A propeller blade angle control circuit for a turboprop engine includes a propeller control unit controlling a supply of oil to modify an angle of propeller blades, a pump located upstream of the propeller control unit and providing the supply of oil from an engine oil return system to the propeller control unit, and a flow regulator between the pump and the propeller control unit, the flow regulator modulating a supply of oil to the propeller control unit. A bypass, downstream of the pump in the propeller blade angle control circuit, has an inlet fluidly coupled to the pump. The bypass is operable between a closed position and an open position in which a portion of the oil supplied to the propeller control unit is diverted away from the propeller blade angle control circuit. The open position is engaged when an oil pressure reaches a predetermined threshold.
B64C 11/38 - Mécanismes de changement de pas des pales par fluide, p.ex. hydrauliques
B64C 11/30 - Mécanismes de changement de pas des pales
F15B 13/02 - Dispositifs de distribution ou d'alimentation du fluide caractérisés par leur adaptation à la commande de servomoteurs
F15B 20/00 - Dispositions propres à la sécurité pour systèmes de manœuvre utilisant les fluides; Utilisation des dispositifs de sécurité dans les systèmes de manœuvre utilisant des fluides; Mesures d'urgence pour les systèmes de manœuvre utilisant des fluides
F15B 21/08 - Systèmes de servomoteur comportant des moyens de commande actionnés électriquement
2.
METHOD OF MITIGATING CORROSION AND EROSION IN AN AIRCRAFT ENGINE
A method of mitigating corrosion and erosion in an aircraft engine, includes: receiving a concentration of contaminants contained within a sample of an environmental medium ingested by the aircraft engine; determining a frequency of corrosion and erosion mitigation actions based on the concentration of the contaminants; and instructing a performance of the corrosion and erosion mitigation actions at the frequency.
A rotor assembly has: blades having airfoils and roots protruding from platform segments; a rotor disc having a peripheral face defining recesses, and slots, a recess located between two adjacent ones of the slots and bounded by a step; feather seals located radially between the peripheral face and the platform segments, a feather seal having a core extending from a trailing end to a leading end and overlapping a gap defined between two platform segments and tabs protruding from the core, the tabs including: trailing tabs positioned axially outside the recess; and leading tabs, a leading tab extending from a root to a tip and having one or more of: the tip axially positioned outside of the recess; and a fillet at an intersection between the tip and an edge of the leading tab, the edge extending between the tip and the core, and facing the step.
A hybrid-electric powerplant (HEP) of an aircraft comprises a thermal engine providing a first torque input to the HEP and an electric motor providing a second torque input to the HEP. A power management system connected to one or both of the thermal engine and the electric motor comprises an engine control unit (ECU) connected to the thermal engine. The ECU controls fuel supplied to the thermal engine. An electric propulsion control (EPC) is connected to the electric motor and controls power supplied to the electric motor. The EPC includes an EPC protection module in communication with a power source for the electric motor. The EPC protection module disables power supplied to the electric motor upon receipt of a signal indicative of one or more of an over-speed condition and an over-torque condition detected in the HEP.
B64D 31/00 - Commande des groupes moteurs; Leur disposition
B60K 6/00 - Agencement ou montage de plusieurs moteurs primaires différents pour une propulsion réciproque ou commune, p.ex. systèmes de propulsion hybrides comportant des moteurs électriques et des moteurs à combustion interne
B64D 27/00 - Disposition du montage des groupes moteurs sur aéronefs; Aéronefs ainsi caractérisés
5.
SYSTEMS AND METHODS FOR IDENTIFYING A CONDITION OF GAS TURBINE ENGINE SEALS
An assembly for an aircraft propulsion system includes a case assembly, at least one seal, a first pressure sensor, and a computing system. The case assembly forms a cavity. The at least one seal is disposed on the case assembly. The at least one seal is configured to seal the cavity. The first pressure sensor is in fluid communication with the cavity. The first pressure sensor is configured to measure a first pressure within the cavity. The computing system is in signal communication with the first pressure sensor. The computing system includes a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to compare the first pressure to a pressure threshold value to identify a wear condition of the at least one seal.
A stator vane for a gas turbine stator vane stage is provided that includes an airfoil having leading and trailing edges, a vane tip, suction and pressure side surfaces, and at least one aero passage. The leading and trailing edges are chordwise spaced apart. The vane tip is spanwise spaced apart from a radial base end. The suction side surface extends chordwise between the leading and trailing edges, and extends spanwise between the radial base end and the vane tip. The pressure side surface extends chordwise between the leading and trailing edges, and extends spanwise between the radial base end and the vane tip. The at least one aero passage extends through the airfoil between the suction and pressure side surfaces, and is disposed proximate and spanwise separated from the vane tip. The stator vane is configured to be cantilevered with the vane tip being unsupported.
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes ; Entrées d'air pour ensembles fonctionnels de propulsion par réaction
A seal assembly for an aircraft engine includes a first seal having an upstream end exposed to a first pressure area, and a downstream end exposed to a second pressure area, the first seal at least partially defining an intermediate pressure area and a chamber being fluidly connected to the second pressure area and to the intermediate pressure area through passages defined in the first seal, the chamber allowing for a mixing of a first portion of a stream of air with air from the second pressure area, and a second seal connected to the first seal, the second seal biasing the first portion of the stream of air toward the passages, and allowing a second portion of the stream of air from flowing therethrough toward the second pressure area. A method of flowing air through an aircraft seal assembly is also described.
F02C 7/28 - Agencement des dispositifs d'étanchéité
F16J 15/3232 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre par joints élastiques, p.ex. joints toriques avec au moins une lèvre ayant plusieurs lèvres
F16J 15/3284 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre par joints élastiques, p.ex. joints toriques caractérisés par leur structure; Emploi des matériaux
F02F 11/00 - Aménagements des garnitures d'étanchéité dans les moteurs à combustion
A fastening system for an aircraft includes first and second parts of the aircraft, a bolt hole including a bolt countersink and being defined by a bolt hole surface of the first part, a nut hole including a nut countersink and being defined by a nut hole surface of the second part, a bolt having a shank including threads, and a bolt head with an undersurface complementarily shaped to the bolt countersink, and a nut having threads and having a nut chamfer complementarily shaped to the nut countersink, the shank being dimensioned relative to the first and second parts, the bolt hole and the nut hole such that the shank is spaced from one of or both of the bolt hole surface and the nut hole surface. A method of fastening a first aircraft part with a second aircraft part using a bolt and a nut is also described.
F16B 31/06 - Assemblages à vis spécialement modifiés en vue de résister à une charge de traction; Boulons de rupture eu égard aux possibilités de rupture par fatigue
F16B 33/00 - Caractéristiques communes aux boulons et aux écrous
F16D 1/033 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour liaison bout à bout de deux arbres ou de deux pièces analogues par serrage de deux surfaces perpendiculaires à l'axe de rotation, p.ex. avec des brides boulonnées
F16D 1/076 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour montage d'un organe sur un arbre ou à l'extrémité d'un arbre par serrage de deux surfaces perpendiculaires à l'axe de rotation, p.ex. avec des brides boulonnées
9.
ASSEMBLIES AND METHODS FOR CONTROLLING LUBRICATION FOR ROTARY ENGINE APEX SEALS
An assembly includes a rotor housing, a first rotor, a lubrication system, a first vibration sensor, and an engine control system. The rotor housing forms a first rotor cavity. The first rotor is configured for rotation within the first rotor cavity. The first rotor includes the plurality of apex seals. The lubrication system is configured to supply a lubrication flow for lubrication of the plurality of apex seals. The first vibration sensor is on the rotor housing. The first vibration sensor is configured to generate a vibration measurement signal. The engine control system includes a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: identify that the vibration measurement signal exceeds a first vibration threshold, and increase a flow rate of the lubrication flow based on an identification of the vibration measurement signal exceeding the first vibration threshold.
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
A turbine exhaust case (TEC) for a gas turbine engine, has: an inner case extending circumferentially about a central axis; an outer case disposed radially outward from the inner case and extending circumferentially about the central axis; struts extending between the inner case and the outer case, a strut of the struts having an airfoil extending from an inner end to an outer end along a span and from a leading edge to a trailing edge along a chord, the airfoil being cambered and having a pressure side being concave and a suction side being convex, and a slot defined through the airfoil downstream of the leading edge, the slot extending from a slot inlet on the suction side to a slot outlet on the pressure side, the slot defining a fluid flow passage for directing fluid flow from the suction side to the pressure side through the airfoil.
B64D 27/12 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz à l'intérieur de la voilure ou fixés à celle-ci
F16M 1/04 - Châssis, carters ou carcasses pour moteurs, machines ou appareils; Châssis servant de bâtis de machines pour moteurs rotatifs ou machines similaires
11.
EXHAUST NOZZLE ASSEMBLY FOR AN AIRCRAFT PROPULSION SYSTEM
An exhaust nozzle assembly for a propulsion system include a primary nozzle, an outer shroud, an ejector nozzle, and an actuator. The primary nozzle extends along an exhaust centerline. The primary nozzle includes a downstream axial end. The outer shroud surrounds the primary nozzle. The ejector nozzle extends axially between a first axial end and a second axial end. The second axial end forms a nozzle exit plane for the exhaust nozzle assembly. The ejector nozzle converges in a direction from the first axial end to the second axial end. The ejector nozzle forms a mixing cross-sectional area between the primary nozzle and the ejector nozzle at the downstream axial end. The actuator is mounted on the ejector nozzle. The actuator is configured to move the ejector nozzle between a first position and a second position, relative to the outer shroud, to control an area of the mixing cross-sectional area.
