electric motor; operating the thermal engine to drive the primary gearbox in the output disengaged configuration; and regenerating the electric power storage unit by providing electrical energy produced by the electric motor to the electric power storage unit.
H02P 9/06 - Commande s'exerçant sur un embrayage ou un autre moyen mécanique de transmission de la puissance et dépendant de la valeur d'une caractéristique électrique à la sortie de la génératrice
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64D 35/022 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride
H02K 7/108 - Association structurelle avec des embrayages, des freins, des engrenages, des poulies ou des démarreurs mécaniques avec des embrayages à friction
H02K 7/116 - Association structurelle avec des embrayages, des freins, des engrenages, des poulies ou des démarreurs mécaniques avec des engrenages
H02K 7/18 - Association structurelle de génératrices électriques à des moteurs mécaniques d'entraînement, p. ex. à des turbines
H02P 101/25 - Adaptation particulière des dispositions pour la commande de génératrices pour moteurs à combustion
H02P 101/30 - Adaptation particulière des dispositions pour la commande de génératrices pour aéronefs
51.
Electromagnetic inducer for gas turbine rotating turbomachinery
Gas turbine engines include fan, compressor, combustor, and turbine sections. An electric machine assembly includes a rotating component comprising a plurality of airfoils and a plurality of first electric machine elements, with each airfoil including a respective first magnetic material electric machine element. A housing is arranged radially outward from the airfoils and includes a circumferential slot. A second electric machine element is arranged within the circumferential slot in the form of a circumferential winding arranged radially outward from the rotating component. A control system is electrically connected to the second electric machine element. Rotation of the rotating component causes the first electric machine elements to generate a magnetic flux and induce a current within the second electric machine element and the control system is configured to distribute the induced current to at least one of a storage device or electrical components of the gas turbine engine.
An assembly includes a rotating seal land, a seal element, a stationary seal support and a mating plate. The seal element axially overlaps, circumscribes and radially sealingly engages the rotating seal land. The stationary seal support axially overlaps and circumscribes the rotating seal land and the seal element. The stationary seal support includes a support notch, a support shoulder, a support inner interior surface and a support outer interior surface. The support notch projects axially into the stationary seal support from a distal end of the stationary seal support to the support shoulder. The support notch projects radially outward away from an axis into the stationary seal support from the support inner interior surface to the support outer interior surface. The mating plate is seated in the support notch and mounted to the stationary seal support. The mating plate axially contacts the support shoulder and the seal element.
A method for fabricating a component from a hybrid raw material, comprising the steps of: forging a first portion of a raw material for the component; additive manufacturing a second portion of a raw material to the first portion to produce a hybrid raw material; and machining the hybrid raw material to produce the component. A repair method is also disclosed.
B22F 7/08 - Fabrication de couches composites, de pièces ou d'objets à base de poudres métalliques, par frittage avec ou sans compactage de pièces ou objets composés de parties différentes, p. ex. pour former des outils à embouts rapportés avec une ou plusieurs parties non faites à partir de poudre
B22F 10/66 - Traitement de pièces ou d'articles après leur formation par des moyens mécaniques
The method can include operating the hybrid aircraft power plant in a first mode of operation in which a plurality of controllable parameters including a fuel flow to the thermal engine and an electric power to the electric motor has a first set of values corresponding to a first power output; receiving a requested power output; determining a second set of values of the one or more of the controllable parameters to bring first power output to the requested power output; determining whether said second set of values poses a risk of undesired clutch behavior; and when the second set of values poses the risk, modulating the second set of values to avoid the risk and setting the plurality of controllable parameters to the modulated second set of values.
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64C 11/30 - Mécanismes de changement de pas des pales
An apparatus is provided for a powerplant. This powerplant apparatus includes a fluid filter assembly with a filter housing. The filter housing extends axially along an axis between a housing first end and a housing second end. The filter housing includes a mount and a sidewall. The mount is connected to the sidewall. The mount includes a flange, an attachment and a pilot. The flange projects radially out from the sidewall and the attachment to a flange distal outer end. The attachment extends axially between the flange and the pilot. The attachment is configured with an external threaded section. The pilot projects axially out from the attachment to the housing first end. The pilot includes a cylindrical external surface located radially inboard of the external threaded section. An axial length of the pilot is greater than an axial length of the attachment.
F16M 7/00 - Détails de fixation ou de réglage des bâtis, châssis ou pièces de supports des moteurs sur leurs fondations ou leur baseFixation des parties fixes des moteurs, p. ex. des blocs cylindres
B01D 29/11 - Filtres à éléments filtrants stationnaires pendant la filtration, p. ex. filtres à aspiration ou à pression, non couverts par les groupes Leurs éléments filtrants avec des éléments filtrants en forme de sac, de cage, de tuyau, de tube, de manchon ou analogue
A vane including an airfoil body extending from a root to a tip defining a longitudinal axis therebetween, wherein the airfoil body includes a leading edge between the root and the tip. The vane includes a pressure side and a suction side and a recess area on the leading edge. The recess area includes a pressure side recess area and a suction side recess area, and a metal is deposited on the recess area. A method of manufacturing the vane and a gas turbine engine including the vane are also disclosed.