F02K 1/06 - Variation de la section utile de la tubulure de jet ou de la tuyère
F01N 13/00 - Silencieux ou dispositifs d'échappement caractérisés par les aspects de structure
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
F01N 3/00 - Silencieux ou dispositifs d'échappement comportant des moyens pour purifier, rendre inoffensifs ou traiter les gaz d'échappement
F02K 1/30 - Ensembles fonctionnels caractérisés par la forme ou la disposition de la tubulure de jet ou de la tuyère; Tubulures de jet ou tuyères particulières à cet effet utilisant des jets de fluide pour influencer l'écoulement du jet pour faire varier la section utile de la tubulure de jet, ou de la tuyère
F02K 1/40 - Tuyères comportant des moyens pour diviser le jet en plusieurs jets partiels ou possédant une section de sortie allongée
F02K 1/78 - Autres structures des tubulures de jet
12.
SYSTEMS AND METHODS FOR DETERMINING GAS TURBINE ENGINE TEMPERATURES
A system for determining an indicated turbine temperature (ITT) for a gas turbine engine includes an engine control system. The engine control system includes a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: determine a first estimated outlet temperature value for a high-pressure turbine of the gas turbine engine, determine an estimated work (WHPT) of the high-pressure turbine, determine an estimated inlet temperature value for the high-pressure turbine using the estimated work (WHPT), and determine the ITT by calculating a second estimated outlet temperature value using the estimated inlet temperature value, the second estimated outlet temperature value different than the first estimated outlet temperature value.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p.ex. à la température, à la pression, à la vitesse du rotor
An aircraft engine, has: a pressure probe having: a static member having a front face and a back face, an inlet and an outlet fluidly connected to the inlet, the front face defining a curved surface; a movable member movably engaged to the static member and movable relative to the static member about a center of rotation, the movable member having a central axis, the movable member having an engagement section matingly engaged to the front face to slide against the curved surface, the engagement section having an opening, and an orientation section protruding from the engagement section and located rearward of the center of rotation, the orientation section defining an external surface exposed to the flow, wherein the movable member is movable relative to the static member as a result of a force imparted by the flow on the external surface.
A propeller shaft assembly for an aircraft engine includes a shaft having a bore extending through the shaft at a front end thereof, the front end of the shaft having an outer surface facing radially outwardly from the shaft and an inner surface spaced apart from the outer surface and facing radially inwardly to the bore. The shaft has a front flange extending radially outwardly on the outer surface, the front flange having a base merging with the outer surface of the shaft. A sleeve is coupled to the shaft within the bore by an interference fit between the sleeve and the shaft, at least part of the sleeve axially aligned with the front flange. The sleeve axially extends from a front to a rear sleeve end, the rear sleeve end axially offset from the engine side surface of the front flange at the base of the front flange.
B64D 35/00 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions
F16D 1/033 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour liaison bout à bout de deux arbres ou de deux pièces analogues par serrage de deux surfaces perpendiculaires à l'axe de rotation, p.ex. avec des brides boulonnées
A propeller shaft assembly for an aircraft engine includes a shaft having: an annular wall extending circumferentially about a shaft axis and circumscribing a hollowed interior defining a cavity in a front end portion of the shaft, the annular wall having an outer surface and an inner surface facing radially inwardly to the cavity; and a front flange projecting radially outwardly from the annular wall. The front flange includes a hub side surface defining an interface plane and adapted to abut with a propeller hub. The shaft also includes a reinforcement web defining an end wall of the cavity, the reinforcement web extending radially inwardly from the inner surface of the annular wall. At least part of the reinforcement web is radially aligned with the front flange. At least one perforation extends axially through the reinforcement web.
B63H 23/35 - Freinage ou verrouillage des arbres, c. à d. moyens pour ralentir ou arrêter la rotation des arbres porte-hélices ou pour les empêcher de commencer à tourner
B64D 35/00 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions
F16D 1/076 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour montage d'un organe sur un arbre ou à l'extrémité d'un arbre par serrage de deux surfaces perpendiculaires à l'axe de rotation, p.ex. avec des brides boulonnées
A propeller shaft assembly for an aircraft engine includes a shaft having a bore extending through the shaft at a front end thereof, the front end of the shaft having an outer surface facing radially outwardly from the shaft and an inner surface spaced apart from the outer surface and facing radially inwardly to the bore. A front flange extends radially outwardly on the outer surface, the front flange defining a hub side surface adapted to abut with a propeller hub. A reinforcement rib extends radially inwardly towards a central axis of the shaft. At least part of the reinforcement rib is radially aligned with the front flange.
B64D 35/00 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions
F16D 1/033 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour liaison bout à bout de deux arbres ou de deux pièces analogues par serrage de deux surfaces perpendiculaires à l'axe de rotation, p.ex. avec des brides boulonnées
F16D 1/076 - Accouplements pour établir une liaison rigide entre deux arbres coaxiaux ou d'autres éléments mobiles d'une machine pour montage d'un organe sur un arbre ou à l'extrémité d'un arbre par serrage de deux surfaces perpendiculaires à l'axe de rotation, p.ex. avec des brides boulonnées
A fuel nozzle for a turbine engine, comprising: a flange defining at least one flange passage; a tip spaced from the flange, the tip defining at least one tip passage; a stem having a first stem end fixedly joined to the flange and a second stem end fixedly joined to the tip, the stem having a peripheral wall extending lengthwise between the first stem end and the second stem end and peripherally around a stem chamber, the tip sealing the stem chamber at the second stem end; and at least one fuel line extending at least partially inside the stem chamber and having a first line end fluidly connected to the at least one flange passage and a second line end fluidly connected to the at least one tip passage.
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
F02C 7/22 - Systèmes d'alimentation en combustible
F23R 3/02 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par la configuration du flux d'air ou du flux de gaz
18.
ADAPTIVELY DEPOSITING BRAZE MATERIAL(S) USING CT SCAN DATA
A method is disclosed for providing a component. During this method, a substrate is scanned using computed tomography to provide substrate scan data. The substrate scan data is compared to substrate reference data to provide additive manufacturing data. Braze powder is deposited with the substrate based on the additive manufacturing data. 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.
A method is disclosed during which a substrate is provided. Braze powder is deposited with the substrate using an additive manufacturing device. The braze powder is sintered together and to the substrate during the depositing of the braze powder to provide the substrate with sintered braze material. The substrate and the sintered braze material are heated to melt the sintered braze material and diffusion bond the sintered braze material to the substrate.
B22F 3/00 - Fabrication de pièces ou d'objets à partir de poudres métalliques, caractérisée par le mode de compactage ou de frittage; Appareils spécialement adaptés à cet effet
B22F 7/00 - Fabrication de couches composites, de pièces ou d'objets à base de poudres métalliques, par frittage avec ou sans compactage
A method is disclosed for providing a component. During this method, a first object is additive manufactured. The first object is scanned using computed tomography to provide first object scan data. 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.
B29C 64/188 - Procédés de fabrication additive impliquant des opérations supplémentaires effectuées sur les couches ajoutées, p.ex. lissage, meulage ou contrôle d’épaisseur
B33Y 50/02 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
B29C 64/386 - Acquisition ou traitement de données pour la fabrication additive
B22F 12/00 - Appareils ou dispositifs spécialement adaptés à la fabrication additive; Moyens auxiliaires pour la fabrication additive; Combinaisons d’appareils ou de dispositifs pour la fabrication additive avec d’autres appareils ou dispositifs de traitement ou de fabrication
A method is disclosed for providing a component. During this method, a substrate is scanned using structured light to provide substrate scan data. The substrate scan data is compared to substrate reference data to provide additive manufacturing data. Braze powder is deposited with the substrate based on the additive manufacturing data. 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.
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.
B22F 7/06 - 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
A deaeration rotor for an aircraft engine lubrication system comprising: an internal ring about an axis having a radially outer internal ring surface defining an inner boundary of an inner passage of the deaeration rotor; an external ring about the axis having a radially inner external ring surface defining an outer boundary of an outer passage of the deaeration rotor; a disc about the axis radially between the internal ring and the external ring, the disc having a radially inner disc surface defining an outer boundary of the inner passage and a radially outer disc surface defining an inner boundary of the outer passage; and blades circumferentially spaced from one another relative to the axis extending in the outer passage from at least one of the external ring and the disc, the blades located radially inward of an annular portion of the outer passage immediately downstream of the blades.
F01M 11/08 - Séparation du lubrifiant de l'air ou du mélange air-carburant avant introduction dans le cylindre
B64D 33/00 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs
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
A deaeration system for an engine lubrication system, the deaeration system comprising: a deaeration rotor rotatable about an axis and including: a rotor inlet extending circumferentially around the axis, a first and a second rotor outlet, a first rotor passage in fluid communication between the rotor inlet and the first rotor outlet, and a second rotor passage in fluid communication between the rotor inlet and the second rotor outlet in parallel to the first rotor passage; and a deaeration conduit including: a conduit inlet, a splitter downstream of the conduit inlet relative to a flow of lubricant through the deaeration conduit, a first conduit outlet and a second conduit outlet downstream of the splitter, the first conduit outlet in fluid communication with the rotor inlet, the conduit inlet having a curved portion extending away from the splitter.
A method of overhaul is provided. During this overhaul method, a substrate is scanned using structured light to provide substrate scan data. The substrate is from a component previously installed within an engine. The substrate scan data is compared to substrate reference data to provide additive manufacturing data. Material is deposited with the substrate using an additive manufacturing device based on the substrate scan data to provide a first object. The first object is scanned using the structured light to provide first object scan data. 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.
B23P 6/00 - Remise en état ou réparation des objets
B23K 28/00 - Soudage ou découpage non couvert par l'un des groupes
B23P 9/00 - Traitement ou finition mécanique des surfaces, avec ou sans calibrage, dans le but primordial de mieux résister à l'usure ou aux chocs, p.ex. traitement des aubes de turbines ou des paliers pour les rendre lisses ou rugueux; Caractéristiques, non prévues ailleurs, de telles surfaces lorsque leur traitement n'est pas précisé
26.
ADAPTIVE MANUFACTURING USING STRUCTURED LIGHT DATA
A method is disclosed for providing a component. During this method, braze powder is additively 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 structured light 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.
B22F 7/06 - 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
B29C 64/386 - Acquisition ou traitement de données pour la fabrication additive
B22F 3/105 - Frittage seul en utilisant un courant électrique, un rayonnement laser ou un plasma
B22F 3/24 - Traitement ultérieur des pièces ou objets
B23K 1/20 - Traitement préalable des pièces ou des surfaces destinées à être brasées, p.ex. en vue d'un revêtement galvanique
B23K 26/03 - Observation, p.ex. surveillance de la pièce à travailler
B23K 35/02 - Baguettes, électrodes, matériaux ou environnements utilisés pour le brasage, le soudage ou le découpage caractérisés par des propriétés mécaniques, p.ex. par la forme
A method is provided during which first braze powder is deposited with a substrate. The first braze powder is sintered to the substrate during the depositing of the first braze powder to provide the substrate with sintered first braze material. Second braze powder is deposited with the substrate. The second braze powder is different than the first braze powder. The second braze powder is sintered to the substrate during the depositing of the second braze powder to provide the substrate with sintered second braze material. The sintered first braze material and the sintered second braze material are heated to melt the sintered first braze material and the sintered second braze material and to diffusion bond the sintered first braze material and the sintered second braze material to the substrate.
B22F 7/06 - 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
A system and method for controlling an aircraft engine is provided. The method includes a) producing first sensor data using a first sensor sensing a first parameter during operation of the aircraft engine on a flight mission; b) producing other sensor data using a plurality of second sensors sensing a plurality of other parameters, during operation of the aircraft engine; c) providing the first and other sensor data to a control unit during operation of the aircraft engine; d) storing the first and other sensor data during operation of the aircraft engine; e) using an artificial intelligence (AI) model that is trained using the stored first and other sensor data produced during operation of the aircraft engine, to produce one or more derived first parameter values; and f) selectively providing the one or more derived first parameter values to the control unit for use in controlling the aircraft engine.
B64D 31/00 - Commande des groupes moteurs; Leur disposition
G06N 3/004 - Vie artificielle, c. à d. agencements informatiques simulant la vie
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gaz; Commande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
A system is provided for an aircraft. This aircraft system includes a propulsion system, and the propulsion system includes a first thermal engine, a second thermal engine and a first electric machine. The propulsion system is configured to operate the first thermal engine and the second thermal engine, without operating the first electric machine, during a first mode of operation to provide aircraft thrust. The propulsion system is configured to operate the first electric machine and the second thermal engine, without operating the first thermal engine, during a second mode of operation to provide the aircraft thrust.
Health monitoring systems and associated methods for gas turbine engines are provided. A health monitoring method includes using a microphone to acquire operation data indicative of acoustic energy generated in a core gas path of the gas turbine engine. The operation data is compared to reference data indicative of an acoustic signature of fluid noise associated with a non-normal condition in the core gas path of the gas turbine engine. Based on the comparing of the operation data to the reference data, the non-normal condition is determined to exist within the core gas path of the gas turbine engine. A signal indicative of the existence of the non-normal condition within the core gas path of the gas turbine engine is output.
31.
GAS TURBINE ENGINE COMPONENT WITH COPPER OXIDE COATING
THE ROYAL INSTITUTION FOR THE ADVANCEMENT OF LEARNING/MCGILL UNIVERSITY (Canada)
Inventeur(s)
Larose, Joel
Roy, Amit
Sharifi, Navid
Stoyanov, Pantcho
Moreau, Christian
Chromik, Richard
Makowiec, Mary
Abrégé
A formation method is provided. During this formation method, a metallic substrate is provided. A coating is deposited onto the metallic substrate using a suspension plasma spray process. The coating is formed from or otherwise includes copper oxide.
A system for a gas turbine engine includes an engine control system. The engine control system includes a processor and a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: obtain a current engine installation configuration for the gas turbine engine and the aircraft, determine an expected normalized value of the engine parameter based on the current engine installation configuration and one or more of a normalized engine power (SHPN) of the gas turbine engine, an engine inlet temperature, an airspeed, or an altitude, determine a fully deteriorated engine (FDE) value of the engine parameter using the expected normalized value of the engine parameter, determine a current value of the engine parameter for the gas turbine engine, and determine the engine operating margin for the engine parameter based on the FDE value of the engine parameter and the current value of the engine parameter.
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gaz; Commande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
A system for a gas turbine engine includes an engine control system. The engine control system includes a processor in communication with a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: obtain a current engine installation configuration for the gas turbine engine, determine a normalized value of the engine parameter for an uninstalled gas turbine engine based on the current engine installation configuration and one or more of a normalized engine power (SHPN) of the gas turbine engine, an airspeed, or an altitude, determine a fully deteriorated engine (FDE) value of the engine parameter using the normalized value of the engine parameter, determine a current value of the engine parameter for the gas turbine engine, and determine the engine operating margin for the engine parameter based on the FDE value of the engine parameter and the current value of the engine parameter.
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
B64D 45/00 - Indicateurs ou dispositifs de protection d'aéronefs, non prévus ailleurs
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gaz; Commande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
34.
SIMULTANEOUSLY DISASSEMBLING ROTOR BLADES FROM A GAS TURBINE ENGINE ROTOR DISK
A method is provided for disassembling a rotor of a gas turbine engine. During this method, the rotor is provided which includes a rotor disk and a plurality of rotor blades arranged circumferentially about an axis. The rotor blades include a plurality of airfoils and a plurality of attachments that mount the rotor blades to the rotor disk. Each of the rotor blades includes a respective one of the airfoils and a respective one of the attachments. A press is arranged against the rotor. The press axially engages each of the rotor blades. The press moves axially along the axis to simultaneously push the rotor blades and remove the attachments from a plurality of slots in the rotor disk.
A method is provided for assembling a rotor of a gas turbine engine. During this method, a rotor disk is provided that includes an axis and a plurality of slots arranged circumferentially about the axis in an array. A plurality of rotor blades are provided that include a plurality of airfoils and a plurality of attachments. Each of the rotor blades includes a respective one of the airfoils and a respective one of the attachments. Each of the attachments is inserted partially into a respective one of the slots. The rotor blades are rested on top of a blade support structure. The blade support structure is lowered axially downward along the rotor disk to simultaneously seat the attachments into the slots.
An exhaust assembly for a gas turbine engine includes an outer exhaust case, an inner exhaust case, and a hollow strut. The outer exhaust case forms an outer cavity radially outward of the outer exhaust case. The inner exhaust case is positioned radially inward of the outer exhaust case. The outer exhaust case and the inner exhaust case form a core flow path. The inner exhaust case forms a centerbody. The hollow strut includes a strut body, an inlet, an outlet, and an internal passage. The strut body is connected to the outer exhaust case and the inner exhaust case. The internal passage extending through the strut body from the inlet to the outlet. The inlet is located at the outer radial end. The inlet is in fluid communication with the outer cavity. The internal passage is configured to direct gas from the outer cavity to the outlet.
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
F01D 25/30 - Têtes d'évacuation, chambres ou parties analogues
F01K 1/04 - Accumulateurs de vapeur avec emmagasinage de la vapeur dans un liquide, p.ex. accumulateur type Ruth
A fuel assembly for a gas turbine engine includes a fuel supply tube, a fuel port, a fuel manifold, and a fuel manifold adapter. The fuel supply tube is configured to convey a fuel. The fuel port is fluidly coupled to the fuel supply tube and configured to receive the fuel from the fuel supply tube. The fuel manifold includes a fuel inlet and a plurality of fuel outlets. The fuel inlet is fluidly coupled to the fuel port and configured to receive the fuel from the fuel port. The fuel manifold adapter includes a first mount portion and a second mount portion. The first mount portion is connected to the fuel port. The first mount portion is moveable relative to the second mount portion.
A compressor for an aircraft engine, has: a dual-impeller having: a first impeller having a first inlet and a first outlet located radially outwardly of the first inlet, and a second impeller rotatable with the first impeller, the second impeller having a second inlet and a second outlet located radially outwardly of the second inlet, the first inlet and the second inlet facing opposite axial directions; and first conduits having first conduit inlets and first conduit outlets, the first conduit inlets fluidly connected to the first outlet of the first impeller, the first conduit outlets fluidly connected to the second inlet of the second impeller; and second conduits having second conduits inlets fluidly connected to the second outlet of the second impeller, a second conduit of the second conduits disposed circumferentially between two adjacent first conduits of the first conduits.