A gas turbine engine includes a turbine section, a turbine case, an inner case, and a damping seal baffle. The turbine section extends along a rotational axis of the gas turbine engine. The turbine case includes a cantilevered inner wall. The cantilevered inner wall extends circumferentially about the rotational axis. The cantilevered inner wall extends between and to an upstream axial end and a downstream axial end. The upstream axial end is disposed at and downstream of the turbine section. The cantilevered inner wall forms a first sealing surface and a second sealing surface. The inner case extends and the damping seal baffle extend circumferentially about the rotational axis. The damping seal baffle is mounted to the inner case. The damping seal baffle is disposed in contact with one or both of the first sealing surface or the second sealing surface.
A turboshaft engine is provided and includes a turbine shaft, an output shaft and a combined speed ratio stage and torque measuring system. The combined speed ratio stage and torque measuring system includes a rigid assembly, a sensor system and first and second bearings. The rigid assembly includes a torque tube, a first gear-speed ratio stage to transmit gear-reduced or gear-increased torque from the turbine shaft and to the torque tube and a second gear-speed ratio stage to transmit gear-reduced or gear-increased torque from the torque tube to the output shaft. The sensor system measures twist of the torque tube. The first and second bearings support the rigid assembly.
G01L 5/12 - 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 la poussée axiale d'un arbre tournant, p. ex. matériel de propulsion
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F16H 1/20 - Transmissions à engrenages pour transmettre un mouvement rotatif sans engrenages à mouvement orbital comportant plus de deux organes engrenés
F16H 57/04 - Caractéristiques relatives à la lubrification ou au refroidissement
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
59.
System and method for applying trim balance to a module of the gas turbine engine
A system for applying balancing trim to a module of a gas turbine engine, the module having a compressor having a forward shaft and an aft hub, a turbine connected to the aft hub, and the turbine having a turbine case and an aft shaft, the system having: a support bracket that mounts to the turbine case and supports the module when testing for unbalances, wherein the support bracket has: an outer periphery; outer mounting apertures at the outer periphery, wherein the outer mounting apertures are sized to receive fasteners to connect the support bracket to the turbine case; a center aperture defined by an inner periphery, wherein the center aperture is sized to receive the aft hub; and weight positioning slots defined between the outer mounting apertures and the center aperture; and wherein the outer periphery and the inner periphery are axially spaced apart from each other.
An assembly for an aircraft propulsion system includes an engine, an aircraft aerostructure, and a heat exchanger. The engine includes a coolant system. The aircraft aerostructure forms a fuel tank for the engine. The aircraft aerostructure includes an aerostructure skin including an interior skin side and an exterior skin side. The interior skin side forms a bottom side of the fuel tank. The exterior skin side forms an aerodynamic surface of the aircraft aerostructure. The heat exchanger is disposed at the aerostructure skin. The heat exchanger includes a heat exchanger body forming a coolant passage. The coolant passage extends through the heat exchanger body between and to an inlet and an outlet. The inlet and the outlet are connected in fluid communication with the coolant system.
A housing assembly has: a rotor housing extending around a rotation axis and from a first side to a second side, and having an inner face; a housing secured to the first side; and a seal received within a groove at an interface between the rotor housing and the housing, the groove extending around a perimeter of the first side and located radially outwardly of the inner face, the seal including: an elastomeric member; and a shield disposed inwardly of the elastomeric member and in contact with both the housing and the rotor housing, the shield including a thermoset plastic member extending along at least a portion of a perimeter of the groove and axially overlapping the elastomeric member, the thermoset plastic member circumferentially overlapping a combustion region of the rotor cavity where combustion occurs during operation of the rotary internal combustion engine.
A machining system includes a WEDM assembly, a probe assembly, a calibration artifact, and a controller. The WEDM assembly includes a conductive wire. The conductive wire extends between and to the upper guide head and the lower guide head. The probe assembly includes a touch probe. The calibration artifact includes an artifact body forming a probe cavity and a wire cavity. The controller is configured to identify a probe X-position and a probe Y-position of the touch probe by controlling the WEDM assembly to move the touch probe to contact the calibration artifact with the touch probe disposed within the probe cavity, identify a wire X-position and a wire Y-position of the conductive wire by controlling the WEDM assembly to move the conductive wire to contact the calibration artifact with the conductive wire disposed within the wire cavity, and calibrate the machining system by determining an X-offset and a Y-offset of the conductive wire relative to the touch probe and storing the X-offset and the Y-offset.
G05B 19/401 - Commande numérique [CN], c.-à-d. machines fonctionnant automatiquement, en particulier machines-outils, p. ex. dans un milieu de fabrication industriel, afin d'effectuer un positionnement, un mouvement ou des actions coordonnées au moyen de données d'un programme sous forme numérique caractérisée par des dispositions de commande pour la mesure, p. ex. étalonnage et initialisation, mesure de la pièce à usiner à des fins d'usinage
A riveting yoke is provided that includes a body, an anvil segment, and a cavity. The body has a height extending between a base side surface and a top side surface, a width extending between a first end surface and a second end surface, and a thickness extending between a first lateral side surface and a second lateral side surface. The anvil segment (AS) has an AS top surface and an AS base surface. The cavity is disposed in the first end surface, and is defined by a cavity base surface, a cavity inner surface, a gusset surface, and the AS base surface. The gusset surface extends from the cavity inner surface to the AS base surface. The cavity base surface is opposite the AS base surface and a distance between the cavity base surface and the AS base surface defines an opening of the cavity.
A telemetry module system for mounting on a rotatable shaft is provided that includes a telemetry module and a connector ring. The telemetry module includes an annular TM housing, a plurality of transmitters, and a first connectivity board. The TM housing extends axially between first and second TM axial ends. Each transmitter is configured to produce RF signals based on sensor input signals. The first connectivity board has a ring configuration and is in signal communication with the transmitters. The connector ring (CR) has an annular CR housing and a second connectivity board. The CR housing extends axially between first and second CR axial ends. The second connectivity board has a ring configuration and is attached to the CR housing. The second connectivity board is configured to receive sensor input signals. The telemetry module and the connector ring are configured to be disposed in a coupled configuration.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F02C 7/00 - 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
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
H04Q 9/00 - Dispositions dans les systèmes de commande à distance ou de télémétrie pour appeler sélectivement une sous-station à partir d'une station principale, sous-station dans laquelle un appareil recherché est choisi pour appliquer un signal de commande ou pour obtenir des valeurs mesurées
A method includes obtaining multiple parameters associated with operation of a turbine engine. The method also includes applying a model predictive control algorithm to the parameters to predict a future torque value of a shaft of the turbine engine, the future torque value indicating a likelihood of future shaft shear. The method further includes, in response to the future torque value being outside of an expected range, performing an action to reduce risk of the future shaft shear.
G05B 13/04 - Systèmes de commande adaptatifs, c.-à-d. systèmes se réglant eux-mêmes automatiquement pour obtenir un rendement optimal suivant un critère prédéterminé électriques impliquant l'usage de modèles ou de simulateurs
An assembly is provided for an aircraft propulsion system. This assembly includes an outer wall structure and an inner wall structure. The outer wall structure extends axially along and circumferentially about an axis. The outer wall structure forms an outer peripheral boundary of a flowpath. The inner wall structure extends axially along and circumferentially about the axis. The inner wall structure is spaced radially in from the outer wall structure. The inner wall structure includes a first case, a second case and a shield. The first case forms a first section of an inner peripheral boundary of the flowpath. The second case is attached to the first case at a joint. The shield covers the joint and is arranged radially between the joint and the flowpath.
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
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
F23R 3/00 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux
67.
SYSTEM AND METHOD FOR ESTIMATING A SERVICE LIFE OF A COMPONENT OF AN AIRCRAFT POWER PLANT
Methods and systems for estimating a remaining service life of a component of an aircraft power plant are provided. A method includes inducing a vibration in the component, and measuring a response to the induced vibration in the component. A computer-implemented trained model is used to determine an estimated remaining service life for the component based on the measured response. An output indicative of the estimated remaining service life of the component is generated.
A lubrication system of a gas turbine engine includes a lubricant reservoir with a housing having a volume of oil located therein. A reservoir passage extends through the housing. One or more side passages extend from the reservoir passage and are configured to selectably fill or drain the volume of oil. An output passage is separate and distinct from the reservoir passage, and a separator wall extends into the interior of the housing from the first housing wall toward the second housing wall and between the reservoir passage and the output passage. When the lubricant reservoir is in a first orientation, the lubricant reservoir is filled via oil flowed through the reservoir passage from the first housing end, and when operated in a negative G orientation, the lubricant reservoir is drained via oil flowed through the reservoir passage and out of the second housing end.
F01M 11/06 - Dispositifs pour maintenir constant le niveau du lubrifiant ou pour l'affranchir du mouvement ou de la position de la "machine" ou du moteur
F16H 57/04 - Caractéristiques relatives à la lubrification ou au refroidissement
F16N 17/00 - Lubrification des machines ou des appareils travaillant dans des conditions très sévères
An aircraft engine, comprising: a thermal engine having an outlet; a heat exchanger having a first conduit connected to a source of a fluid, and a second conduit connected to the outlet; an inlet conduit connecting the source to the first conduit; an outlet conduit connected to an outlet of the first conduit; a component connected to the outlet conduit; a bypass conduit and a recirculation conduit both connecting the inlet conduit to the outlet conduit while bypassing the heat exchanger; a valve communicating with the recirculation conduit and the bypass conduit, the valve operable to selectively allow the fluid to flow in one of the recirculation conduit and the bypass conduit while limiting the fluid from flowing in the other of the recirculation conduit and the bypass conduit; and a flow inducer communicating with the recirculation conduit for inducing a flow from the outlet conduit to the inlet conduit.