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
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
39.
GAS TURBINE ENGINE EXHAUST CASE WITH BLADE SHROUD AND STIFFENERS
An assembly is provided for a gas turbine engine. This engine assembly includes a bladed rotor rotatable about an axis, and an engine case. The engine case includes an outer duct wall, a first circumferential stiffener, a second circumferential stiffener and a plurality of axial stiffeners. The outer duct wall forms a shroud around the bladed rotor. The first circumferential stiffener extends circumferentially about the outer duct wall. The second circumferential stiffener extends circumferentially about the outer duct wall. The axial stiffeners are arranged circumferentially about the outer duct wall. Each of the axial stiffeners extends axially between the first circumferential stiffener and the second circumferential stiffener.
F16M 1/04 - Châssis, carters ou carcasses pour moteurs, machines ou appareils; Châssis servant de bâtis de machines pour moteurs rotatifs ou machines similaires
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
A gas turbine engine apparatus includes an engine flowpath, a protuberance and a variable vane. The protuberance projects into the engine flowpath. The variable vane extends across the engine flowpath. The variable vane includes a pivot axis and an airfoil. The variable vane is configured to pivot about the pivot axis between a first position and a second position. The airfoil extends spanwise along a span line between a first end and a second end. The airfoil extends chordwise along a chord line between a leading edge and a trailing edge. A recess extends spanwise into the airfoil from the first end. The airfoil, at the first end, is spaced from the protuberance when the variable vane is in the first position. The airfoil, at the first end, is aligned with the protuberance and the protuberance projects into the recess when the variable vane is in the second position.
F01D 9/02 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage
F01D 1/02 - "Machines" ou machines motrices à déplacement non positif, p.ex. turbines à vapeur avec des moyens stationnaires de guidage de fluide de travail et un rotor à ailettes ou de structure analogue
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
F04D 29/56 - Moyens de guidage du fluide, p.ex. diffuseurs réglables
41.
VARIABLE VANE AIRFOIL WITH AIRFOIL TWIST TO ACCOMMODATE PROTUBERANCE
A gas turbine engine apparatus includes a variable vane. The variable vane includes a pivot axis and an airfoil. The variable vane is configured to pivot about the pivot axis between a first position and a second position. The airfoil extends spanwise along a span line between a first end and a second end. The airfoil extends chordwise along a chord line between a leading edge and a trailing edge. The chord line is angularly offset from a reference plane containing the pivot axis by a twist angle. A first section of the airfoil is disposed at the first end. The twist angle varies as the first section extends spanwise along the span line. A second section of the airfoil is disposed spanwise between the first section and the second end. The twist angle is uniform as the second section extends spanwise along the span line.
F01D 1/02 - "Machines" ou machines motrices à déplacement non positif, p.ex. turbines à vapeur avec des moyens stationnaires de guidage de fluide de travail et un rotor à ailettes ou de structure analogue
F01D 9/02 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage
42.
AIRCRAFT CONTRAIL MONITORING AND TARGETED MITIGATION
A system is provided for an aircraft. This aircraft system includes an aircraft powerplant, a powerplant sensor system, an environment sensor system and a monitoring system. The aircraft powerplant includes a heat engine. The powerplant sensor system is configured to provide engine data indicative of one or more operating parameters of the heat engine. The environment sensor system is configured to provide environment data indicative of one or more environmental parameters of an environment in which the heat engine is operating. The monitoring system is configured to determine formation of a contrail and quantify an impact of the contrail when formed based on the engine data and the environment data.
An aircraft engine, has: an inlet extending circumferentially around a central axis; an annular inlet duct having a duct inlet fluidly connected to an environment outside of the aircraft engine and a duct outlet fluidly connected to the inlet; a flow restrictor extending across the annular inlet duct and being movable within the annular inlet duct; an actuator engaged to the flow restrictor and operable to move the flow restrictor; and a controller operatively connected to at least one sensor and the actuator, the controller having a processing unit and a computer- readable medium operatively connected to the processing unit and containing instructions for: receiving a signal indicative of a pressure difference between opposite sides of the flow restrictor; and powering the actuator to move the flow restrictor with the actuator from a first position to a second position offset form the first position as a function of the pressure difference.
F02C 9/16 - Commande du débit du fluide de travail
F02D 13/00 - Réglage de la puissance du moteur par variation des caractéristiques de fonctionnement de la soupape d'admission ou de la soupape d'échappement, p.ex. réglage de la durée d'admission ou d'échappement
44.
AIRCRAFT INTAKE DUCT WITH PASSIVELY MOVABLE FLOW RESTRICTOR
An aircraft engine, has: an inlet leading to a compressor section, the inlet extending circumferentially around a central axis; an annular inlet duct extending circumferentially around the central axis, the annular inlet duct having a duct inlet fluidly connected to an environment outside of the aircraft engine and a duct outlet fluidly connected to the inlet, the duct outlet extending circumferentially around the central axis; and a flow restrictor located within the annular inlet duct, the flow restrictor extending across the annular inlet duct, being movable within the annular inlet duct along a circumferential direction relative to the central axis in response to a fluid pressure differential on opposed sides of the flow restrictor.
F02C 7/042 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction à géométrie variable
B64D 33/02 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des entrées d'air de combustion
F02C 9/16 - Commande du débit du fluide de travail
45.
POROUS COVER FOR A TAKEOFF PORT OF A GAS TURBINE ENGINE
A system is provided for a gas turbine engine. This engine system includes a flowpath wall, a takeoff conduit and a porous cover. The flowpath wall forms a peripheral boundary of an internal engine flowpath. The flowpath wall includes a takeoff port. The takeoff conduit includes an internal conduit passage fluidly coupled with the internal engine flowpath through the takeoff port. The takeoff conduit projects out from the flowpath wall. The porous cover for the internal conduit passage is disposed at the takeoff port.
46.
VARIABLE GUIDE VANE ASSEMBLY FOR GAS TURBINE ENGINE
A variable guide vane assembly for a gas turbine engine stator is provided. The variable guide vane assembly includes a plurality of vanes and a plurality of RT mechanisms. The vanes extend between a shroud and hub. The vanes are circumferentially disposed and spaced apart from one another. Each vane includes inner and outer radial ends, and inner and outer radial posts. Each vane is pivotally mounted to rotate about its rotational axis. Each RT mechanism is in communication with the inner or outer radial post of a respective vane. The RT mechanism includes a pin connected to the vane that is disposed in a ramp slot non- rotational relative to the pivotable vane. The ramp slot extends between first and second lengthwise ends. Rotation of the vane relative to the ramp slot causes the pin to travel within the ramp slot and the vane to translate linearly.
F01D 17/16 - Organes de commande terminaux disposés sur des parties du stator faisant varier l'aire effective de la section transversale des injecteurs ou tuyères de guidage en obturant les injecteurs
A method and system for processing parameter values from a redundant sensor configured to sense a parameter used in the control of an aircraft engine is provided. The method includes: a) receiving a plurality of parameter values from a redundant sensor by sensing the same parameter at the same time; b) identifying mismatched parameter values; c) producing a predicted parameter value using an artificial intelligence (AI) model having a database of parameter values representative of the sensed parameter; d) providing the predicted parameter value to a control unit; and e) operating the control unit to select a first parameter value or a second parameter value using the predicted parameter for use in the control of the aircraft engine.
A method is provided for operating an aircraft system. During this method, an electric machine of an electrical system is operated onboard an aircraft. A first set of electrical system waves produced by the electrical system is sensed. A set of electrical system attenuation waves is produced to attenuate a second set of electrical system waves produced by the electrical system in response to sensing the first set of the electrical system waves.
G10K 11/178 - Procédés ou dispositifs de protection contre le bruit ou les autres ondes acoustiques ou pour amortir ceux-ci, en général utilisant des effets d'interférence; Masquage du son par régénération électro-acoustique en opposition de phase des ondes acoustiques originales
A system is provided for an aircraft. This aircraft system includes a gas turbine engine and a sensor system. The gas turbine engine includes an inlet and a compressor section. A flowpath projects radially inward into the gas turbine engine from the inlet and extends through the compressor section. The sensor system includes a plurality of static pressure sensors at least partially within the flowpath. The sensor system is configured to determine a total pressure characteristic within the flowpath using the plurality of static pressure sensors.
A method is provided for operating an aircraft system. During this operating method, a plurality of drive units are provided that include a thermal engine drive unit and an electric machine drive unit. A mechanical load is powered using a first of the drive units. The first of the drive units includes a first rotating structure. A parameter of the first rotating structure is monitored. A failure of the first of the drive units is detected based on the monitored parameter. A switch is made from the first of the drive units to a second of the drive units to power the mechanical load where the failure of the first of the drive units is detected.
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
B64D 35/02 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le type de groupe moteur
B64D 35/08 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le fait que la transmission est entraînée par plusieurs groupes moteurs
B64D 31/09 - en réponse à une défaillance des groupes moteurs
B64D 31/18 - pour les groupes moteurs hybrides-électriques
51.