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F28F 27/02 - Commandes ou dispositifs de sécurité spécialement adaptés pour les appareils d'échange ou de transfert de chaleur pour commander la répartition des sources de potentiel calorifique entre des canaux différents
70.
HYBRID ELECTRIC AIRCRAFT PROPULSION SYSTEM AND METHOD
There is provided a hybrid electric aircraft propulsion system and method for operating same. The method comprises providing, to a first electric motor and a second electric motor, alternating current (AC) electric power from a generator, the generator receiving rotational power from a thermal engine, providing, to the first electric motor and the second electric motor, AC electric power from at least one motor inverter, the at least one motor inverter configured to convert DC electric power from a DC power source into AC electric power, and selectively driving the first and second electric motors from the generator, the at least one motor inverter, or a combination thereof, wherein the first electric motor drives a first rotating propulsor and the second electric motor drives a second rotating propulsor.
H02P 27/06 - Dispositions ou procédés pour la commande de moteurs à courant alternatif caractérisés par le type de tension d'alimentation utilisant une tension d’alimentation à fréquence variable, p. ex. tension d’alimentation d’onduleurs ou de convertisseurs utilisant des convertisseurs de courant continu en courant alternatif ou des onduleurs
B60L 50/13 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant la puissance de propulsion fournie par des générateurs entraînés par le moteur, p. ex. des générateurs entraînés par des moteurs à combustion utilisant des générateurs à courant alternatif et des moteurs à courant alternatif
B60L 50/51 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant de la puissance de propulsion fournie par des batteries ou des piles à combustible caractérisée par des moteurs à courant alternatif
B60L 50/60 - Propulsion électrique par source d'énergie intérieure au véhicule utilisant de la puissance de propulsion fournie par des batteries ou des piles à combustible utilisant de l'énergie fournie par des batteries
B60L 53/24 - Procédés de chargement de batteries spécialement adaptées aux véhicules électriquesStations de charge ou équipements de charge embarqués pour ces batteriesÉchange d'éléments d’emmagasinage d'énergie dans les véhicules électriques caractérisés par des convertisseurs situés dans le véhicule utilisant le convertisseur de propulsion du véhicule pour la charge
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 27/12 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz à l'intérieur des ailes ou fixés à celles-ci
B64D 27/40 - Aménagements pour le montage de groupes moteurs sur aéronefs
B64D 31/09 - Dispositifs amorçant la mise en œuvre actionnés automatiquement en réponse à une défaillance des groupes moteurs
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64D 35/00 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions
B64D 35/023 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride du type en série-parallèle
B64D 35/024 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride du type en série
B64D 35/025 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride du type en parallèle
B64D 35/08 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions caractérisée par le fait que la transmission est entraînée par plusieurs groupes moteurs
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
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
G08C 19/38 - Systèmes de transmission de signaux électriques utilisant des dispositifs dynamo-électriques
H02K 51/00 - Transmissions dynamo-électriques, c.-à-d. moyens dynamo-électriques pour la transmission de la puissance mécanique d'un arbre conducteur à un arbre conduit, comprenant des éléments moteur et générateur en corrélation constructive
H02P 5/74 - Dispositions spécialement adaptées à la régulation ou la commande de la vitesse ou du couple d’au moins deux moteurs électriques pour commander au moins deux moteurs dynamo-électriques à courant alternatif
H02P 6/00 - Dispositions pour commander les moteurs synchrones ou les autres moteurs dynamo-électriques utilisant des commutateurs électroniques en fonction de la position du rotorCommutateurs électroniques à cet effet
H02P 6/12 - Contrôle de la commutationIndication d'un défaut de commutation
H02P 9/00 - Dispositions pour la commande de génératrices électriques de façon à obtenir les caractéristiques désirées à la sortie
H02P 17/00 - Dispositions pour la commande des transmissions dynamo-électriques
H02P 29/024 - Détection d’un défaut, p. ex. court circuit, rotor bloqué, circuit ouvert ou perte de charge
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
H02P 101/25 - Adaptation particulière des dispositions pour la commande de génératrices pour moteurs à combustion
H02P 101/30 - Adaptation particulière des dispositions pour la commande de génératrices pour aéronefs
71.
ROTORCRAFT WITH ANTI-TORQUE AIR SYSTEM AND TAIL ROTOR(S)
An assembly is provided for a rotorcraft. This rotorcraft assembly includes a fuselage, a tail structure, an air system and a plurality of tail rotors. The tail structure is configured as or otherwise includes a tail boom. The tail boom projects longitudinally along a centerline out from the fuselage to a distal end. The air system includes an air flowpath. The air flowpath passes from the fuselage into the tail boom and extends longitudinally within the tail boom towards the distal end. The tail rotors are connected to the tail structure at the distal end.
B64C 27/82 - GiravionsRotors propres aux giravions caractérisés par l'existence d'un rotor auxiliaire ou d'un dispositif à jet fluide pour contrebalancer le couple du rotor de sustentation ou faire varier la direction du giravion
72.
SYSTEM AND METHOD FOR GENERATING A PRELIMINARY DESIGN OF A STRUCTURAL ARCHITECTURE FOR AN AIRCRAFT PROPULSION SYSTEM
A method for generating a preliminary design of a propulsion system structural architecture for an aircraft propulsion system includes generating, with an artificial intelligence (AI) model at a computer system, one or both of at least one two-dimensional image or at least one three-dimensional model of the preliminary design including a selected combination of geometric parameters. The AI model has been trained for both a propulsion system geometry using the geometric parameters extracted from historical geometry data and a propulsion system structural design using the geometric parameters and operational parameters of historical operational data. The operational parameters are associated with the geometric parameters. Training the AI model included identifying the selected combination of geometric parameters of the preliminary design, with the AI model, by determining a plurality of combinations of the extracted geometric parameters and the associated operational parameters which satisfy each of at least one technical constraint and at least one customer constraint and selecting the selected combination of geometric parameters from the plurality of combinations of the extracted geometric parameters using the associated operational parameters.
An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet connected to a source of combustion gases and an outlet fluidly connected to the turbine, and a conduit extending around the central axis from the inlet to the outlet; a bearing housing; a turbine support case secured to the bearing housing, the scroll case disposed axially between the turbine support case and the bearing housing, the turbine support case having spokes extending through the scroll case and radially supported by the bearing housing; and heat shields distributed around the central axis in circumferential alignment with the spokes and extending axially through the scroll case for thermally shielding the spokes, a heat shield of the heat shields disposed around a spoke of the spokes, the heat shield defining an air gap surrounding the spoke inside the scroll case.
A variable pitch stator vane stage of a gas turbine engine includes a plurality of stator vanes extending across a flowpath of the gas turbine engine, and an actuation assembly operably connected to the plurality of start vanes. The actuation assembly includes an actuation ring extending circumferentially around an engine central longitudinal axis of the gas turbine engine, and a plurality of vane links. Each vane link operably connects a corresponding stator vane of the plurality of stator vanes to the actuation ring. A vane link of the plurality of vane links includes a link body, and a first link end connected the link body. The first link end includes a first spherical connection to the actuation ring. Circumferential movement of the actuation ring about the engine central longitudinal axis urges rotation of the plurality of vanes about their respective vane axes.
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
75.
AIRCRAFT ACOUSTIC PANEL WITH INTEGRATED FIRE-RESISTANT MATERIAL
An apparatus is provided for an aircraft. This apparatus includes an acoustic panel, and the acoustic panel includes a face skin, a back skin and a cellular core disposed between and connected to the face skin and the back skin. The back skin laterally overlaps the face skin and includes an aerogel layer. The cellular core includes a plurality of cavities that extend through the cellular core from the face skin to the back skin. Each of the cavities is fluidly coupled with one or more perforations in the face skin.
A rotor blade includes an airfoil, a root, a cooling air passage and a cooling air deflector. The root extends axially along an axis between a first end of the root and a second end of the root. The root extends laterally between a first side of the root and a second side of the root. The root projects radially inward and away from the airfoil to an inner end of the root. The cooling air passage includes a passage inlet disposed at the inner end. The cooling air passage projects radially into the rotor blade from the passage inlet. The cooling air deflector projects radially inward from the inner end of the root to an inner end of the cooling air deflector. The cooling air deflector is disposed at the second side and is spaced laterally from the first side. The cooling air deflector includes a cooling air aperture.
A de-aerator is provided that includes a body, a cover panel, a fluid inlet, a helical fluid passage, at least one partition, and at least one fluid outlet. The body extends between opposing first and second axial ends. The body has a sidewall, a base panel, and an internal cavity. The fluid inlet is in communication with the body and is configured to direct fluid tangentially into the internal cavity. The helical fluid passage is disposed within the internal cavity and has entry and exit ends. The helical fluid passage has circumferential turns that each include an air passage. The at least one partition is disposed within the internal cavity at or below the passage exit end, spaced above the base panel. The at least one fluid outlet is configured to permit liquid passage from the internal cavity of the body to outside the body.
A method of assessing the quality of a manufactured component includes the steps of manufacturing a component and generating data at each of a plurality locations through the component. The generated data is passed to a machine learning branch wherein the generated data is compared to training data at each of the plurality of locations to determine whether the component is of functionally tolerant dimensions at each of the plurality of locations. The manufactured component is accepted should the component be within functionally tolerant dimensions at each of the plurality of locations, and the component is rejected if the component fails to be within functionally tolerant dimensions respectively, at each of the plurality of locations. A system is also disclosed.
A method of assessing the quality of a manufactured component includes the steps of manufacturing a component and generating data at each of a plurality of locations on the component. The generated data is passed to a machine learning branch wherein the generated data is compared to training data at each of the plurality of locations and across a plurality of folds using K-fold validation to determine whether the component is of a functionally tolerant dimension at each of the plurality of sections. A system is also disclosed.