DIFFUSER AND ASSOCIATED COMPRESSOR SECTION OF AIRCRAFT ENGINE
The compressor section can have a centrifugal impeller operable to rotate around an axis, the centrifugal impeller having blades, a compressor inlet oriented towards the front and axially relative the axis, a compressor outlet oriented radially outwardly relative the axis, a diffuser having a diffusion flow path, a diffuser inlet in fluid flow communication with the compressor outlet, a diffusion flow path between a rear wall and a front wall; a collector extending circumferentially around the axis, having a collector inlet in fluid communication with the diffuser outlet, and a collector outlet; and hollow structural members protruding rearwardly from the rear wall, the hollow structural members being circumferentially interspaced from one another, each hollow structural member having a length extending radially along the rear wall and having an internal conduit extending radially inwardly along the length.
F02C 9/16 - Commande du débit du fluide de travail
F01D 1/02 - "Machines" ou machines motrices à déplacement non positif, p.ex. turbines à vapeur avec des moyens stationnaires de guidage de fluide de travail et un rotor à ailettes ou de structure analogue
F01D 9/02 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage
F02C 3/08 - 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 comprenant au moins un étage radial
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes ; Entrées d'air pour ensembles fonctionnels de propulsion par réaction
F02C 9/20 - Commande du débit du fluide de travail par réglage des aubes
F04D 29/44 - Moyens de guidage du fluide, p.ex. diffuseurs
52.
FLOW DEFLECTOR FOR APERTURE IN GAS TURBINE ENGINE FLOWPATH WALL
A system is provided for a gas turbine engine. This gas turbine engine system includes a flowpath wall and a deflector. The flowpath wall includes a surface and an opening to a blind aperture. The surface fomis a peripheral boundary of an internal engine flowpath. The opening is disposed in the surface. The blind aperture extends vertically into the flowpath wall from the opening. The deflector projects vertically out from the flowpath wall into the internal engine flowpath. The deflector is configured to deflect gas flowing within the internal engine flowpath over the opening.
An aircraft system is provided that includes a first propulsor rotor, a second propulsor rotor, a drivetrain and an intermittent combustion engine. The first propulsor rotor is rotatable about a first propulsor axis. The second propulsor rotor is rotatable about a second propulsor axis. The drivetrain includes a drive structure and a transmission. The drive structure is rotatable about a drive axis that is angularly offset from the first propulsor axis and the second propulsor axis. An output of the transmission is coupled to the first propulsor rotor and the second propulsor rotor through the drive structure. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor and the second propulsor rotor through the drivetrain.
A method for developing a numerical control manufacturing program for a common geometric feature of a first component includes obtaining manufacturing process data for the common geometric feature. The manufacturing process data is associated with one or more numerical control manufacturing processes for the common geometric feature of one or more second components. Each of the one or more second components includes the common geometric feature. The method further includes determining one or more manufacturing constraints for the numerical control manufacturing program for the common geometric feature of the first component, selecting a numerical control manufacturing process of the one or more numerical control manufacturing processes, obtaining manufacturing process parameters for the selected one or more numerical control manufacturing processes, and developing the numerical control manufacturing program for the common geometric feature of the first component. The developed numerical control manufacturing program includes the manufacturing process parameters.
G05B 19/4097 - 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'u caractérisée par l'utilisation de données de conception pour commander des machines à commande numérique [CN], p.ex. conception et fabrication assistées par ordinateur CFAO
55.
AIRCRAFT PROPULSION SYSTEM WITH INTERMITTENT COMBUSTION ENGINE(S)
An aircraft system is provided that includes an aircraft fuselage, a first propulsor, a first drivetrain, a second propulsor, a second drivetrain and an intermittent combustion engine. The first propulsor is outside of the aircraft fuselage. The first propulsor includes a first propulsor rotor and a first vane array. The first drivetrain is coupled to the first propulsor rotor. The second propulsor is outside of the aircraft fuselage. The second propulsor includes a second propulsor rotor and a second vane array. The second drivetrain is coupled to the second propulsor rotor. The intermittent combustion engine is within the aircraft fuselage. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor, independent of the second propulsor rotor, through the first drivetrain. The intermittent combustion engine is configured to drive rotation of the second propulsor rotor, independent of the first propulsor rotor, through the second drivetrain.
B64D 35/08 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le fait que la transmission est entraînée par plusieurs groupes moteurs
B64D 27/00 - Disposition du montage des groupes moteurs sur aéronefs; Aéronefs ainsi caractérisés
56.
AIRCRAFT PROPULSION SYSTEM WITH INTERMITTENT COMBUSTION ENGINE(S)
An aircraft system is provided that includes a first propulsor, a second propulsor, a drivetrain and an intermittent combustion engine. The first propulsor includes a first propulsor rotor and a first vane array. The second propulsor includes a second propulsor rotor and a second vane array. The drivetrain includes a drive structure and a transmission. An output of the transmission is coupled to the first propulsor rotor and the second propulsor rotor through the drive structure. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor and the second propulsor rotor through the drivetrain.
B64D 35/04 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le fait que la transmission entraîne plusieurs hélices ou rotors
B64D 27/00 - Disposition du montage des groupes moteurs sur aéronefs; Aéronefs ainsi caractérisés
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
57.
AIRCRAFT PROPULSION SYSTEM WITH INTERMITTENT COMBUSTION ENGINE(S)
An aircraft system is provided that includes a first propulsor rotor, a first transmission, a second propulsor rotor, a second transmission and an intermittent combustion engine. The first propulsor rotor is rotatable about a first propulsor axis. The first transmission is coupled to the first propulsor rotor. The second propulsor rotor is rotatable about a second propulsor axis. The second transmission is coupled to the second propulsor rotor. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor through the first transmission. The intermittent combustion engine is configured to drive rotation of the second propulsor rotor through the second transmission.
B64D 35/04 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le fait que la transmission entraîne plusieurs hélices ou rotors
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
Aircraft power plants including hydrogen turbo-expanders, and associated methods are provided. One method of operating an aircraft power plant includes: driving a load onboard an aircraft with a combustion engine; heating a solid metal hydride onboard the aircraft to cause hydrogen gas to be released from the solid metal hydride; expanding the hydrogen gas through a turbo-expander to produce work; and using the work produced by the turbo-expander to drive the load.
A service tube assembly for an aircraft engine, comprising: a service tube having a threaded end portion, an opposed end portion and an annular tube surface proximate to the threaded end portion; a housing having an outer surface defining a tube socket extending in the outer surface, and a ramp extending toward the tube socket so as to define an engagement direction, the tube socket engaged with the threaded end portion of the service tube; a locking member having a bottom surface disposed against the ramp and an engagement surface facing toward the service tube, the locking member slidable along the ramp in the engagement direction between a first member position in which the engagement surface is spaced from the annular tube surface and a second member position in which the engagement surface contacts the annular tube surface; and a fastener releasably holding the locking member against the ramp.
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
B64D 33/00 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs
F01D 25/00 - "MACHINES" OU MACHINES MOTRICES À DÉPLACEMENT NON POSITIF, p.ex. TURBINES À VAPEUR - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
F16B 39/02 - Blocage des vis, boulons ou écrous dans lequel le verrouillage s'effectue après vissage
F16L 15/00 - Raccords avec filetage; Formes des filetages pour ces raccords
F16L 41/10 - Raccordements des tuyaux aux parois ou à d'autres tuyaux, dans lesquels l'axe du tuyau est perpendiculaire au plan de la paroi ou à l'axe de l'autre tuyau l'embout du tuyau étant vissé dans la paroi
60.
SYSTEM AND METHOD FOR DETERMINING ROTOR WHIRL DISPLACEMENT
An assembly for rotational equipment includes a rotor, at least one sensor, and a controller. The rotor includes a first plurality of teeth arranged on the rotor in a first circumferential array of teeth at a first axial position. The at least one sensor includes a first sensor positioned radially adjacent the first circumferential array of teeth at the first axial position. The at least one sensor is configured to generate an output signal waveform. The controller is in signal communication with the at least one sensor. The controller includes a processor and non- transitory memory in signal communication with the processor. The non-transitory memory stores instructions which, when executed by the processor, cause the processor to measure a dynamic whirl displacement of the rotor at the first axial position using the output signal waveform.
61.
HYBRID-ELECTRIC AIRCRAFT PROPULSION SYSTEM AND METHOD
A propulsion system for an aircraft is provided that includes an electric generator, a compressor, an internal combustion (IC) engine, a turbine, an electric power storage unit, and an electric motor. The compressor is configured to selectively produce a flow of compressor air at an air pressure greater than an ambient air pressure. The IC engine is configured to selectively intake compressor air during operation and produce an exhaust gas flow during operation. The turbine, powered by exhaust gas flow, is in communication with and configured to power the compressor and the electric generator. The electric power storage unit is in communication with the electric generator. The electric motor is in communication with the IC engine. The electric motor is powered by the electrical power produced by the electric generator, and the electric motor is configured to selectively provide motive force to the IC engine.
B64D 27/00 - Disposition du montage des groupes moteurs sur aéronefs; Aéronefs ainsi caractérisés
B64D 35/08 - Transmission de la puissance du groupe moteur aux hélices ou aux rotors; Aménagements des transmissions caractérisée par le fait que la transmission est entraînée par plusieurs groupes moteurs
62.
GAS TURBINE INTAKE FOR AIRCRAFT ENGINE AND METHOD OF INSPECTION THEREOF
The gas turbine intake can have a swirl housing assembly with a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending around the central axis and between the tangential inlet and the annular outlet, the swirl housing assembly having a proximal portion defining a first portion of the swirl path, a distal portion defining a second portion of the swirl path, vanes located in the swirl housing assembly, the vanes circumferentially interspaced from one another relative the central axis and extending between the proximal portion and the distal portion, the proximal portion fastened to the distal portion via a plurality of fasteners, a gasket sandwiched between the proximal portion and the distal portion by the plurality of fasteners, the gasket extending in a radial plane relative the central axis.
F01D 1/06 - "Machines" ou machines motrices à déplacement non positif, p.ex. turbines à vapeur avec des moyens stationnaires de guidage de fluide de travail et un rotor à ailettes ou de structure analogue traversées par le fluide de travail principalement dans le sens radial
F01D 9/06 - Conduits d'admission du fluide à l'injecteur ou à l'organe analogue
F01D 25/30 - Têtes d'évacuation, chambres ou parties analogues
F02C 6/12 - Turbocompresseurs de suralimentation, c. à d. ensembles fonctionnels destinés à augmenter la sortie de puissance mécanique des moteurs à piston à combustion interne en augmentant la pression de suralimentation
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
63.
DAMPER SEGMENT FOR PRESSURIZED GAS PIPE OF AIRCRAFT ENGINE
The damper segment can be assembled between adjacent segments of a pressurized gas pipe of an aircraft engine. The damper segment can have a proximal end, a distal end, a rigid tube at the proximal end, a damper tube extending between the rigid tube and the distal end, the damper tube being made of a metal mesh, a proximal catch structurally connecting a proximal end of the damper tube to the rigid tube, and a distal catch structurally connected between a distal end of the damper tube and the distal end, the damper tube having an unsupported length extending between the distal catch and the proximal catch, the rigid tube having a liner portion projecting into the distal segment, the liner portion extending internally relative the damper tube.
F02C 6/04 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p.ex. sans sortie de puissance mécanique
F01D 9/06 - Conduits d'admission du fluide à l'injecteur ou à l'organe analogue
F01D 13/02 - Couplage à fluide énergétique commun entre "machines" ou machines motrices
F01D 25/30 - Têtes d'évacuation, chambres ou parties analogues
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
A fuel system of an aircraft engine, has: a fuel manifold having a first manifold inlet and a second manifold inlet; a transfer tube assembly having a first tube slidably engaged to the fuel manifold and fluidly connected to the first manifold inlet, and a second tube slidably engaged to the fuel manifold and fluidly connected to the second manifold inlet; and an adaptor having: a body slidably engaged by the first tube and by the second tube, a first member defining a first fuel conduit fluidly connected to the first manifold inlet via the first tube, and a second member defining a second fuel conduit fluidly connected to the second manifold inlet via the second tube.
A probe shielding arrangement comprises a sleeve having a radially inner end mounted to a turbine housing and a radially outer end floatingly received in a probe boss on an exhaust case. The sleeve circumscribes an annular cavity around the probe. The annular cavity is sealed at opposed ends thereof to form a dead air cavity around the probe for insulation purposes.
A vane system for an aircraft engine, comprising: an inner wall extending circumferentially about a duct axis; an outer wall extending circumferentially about the duct axis radially outward of the inner wall relative to the duct axis; at least one vane extending from an inner end attached to the inner wall to an outer end rotatably connected to the outer wall, the outer end rotatable relative to the outer wall about a vane axis at an angle to the duct axis; a ring extending circumferentially about the duct axis radially outward of the outer wall relative to the duct axis, the ring rotatable about the duct axis; and at least one transmission member located radially outward of the outer wall relative to the duct axis and coupling the ring to the outer end such that rotating the ring about the duct axis rotates the outer end about the vane axis.
F01D 17/16 - Organes de commande terminaux disposés sur des parties du stator faisant varier l'aire effective de la section transversale des injecteurs ou tuyères de guidage en obturant les injecteurs
A shaft assembly for an aircraft powerplant, comprising: a shaft extending along an axis from a first shaft end to a second shaft end; a bearing assembly extending about the axis and supporting the first shaft end of the shaft, the bearing assembly including an inner race secured to the shaft and an outer race radially outward of the inner race relative to the axis; a seal extending about the axis and located radially outward of the shaft, the seal disposed axially between the bearing assembly and the second shaft end; a housing having a housing wall located between the bearing assembly and the seal; and a washer extending about the axis and located axially between the bearing assembly and the seal, the washer extending axially from the outer race to the housing wall.
There is provided an exhaust mixer arrangement for a turbofan engine having a bypass passage for channelling a bypass flow and a core passage for channelling a core flow around a central axis. The exhaust mixer arrangement comprises a mixer body mounted for rotation about the central axis. The mixer body has an annular wall extending around the central axis. The annular wall defines a plurality of circumferentially distributed alternating inner and outer lobes, with each inner lobe protruding into the core passage, and each outer lobe protruding into the annular bypass passage. A driving unit is operatively connected to the mixer body for selectively driving the mixer body in rotation about the central axis. A controller is operatively connected to the driving unit for controlling a rotational speed of the mixer body as a function of a flight operating condition.
F02K 1/46 - Tuyères comportant des moyens pour ajouter de l'air au jet ou pour augmenter la zone de mélange du jet et de l'air ambiant, p.ex. pour réduire le bruit
F02K 1/38 - Introduction d'air à l'intérieur du jet
69.
AIRCRAFT ENGINE, GAS TURBINE INTAKE THEREFORE, AND METHOD OF GUIDING EXHAUST GASSES
A gas turbine intake can have a swirl housing having a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the tangential inlet, and a plurality of vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis, each vane having a twisted and flat body having a length extending from a leading end to a trailing end, the leading end being oriented mainly circumferentially and axially at the swirl path, the trailing end being oriented mainly axially and radially at the annular outlet, the twisted and flat body twisting between the leading end and the trailing end around the central axis, around a radial axis perpendicular to the central axis, and around a tangential axis perpendicular to both the central axis and the radial axis.
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
F01D 1/06 - "Machines" ou machines motrices à déplacement non positif, p.ex. turbines à vapeur avec des moyens stationnaires de guidage de fluide de travail et un rotor à ailettes ou de structure analogue traversées par le fluide de travail principalement dans le sens radial
The gas turbine intake can have a swirl housing having an inlet portion fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the inlet portion to a circumferential outlet, the circumferential outlet fluidly connected back into the inlet portion, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis and located radially inwardly from the swirl path relative the central axis, the swirl path being free of the vanes.
B64D 27/02 - Aéronefs caractérisés par le type ou la position des groupes moteurs
F02M 26/01 - Recirculation des gaz d’échappement internes, c. à d. où les gaz d’échappement résiduels sont piégés dans le cylindre ou repoussés du collecteur d’admission ou d’échappement vers la chambre de combustion sans l’utilisation de passages additionnels
B64D 33/00 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
There is provided a load transfer interface in an aircraft engine for transferring a load from a bearing housing to an engine casing. The load transfer interface comprises a first component operatively coupled to and receiving the load from the bearing housing. The first component has a first annular body with a spigot extending axially from the first annular body. The interface comprises a second component operatively coupled to the first component and to the engine casing. The second component has a second annular body with a spigot-receiving cavity disposed therein. The spigot-receiving cavity is shaped and positioned to receive the spigot of the first component. The second component receives the load from the first component and transfers the load to the engine casing.
F02C 7/06 - Aménagement des paliers; Lubrification
F01D 25/16 - Aménagement des paliers; Support ou montage des paliers dans les stators
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
F02C 3/08 - 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 comprenant au moins un étage radial
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnel; Disposition permettant la dilatation calorifique ou le déplacement
A turbine assembly of an aircraft engine includes a cooling system for optimizing a tip clearance gap defined between an inner surface of a turbine housing and blade tips of the turbine blades. The cooling system includes a cooling airflow passage located radially outward from the turbine housing and being in heat-transfer communication with the turbine housing. The cooling airflow passage receives a flow of cooling air therethrough for cooling the turbine housing. A heat sink is disposed on the outer surface of the turbine housing within the cooling airflow passage, the heat sink including heat transfer elements projecting into the cooling airflow passage away from the outer surface of the turbine housing. The heat transfer elements are in convective heat transfer relationship with the flow of cooling air in the cooling airflow passage.
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
B64D 27/16 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à réaction
F02C 7/12 - Refroidissement des ensembles fonctionnels
73.
AIRCRAFT PROPULSION SYSTEM WITH INTERMITTENT COMBUSTION ENGINE AND ELECTRIC TRANSMISSION SYSTEM AND METHOD FOR OPERATING THE SAME
A propulsion system for an aircraft having a nacelle and a fuselage is provided. The nacelle has a gas flow path and a nacelle interior region. The system includes a compressor section, an intermittent IC engine, a turbine section, a fan, and an IC cooling system. A first electric motor powers the compressor section. The compressor section produces a flow of elevated pressure compressor air. The intermittent IC engine selectively intakes the compressor air flow and produces an exhaust gas flow. The turbine section powered by exhaust gas in turn powers a first electric generator. The fan is driven by a second electric motor. The IC engine cooling system has a heat exchanger disposed within the gas flow path, coolant, coolant piping, and a pump. The heat exchanger is disposed in the nacelle.
B64D 27/00 - Disposition du montage des groupes moteurs sur aéronefs; Aéronefs ainsi caractérisés
B64D 33/00 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs
F02C 1/00 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de gaz chauds ou de gaz sous pression non chauffés, comme fluide de travail
74.
ADJUSTING AIRCRAFT POWERPLANT POWER SPLIT TO CONTROL POWERPLANT VIBRATI ONS
A method is provided for operating a system of an aircraft. During this method, rotation of a propulsor rotor is driven using mechanical power output by a powerplant. The powerplant includes a first drive device and a second drive device. The first drive device generates a first portion of the mechanical power. The second drive device generates a second portion of the mechanical power. A ratio between the first portion of the mechanical power and the second portion of the mechanical power is adjusted to control vibrations of the powerplant.
B64D 31/18 - pour les groupes moteurs hybrides-électriques
B64D 31/00 - Commande des groupes moteurs; Leur disposition
F16F 15/00 - Suppression des vibrations dans les systèmes; Moyens ou dispositions pour éviter ou réduire les forces de déséquilibre, p.ex. dues au mouvement
G05D 19/00 - Commande des oscillations mécaniques, p.ex. de l'amplitude, de la fréquence, de la phase
Blowdown valves and associated methods for separating oil and air in a lubrication system of an aircraft engine are provided. A method includes receiving a mixture of air and oil at a blowdown valve including a valve member movable between a valve- closed position and a valve-open position. With the valve member in the valve-open position, the method includes impinging the mixture against the valve member. A first portion of the mixture having a first fraction of oil is released from the blowdown valve upstream of the valve member. A second portion of the mixture having a second fraction of oil greater than the first fraction of oil is guided around and past the valve member. The second portion of the mixture is released from the blowdown valve downstream of the valve member.
An article has a nickel-based alloy substrate having, in weight percent: 5.4- 7.4 Re; 4.1-5.9 Ru; 3.0-6.2 Cr; 3.0-10.0 Co; 0.5-3.8 Mo; 3.0-6.0 W; 4.6-8.6 Ta; 5.0-6.4 Al; 0.050- 0.30 Hf; no more than 0.50 all other elements, if any, individually; and no more than 2.0 all other elements, if any, combined. A nickel-based coating is on the substrate and comprising, in weight percent: 6.0-10.0 Al; 4.0-15.0 Cr; 11.0-15.0 Co; 0.1-1.0 Hf; 0.1-1.0 Si; 0.1-1.0 Y; up to 1.0 Zr if any; up to 7.0 Ta if any; up to 6.0 W if any; no more than 1.0 all other elements, if any, individually; and no more than 4.0 all other elements, if any, combined.
A turbine assembly in a turbine section of an aircraft engine includes a rotor with blades having blade tips, and a turbine housing radially surrounding the blades. A distance between an inner surface of the housing and the blade tips defines a tip clearance gap. A passive cooling system for optimizing the tip clearance gap includes a cooling airflow passage located radially outward from, and in heat-transfer with, the turbine housing. The cooling airflow passage has an inlet opening located upstream of the rotor and an exit opening located downstream of the rotor. The inlet opening provides air flow into the cooling airflow passage. The exit opening provides air flow communication between the cooling airflow passage and a main gaspath of the turbine section. A flow of cooling air through the cooling airflow passage is induced, to cool the housing.
A seal for an aircraft engine includes an annular body receivable in interference fit in a radial spacing defined between inner and outer cylindrical components of the aircraft engine, the annular body defining a central axis coaxial with the inner and outer cylindrical components, the annular body including an inner portion defining an inner diameter, an outer portion defining an outer diameter, and an intermediate portion extending between the inner and outer portions. The outer portion is slidably engageable to the outer cylindrical component at an outer contact sealing portion of the annular body. One of the inner portion and the outer portion defines at least one cut extending from the one of the inner portion and the outer portion toward another one of the inner portion and the outer portion. A method for sealing a radial spacing between coaxial cylindrical components in an aircraft engine is also described.
F02C 7/28 - Agencement des dispositifs d'étanchéité
F16J 15/3236 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre par joints élastiques, p.ex. joints toriques avec au moins une lèvre ayant plusieurs lèvres dont au moins une lèvre pour chaque surface, p.ex. conditionnements en U
A containment assembly is provided for an aircraft engine having a rotor with a set of blades. The containment assembly comprises a containment casing annularly surrounding the rotor radially outward of the set of blades. The containment casing is made of a material having a density less than that of steel. A layer of thermal insulation is disposed radially inward of the containment casing. The layer of thermal insulation is radially disposed between the containment casing and the set of blades.
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
B64D 29/00 - Nacelles, carénages ou capotages des groupes moteurs
F01D 25/14 - Carcasses d'enveloppe modifiées à cet effet
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
A flow measuring system for one or more oil nozzles of a lubrication system of an aircraft engine comprises an external supply of a testing liquid and a pump fluidly connecting the external supply of the testing liquid to an inlet of the lubrication system via a first conduit. The inlet to the lubrication system is disposed upstream of the one or more oil nozzles. A second conduit fluidly connects an outlet of the lubrication system to the external supply of the testing liquid. The outlet of the lubrication system is disposed downstream of the one or more oil nozzles. A flow measuring device is operable to measure a flow resistance through the one or more oil nozzles.
A valve for an air system in an aircraft engine, comprising: a housing defining a chamber having a valve axis; a body within the chamber about a piston axis collinear with the valve axis, extending from a first surface to a second surface, defining a bore extending from the first to the second surface, having a mating connector defined by the second surface and located radially outward of the bore relative to the piston axis; and a sleeve extending from a first end matingly engaged with the mating connector to a second end along a sleeve axis collinear with the valve axis, the first end axially stacked on the body via the first surface to define a first distance between the first end and the first surface, and via the second surface to define a second distance between the first end and the second surface greater than the first distance.
F16K 11/044 - Soupapes ou clapets à voies multiples, p.ex. clapets mélangeurs; Raccords de tuyauteries comportant de tels clapets ou soupapes; Aménagement d'obturateurs et de voies d'écoulement spécialement conçu pour mélanger les fluides dont toutes les faces d'obturation se déplacent comme un tout comportant uniquement des soupapes ou des clapets à corps de soupape ou de clapet mobiles situés entre des sièges de soupape ou de clapet
F16K 31/122 - Moyens de fonctionnement; Dispositifs de retour à la position de repos actionnés par un fluide le fluide agissant sur un piston
A rotor of an aircraft engine has a plurality of blades extending radially from a disc. At least one of the blades has an airfoil, a root and a tip. The airfoil has a crack-mitigating rib extending chordwise along the airfoil. The crack-mitigating rib is disposed radially closer to the root than to the tip.
An aircraft engine, has: an upstream stator having upstream stator vanes distributed about a central axis; and a downstream stator having downstream stator vanes distributed about the central axis, the downstream stator located downstream of the upstream stator, a number of the upstream stator vanes different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes; and a second vane differing from the first vane by a geometric parameter, the geometric parameter causing the second vane to have one or more of: a stiffness greater than that of the first vane, and a major portion of a leading edge of the second vane circumferentially overlapped by another one of the upstream stator vanes.
An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, the downstream stator vanes including: a first vane made of a first material, a major portion of a leading edge of the first vane circumferentially overlapped by one of the upstream stator vanes, and a second vane made of a second material having a greater stiffness, strength, and/or ductility than that of the first material, a major portion of a leading edge of the second vane exposed via a spacing defined between two of the upstream stator vanes.
F01D 9/02 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage
F01D 5/28 - Emploi de matériaux spécifiés; Mesures contre l'érosion ou la corrosion
F01D 25/00 - "MACHINES" OU MACHINES MOTRICES À DÉPLACEMENT NON POSITIF, p.ex. TURBINES À VAPEUR - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
A valve for an air system in an aircraft engine, comprising: a housing defining a chamber having a valve axis circumscribed by a sealing surface; and a piston assembly within the chamber including: a sealing ring; and a body extending annularly about a piston axis collinear with the valve axis, having a first and a second piston surface axially spaced apart, a radially outer piston surface extending axially and located between the first and second piston surfaces, and an annular groove extending radially inwardly from the radially outer piston surface having first and second groove walls spaced apart and axially facing one another, the sealing ring within the annular groove, the body including: a first member defining the first piston surface and the first groove wall; and a second member defining the second piston surface and the second groove wall, the first member and the second member in mating engagement.
F01D 17/14 - 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
F01D 25/28 - Dispositions pour le support ou le montage, p.ex. pour les carters de turbines
86.
JOINT BETWEEN GAS TURBINE ENGINE COMPONENTS WITH BONDED FASTENER(S)
An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a seal carrier, a seal land, a seal ring, a plate and a fastener. The seal carrier has an annular groove and extends between a first side and a second side. The seal land is opposite the annular groove. The seal ring seals a gap between the seal carrier and the seal land. The seal ring is seated in the annular groove. The plate is at the second side of the seal carrier. The fastener includes a head and an elongated member connected to the head. The head is at the first side of the seal carrier. The elongated member projects out from the head through the seal carrier, the seal ring and the plate. The elongated member is bonded to the plate.
An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a case, a housing, a component and a spring element. The case includes an aperture that extends axially along an axis through the case. The housing is attached to the case with a cavity formed by and axially between the housing and the case. The component includes a base and a projection. The base is disposed within the cavity and axially engages the case. The projection projects out from the base and axially through the aperture. The spring element is disposed within the cavity. The spring element is compressed axially between and engages the base and the housing.