A method of assessing the quality of a manufactured component includes the steps of manufacturing a component and generating data at each of a plurality of locations through the component. The generated data is passed to a machine learning branch. The generated data is compared to training data at each of the plurality of locations and across a plurality of folds using K-fold validation to determine whether the component is of a functionally tolerant dimension at each of the plurality of locations. The training data at each of the plurality of folds at each of the plurality of locations is from a common part. A system is also disclosed.
A method of assessing the quality of a manufactured component includes the steps of manufacturing a component and generating data at each of a plurality of locations on the component. The generating data is passed to a machine learning branch. The generated data is compared to training data at each of the locations to determine whether the component is a functionally tolerant dimension at each of the plurality of locations. The manufactured component is accepted or rejected based upon a determined percentage chance the component is functionally tolerant or fails to be functionally tolerant, respectively, at each of the plurality of locations. A system is also disclosed.
An aircraft propulsion system is provided that includes a thermal engine and an electrical energy generating device. The electrical energy generating device has first and second fluid conduits and a thermoelectric generator (TEG). The TEG is disposed between the first and second fluid conduits. A first side of the TEG is in thermal communication with the first fluid conduit, and a second side of the TEG is in thermal communication with the second fluid conduit. The TEG generator is configured to produce electrical energy as a function of a temperature difference across the TEG. The first fluid conduit contains a first fluid flow, and the second fluid conduit contains a second fluid flow. During operation of the propulsion system, the first fluid flow is at a first higher temperature and the second fluid flow is at a second lower temperature, thereby producing the temperature difference across the TEG.
H10N 10/17 - Dispositifs thermoélectriques comportant une jonction de matériaux différents, c.-à-d. dispositifs présentant l'effet Seebeck ou l'effet Peltier fonctionnant exclusivement par les effets Peltier ou Seebeck caractérisés par la structure ou la configuration de la cellule ou du thermocouple constituant le dispositif
83.
RAPID PART ASSESSMENT WITH TESTING AND MACHINE LEARNING BRANCHES
A method includes the steps of manufacturing and generating data at each of a plurality of locations on the component. Passing the data to a testing branch which performs tests and reaches a conclusion as to whether the component is functionally tolerant at each of the plurality of locations. The data is also passed to a machine learning branch wherein it is compared to training data to determine whether the component is of a functionally tolerant dimension at each of the plurality of locations. The manufactured component accepts should both the testing branch and the machine learning branch determine the component is of functionally tolerant dimensions at the plurality of locations, and rejects the component if either of the testing branch or the machine learning branch determines the component fails to be functionally tolerant dimensions, respectively. A system is also disclosed.
A propulsion system for an aircraft includes an air intake, an engine assembly, a turbocompressor, an electrical assembly, and at least one propulsor. The engine assembly includes an engine. The engine includes an air inlet, an exhaust outlet, and an engine output shaft. The turbocompressor assembly includes a turbocompressor. The turbocompressor includes a turbine and a compressor. The turbine and the compressor form a rotational assembly. The rotational assembly includes a shaft, a bladed turbine rotor of the turbine, and a bladed compressor rotor of the compressor. The shaft interconnects the bladed turbine rotor and the bladed compressor rotor. The turbine includes a turbine inlet and a turbine outlet. The turbine inlet is connected in fluid communication with the exhaust outlet. The compressor includes a compressor inlet and a compressor outlet. The compressor inlet is connected in fluid communication with the air intake. The rotational assembly is mechanically independent of the engine output shaft.
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
B64U 10/14 - Plates-formes volantes comportant quatre axes distincts de rotors, p. ex. quadcoptères
B64U 20/94 - Refroidissement des rotors ou des moteurs de rotor
An aircraft engine, has: a turbine; a scroll case having an inlet connected to a source of combustion gases and an outlet connected to the turbine, and a conduit extending from the inlet to the outlet; a bearing housing including a support flange; an exhaust case downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes extending along a direction having an axial component, the spokes extending through the scroll case and radially supported by the bearing housing, a spoke having a distal end secured to the support flange via: one or more fasteners, and a retaining member at the distal end, the retaining member defining an abutment face facing an axial direction and circumferentially overlapping the support flange, a portion of the support flange located axially between the distal end and the abutment face.
A fir tree structure is provided and includes a part extending radially inwardly from a rotating part and having an inward taper. The part has a spline profile forming fixing lobes and necks, which are interleaved with the fixing lobes. The spline profile has constantly changing radii of curvature characterized in that the radii of curvature are relatively reduced in high stress locations and relatively increased in low stress locations.
A propulsion system for an aircraft includes an engine assembly, an interburner, a turbocompressor, and a flow control assembly. The engine assembly includes an engine. The engine includes an air inlet, an exhaust outlet, and an engine output shaft. The interburner includes an interburner inlet and an interburner outlet. The turbocompressor assembly includes at least one turbocompressor. The at least one turbocompressor includes a turbine and a compressor. The turbine includes a turbine inlet and a turbine outlet. The compressor includes a compressor inlet and a compressor outlet. The flow control assembly includes an engine exhaust diverter subassembly and at least one compressor outlet diverter subassembly. The engine exhaust diverter subassembly is connected to exhaust outlet, the interburner inlet, and the turbine inlet. The at least one compressor outlet diverter subassembly is connected to the compressor outlet, the air inlet, and the interburner inlet.
An aircraft power plant, has: a thermal engine having a combustion chamber inside a housing defining a coolant passage, and an oil passage; and a heat management system including: an oil-to-coolant heat exchanger having an oil conduit in fluid communication with the oil passage and a first coolant conduit in heat exchange relationship with the oil conduit and in fluid communication with the coolant passage; and a coolant-to-air heat exchanger having a second coolant conduit in fluid communication with the coolant passage and an air conduit in heat exchange relationship with the second coolant conduit and in fluid communication with an environment outside the aircraft power plant, wherein the coolant passage inlet and the coolant passage outlet of the coolant passage of the housing are fluidly connected to one another via the first coolant conduit of the oil-to-coolant heat exchanger and the second coolant conduit of the coolant-to-air heat exchanger.
F02B 55/10 - Réfrigération de ces organes et carcasses
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
F28D 21/00 - Appareils échangeurs de chaleur non couverts par l'un des groupes
A casing for a gas turbine engine, including: at least one axial hairpin feature located proximate to either a forward flange or a rearward flange; and a central connecting portion extending from the at least one axial hairpin feature, the at least one axial hairpin feature being located between the central connecting portion and one of the forward flange or a rearward flange.
F01D 21/04 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à une position incorrecte du rotor par rapport au stator, p. ex. indiquant cette position
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
Tooling and methods for assembling a vane ring of a turbine engine are provided. A tool includes a base configured to receive an outer ring and vanes radially inwardly of the outer ring. A heater is operable to apply heat to the outer ring when the outer ring is received on the base. Expandable jaws having a plurality of radially actuatable pads are operable to urge the vanes radially outwardly toward the outer ring. A press is operable to apply an axial force against an inner support of the vane ring to install the inner support radially inwardly of the vanes.
B23P 11/02 - Assemblage ou désassemblage de pièces ou d'objets métalliques par des processus du travail du métal non prévus ailleurs en dilatant une pièce et en l'emmanchant ensuite sur une autre pièce par retrait, ou en rétractant une pièce et en l'emmanchant ensuite par dilatation, p. ex. en employant la pression de fluidesAssemblage ou désassemblage de pièces ou d'objets métalliques par des processus du travail du métal non prévus ailleurs par assemblage à force
B23P 19/027 - Machines effectuant simplement l'assemblage ou la séparation de pièces ou d'objets métalliques entre eux ou des pièces métalliques avec des pièces non métalliques, que cela entraîne ou non une certaine déformationOutils ou dispositifs à cet effet dans la mesure où ils ne sont pas prévus dans d'autres classes pour le montage d'objets à la presse, ou pour le démontage de ces objets utilisant des moyens hydrauliques ou pneumatiques
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
A seal arrangement of a gas turbine engine includes a first component and a second component abutting the first component. An O-ring is positioned in a gland volume defined between the first component and the second component. The O-ring has a hollow circular cross-sectional shape including a ring inner diameter and a ring outer diameter, with a hollow portion defined by the ring inner diameter. The ring inner diameter of the cross-sectional shape is defined as a function of a gland outer diameter, a gland inner diameter and a gland width.
Systems and methods for detecting a performance degradation of a pneumatic bleed-off valve installed in a compressor section of a gas turbine engine are provided. A method includes acquiring a first relationship between a pressure ratio across a compressor of the engine and a rotational speed of a spool of the engine during an acceleration of the spool, and acquiring a second relationship between the pressure ratio across the compressor and the rotational speed of the spool during the deceleration of the spool. A cross-over rotational speed corresponding to an intersection of the first relationship and the second relationship is determined. The performance degradation of the pneumatic bleed-off valve is detected when the cross-over rotational speed is outside a prescribed range.
F04D 27/00 - Commande, p. ex. régulation, des pompes, des installations ou des systèmes de pompage spécialement adaptés aux fluides compressibles
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
An aircraft engine has a flammable fluid containment assembly including a component containing a flammable fluid. A transfer tube is fluidly connected to the component. A sealing ring is provided at an interface between the aircraft engine component and the transfer tube. A first thermal barrier is provided between the sealing ring and a first one of the aircraft engine component and the flammable fluid inside the transfer tube. The first thermal barrier includes a first annular cavity filled with a first thermal insulation medium and axially spanning the sealing ring.
An impeller is provided for a centrifugal compressor of an aircraft engine. The impeller includes an impeller hub and an impeller shroud having an inducer portion, an exducer portion and a bend portion interposed between the inducer portion and the exducer portion. The impeller shroud is swept forwardly in the exducer portion, the impeller shroud and the impeller hub define a gap extending through the inducer portion, the bend portion and the exducer portion and the gap is characterized as having a passage area that exhibits about a 34.375% increase from 10% of the shroud chord to 45% of the shroud chord and exhibits about a −31.1628% decrease from 45% of the shroud chord to 92.5% of the shroud chord.