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
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 25/00 - "MACHINES" OU MACHINES MOTRICES À DÉPLACEMENT NON POSITIF, p.ex. TURBINES À VAPEUR - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
F16J 15/06 - Joints d'étanchéité entre surfaces immobiles entre elles avec garniture solide comprimée entre les surfaces à joindre
88.
GAS TURBINE ENGINE WITH ELECTRIC MACHINE IN ENGINE CORE
A gas turbine engine assembly includes an engine core and an electric machine. The engine core includes a first rotating structure, a second rotating structure, a combustor and a flowpath. The first rotating structure includes a first structure turbine rotor. The second rotating structure includes a second structure compressor rotor, a second structure turbine rotor and a second structure shaft connecting the second structure compressor rotor to the second structure turbine rotor. The second structure compressor rotor, the combustor, the second structure turbine rotor and the first structure turbine rotor are arranged sequentially along the flowpath. The electric machine is arranged within the engine core. The electric machine includes an electric machine rotor and an electric machine stator adjacent the electric machine rotor. The electric machine rotor is rotatable with the second rotating structure and located between the second structure compressor rotor and the first structure turbine rotor.
A method for generating assembly instructions for a plurality of 3D component models including a first 3D component model and a second 3D component model. The first 3D component model includes a first geometric feature, and the second 3D component model includes a second geometric feature. The method includes determining first assembly instructions for assembling the plurality of 3D component models into a first 3D model assembly, determining a plurality of assembly constraints for assembling the plurality of 3D component models into the first 3D model assembly using the first assembly instructions, modifying the plurality of 3D component models, and generating second assembly instructions for assembling the modified plurality of 3D component models into a second 3D model assembly. The second assembly instructions are different than the first assembly instructions.
An aircraft engine oil filler apparatus includes a filler tube configured to be connected to an oil tank such that a bottom portion of the filler tube is disposed inside the oil tank, a valve received in the bottom portion of the filler tube movable between an open position in which the valve hydraulically connects the filler tube to the oil tank, and a closed position in which the valve hydraulically disconnects the filler tube from the oil tank, and a float disposed above the valve and operatively connected to the valve to move the valve from the open position to the closed position when oil inside the oil tank rises to a threshold level. The valve is movable independently from the float when pressure in the oil tank is greater than pressure in the filler tube. A method of operation of an oil filler apparatus is also described.
An aircraft engine, has: an upstream stator having upstream stator vanes circumferentially distributed about a central axis; and a downstream stator having downstream stator vanes circumferentially distributed about the central axis, the downstream stator located downstream of the upstream stator relative to an airflow flowing within a core gaspath of the aircraft engine, a number of the upstream stator vanes being different than a number of the downstream stator vanes, major portions of leading edges of the downstream stator vanes circumferentially overlapped by the upstream stator vanes, the downstream stator vanes including: a first pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a first pitch, and a second pair of circumferentially adjacent vanes of the downstream stator vanes spaced apart by a second pitch different than the first pitch.
A method for checking a closing point for a bleed-off valve for a gas turbine engine includes determining a modulation characteristic curve for the bleed-off valve, determining a nominal closing point value for the bleed-off valve on the modulation characteristic curve, operating the gas turbine engine and increasing an engine power of the gas turbine engine until the gas turbine engine parameter reaches a predetermined testing value, and determining a bleed-off valve measured value and a gas turbine engine measured value when the gas turbine engine parameter reaches the predetermined testing value. The gas turbine engine measured value is different than the nominal closing point value. The method further includes determining an extrapolated closing point value and checking the closing point for the bleed-off valve by comparing the bleed- off valve measured value or the gas turbine engine measured value to the extrapolated closing point value.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p.ex. à la température, à la pression, à la vitesse du rotor
93.
PRESS-FIT COMPONENTS DISASSEMBLY TOOLING AND PROCESS
A tool and an associated process for removing a press-fit component from an associated engine component, wherein the engine component is pushed in a first axial direction while the press-fit component is restrained from axially moving in the first axial direction.
B23P 19/02 - 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 déformation; Outils 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
94.
TOOLING SYSTEM AND METHODS OF ASSEMBLING AND DISASSEMBLING A ROTARY ASSEMBLY THEREWITH
The tooling system can have an extension unit having a pushing member and a pulling member extending along a length, the pushing member and the pulling member in sliding engagement and displaceable relative to one another, the extension unit having an internal passage configured for receiving a portion of the shaft via the first end; a pushing adapter engageable with and disengageable from the pushing member at the first end, and further engageable with and disengageable from one of the shaft and the component; and a pulling adapter engageable with and disengageable from the pulling member at the first end, and further engageable with and disengageable from an other one of the shaft and the component.
B23P 21/00 - Machines pour l'assemblage de nombreuses pièces différentes destinées à composer des ensembles, avec ou sans usinage de ces pièces avant ou après leur assemblage, p.ex. à commande programmée
F02C 7/00 - Caractéristiques, parties constitutives, détails ou accessoires non couverts dans, ou d'un intérêt plus général que, les groupes ; Entrées d'air pour ensembles fonctionnels de propulsion par réaction
A structure is provided for a gas turbine engine. This gas turbine engine structure includes an engine case, an engine pylon and an engine line. The engine case includes a base, a mounting boss and a first support element. The base extends axially along and circumferentially about an axial centerline of the engine case. The mounting boss projects radially out from the base. The first support element projects radially out from the base and laterally out from the mounting boss. The first support element is configured as or otherwise includes a peripheral boss. The engine pylon is mounted to the mounting boss. The engine line is coupled to the peripheral boss.
A structure for a gas turbine engine includes a first engine component, a second engine component and fasteners. The component apertures include first fastener apertures and intergroup apertures. The first fastener apertures are arranged into a plurality of groups. The first group is fomied by Ni-number of the first fastener apertures. The second group is formed by N2-number of the first fastener apertures where the N2-number is different than the Ni- number. Each of the intergroup apertures is disposed circumferentially between and adjacent a respective circumferentially neighboring pair of the groups. The second engine component includes a surface and second fastener apertures. The surface axially engages the first engine component and covers the intergroup apertures. The fasteners attach the first engine component and the second engine component together. Each of the fasteners is mated with one of the first fastener apertures and one of the second fastener apertures.
A gas turbine engine includes an intake device. The intake device includes a snorkel and a filter case. The snorkel includes a tubular body and an inlet aperture. The tubular body extends between a closed end and an open end opposite the closed end. The inlet aperture is formed through the tubular body proximate the closed end. The tubular body forms a first portion of a gas flow path for a bleed gas from the inlet aperture to the open end. The filter case is connected to the tubular body. The filter case extends between a first end and a second end. The filter case includes a sidewall extending from the first end to the second end. The sidewall surrounds a filter cavity. The filter case is configured to receive the bleed gas from the open end of the tubular body. The filter case and the snorkel form a unitary component.
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 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p.ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
A method is provided for an engine. During this method, a database is provided for a parameter of the engine. The database includes a plurality of values for the parameter determined over a period of time. Confidence bands are established using a probability density function on the database. An action is performed in response to a comparison of a first updated value for the parameter to the confidence bands. The engine may be configured as a gas turbine engine or another type of heat engine.
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gaz; Commande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
F01D 21/00 - Arrêt des "machines" ou machines motrices, p.ex. dispositifs d'urgence; Dispositifs de régulation, de commande ou de sécurité non prévus ailleurs
F02C 3/00 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p.ex. à la température, à la pression, à la vitesse du rotor
An inlet guide vane assembly for an aircraft engine includes an array of inlet guide vanes having radially inner ends at a radially inner shroud, radially outer ends at a radially outer shroud, and airfoils extending therebetween. An internal passage extends radially through the airfoil from a vane air inlet at the radially inner end to a vane air outlet at the radially outer end. The vane air inlet is in fluid communication with an inner plenum, disposed radially inwardly of the inlet guide vane and in fluid communication with an anti-icing air source. The vane air outlet in fluid communication with an outer plenum, disposed radially outwardly of the inlet guide vane and having an exhaust port. A vane anti-icing pathway is defined in a radially- outward direction from the inner plenum, through the internal passage of the vane, and into the outer plenum.
B64D 33/02 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des entrées d'air de combustion
B64D 15/02 - Dégivrage ou antigivre des surfaces externes des aéronefs par gaz chaud ou liquide amené par conduit
An intake device for a gas turbine engine includes a snorkel and a particle separator. The snorkel is configured to be mounted to a panel defining at least a portion of a gas flow path within the gas turbine engine. The snorkel includes a tubular body extending between a closed end and an open end opposite the closed end. The snorkel further includes an inlet aperture formed through the tubular body adjacent the closed end. At least a portion of the snorkel is configured to be disposed within the gas flow path. The particle separator is mounted to the snorkel downstream of the inlet aperture. The particle separator includes at least one gas flow passage extending between a flow inlet and a flow outlet. The at least one gas flow passage is configured to remove particulate matter from the at least one gas flow passage upstream of the flow outlet.
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
B01D 50/20 - Combinaisons de dispositifs couverts par les groupes et
B01D 45/04 - Séparation de particules dispersées dans des gaz ou des vapeurs par gravité, inertie ou force centrifuge par inertie
B01D 45/12 - Séparation de particules dispersées dans des gaz ou des vapeurs par gravité, inertie ou force centrifuge en utilisant la force centrifuge
F02C 6/06 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p.ex. sans sortie de puissance mécanique délivrant des gaz comprimés
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