An assembly is provided for an aircraft propulsion system. This assembly includes an electric machine, an exhaust duct and a mounting structure. The electric machine extends axially along an axis. The exhaust duct includes an annular first section and a non-annular second section fluidly coupled with and downstream of the annular first section. The annular first section extends circumferentially around the axis. The non-annular second section axially overlaps the electric machine and extends partially circumferentially about the electric machine. The mounting structure connects the electric machine to the exhaust duct. The mounting structure includes a duct mount, a machine mount and a framework. The duct mount is connected to the annular first section. The machine mount is connected to the electric machine. The framework extends axially between and is connected to the duct mount and the machine mount. The framework includes a plurality of struts arranged in a truss.
B64D 27/40 - Aménagements pour le montage de groupes moteurs sur aéronefs
B64D 35/022 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride
An aerospace actuator is provided and includes an actuator housing, second and third parts respectively including retraction-limiting and extension-limiting surfaces, a piston and a stroke limiting device. The piston is movable within the actuator housing between a retraction-limited position where the piston impinges against the retraction-limiting surface and an extension-limited position where the piston impinges against the extension-limiting surface. The stroke limiting device includes at least one of a first adjustable engagement unit and a second adjustable engagement unit. The first adjustable engagement unit is disposed between the actuator housing and the second part for adjusting a first distance between the actuator housing and the retraction-limiting surface. The second adjustable engagement unit is disposed between the second part and the third part for adjusting a second distance between the extension-limiting surface and the retraction-limiting surface.
F15B 15/24 - Autres parties constitutives pour limiter la course
F15B 15/14 - Dispositifs actionnés par fluides pour déplacer un organe d'une position à une autreTransmission associée à ces dispositifs caractérisés par la structure de l'ensemble moteur le moteur étant du type à cylindre droit
97.
Centrifugal compressor impeller and method of producing the same
A method of producing a centrifugal compressor impeller is provided. The impeller includes a hub with a rotational axis, and forward and aft surfaces. The aft surface extends between an outer radial surface of the impeller and a bore centered on the rotational axis. The method includes: analyzing the impeller to determine a stress field having tangential stress data values and radial stress data values as a function of radial position within the impeller hub and as a function of axial position within the impeller hub; determining biaxiality ratio (BR) data using the determined tangential stress data values and the determined radial stress data values; determining a peak von Mises creep strain location on the impeller; contouring the aft surface at the peak von Mises creep strain location using the radial stress data values and the BR data; and producing the impeller with the contoured aft surface.
A method of operating a hybrid electric powertrain (HEP) may include determining, based on a HEP torque demand and a maximum torque limit for a gas turbine engine having a mechanical output coupled to a power shaft of the HEP, a target torque for the gas turbine engine. The method may also include controlling the gas turbine engine to operate according to a target speed or the target torque. The method may also include, while operating the gas turbine engine according to the target speed or the target torque, determining, based on the HEP torque demand and the target torque, a total torque compensation demand for an electric machine having a mechanical output controllably couplable to the power shaft of the HEP, and operating the electric machine according to the total torque compensation demand.
B64D 31/09 - Dispositifs amorçant la mise en œuvre actionnés automatiquement en réponse à une défaillance des groupes moteurs
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
99.
BLEED-OFF ASSEMBLY INTAKE DEVICE FOR AN AIRCRAFT PROPULSION SYSTEM
An aircraft propulsion system includes a casing, a compressor, and an assembly. The casing forms an annular cavity. The compressor is configured to direct a compressed air flow into and through the annular cavity in an air flow direction. The assembly includes an intake device disposed on the casing. The intake device includes a particle separator body and an air conduit. The particle separator body extends from an upstream end to a downstream end in the air flow direction. The particle separator body forms an interior surface and an exterior surface. The interior surface forms an inlet cavity of the intake device. The inlet cavity has an inlet opening at the downstream end. The exterior surface forms an airfoil of the particle separator body. The air conduit extends through the casing. The air conduit includes a conduit inlet within the particle separator body. The conduit inlet is disposed at the inlet cavity.
F02C 7/052 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes comportant des dispositifs séparateurs de poussière
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
An aircraft power plant, comprising: a fan for providing power to an aircraft; a gas turbine engine located on a first axial side of the fan, the gas turbine engine having a compressor drivingly engaged by a turbine via an engine shaft; an electric motor powered by a power source, the electric motor driving a motor shaft; and a gearbox located on a second axial side of the fan opposite the first axial side, the gearbox in driving engagement with the engine shaft, the motor shaft, and the fan, the gearbox defining a first load path between the gas turbine engine and the fan and a second load path between the electric motor to the fan, the fan in continuous driving engagement with both of the gas turbine engine and the electric motor along the first load path and the second load path.
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 35/022 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride
