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 11/40 - Mécanismes de changement de pas des pales par fluide, p.ex. hydrauliques automatiques
F03D 7/02 - Commande des mécanismes moteurs à vent les mécanismes moteurs à vent ayant l'axe de rotation sensiblement parallèle au flux d'air pénétrant dans le rotor
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
2.
Controlling rate of rotor feather by primary blade angle control system
A rotor blade control system includes a main control valve having an inlet for receiving liquid and an outlet for issuing liquid to a rotor pitch change actuator. The main control valve is configured to control flow of liquid from the inlet to the outlet to modify pitch angle of rotor blades. A feathering system has a first conduit in fluid communication with the outlet of the main control valve, a second conduit in fluid communication with the rotor pitch change actuator, and a drain conduit in fluid communication with a liquid return system. The feathering system has a normal operation mode for supplying liquid from the main control valve to the rotor pitch change actuator, and a feathering mode for allowing drainage from the rotor pitch change actuator to the drain conduit across a range of flow rates.
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
4.
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.
5.
HYBRID ELECTRIC POWERPLANT SYSTEMS AND CONTROLLERS
A hybrid electric propulsion (HEP) system can include a heat engine torque sensor connected between a heat engine and a combining gear box to sense a heat motor input torque input to the combining gear box, an electric motor torque sensor connected between an electric motor and the combining gear box to sense an electric motor input torque input to the combining gear box, and a combining gear box torque sensor connected to an output of the combining gearbox. The system can include a HEP controller operatively connected to each of the heat engine torque sensor, the electric motor torque sensor, and the combining gear box torque sensor to receive one or more torque signals therefrom. The controller can be configured to output one or more output signals as a function of the signals from each of the heat engine torque sensor, the electric motor torque sensor, and the combining gear box torque sensor.
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.
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
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
F01D 25/32 - Recueil de l'eau de condensation; Drainage
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.
An exhaust duct of an aircraft engine includes an annular inlet conduit having an inlet central axis, and at least two outlet conduits in flow communication with the inlet conduit. The at least two outlet conduits are located non-parallel to the inlet central axis. Each of the at least two outlet conduits include an outlet port defining a distal end of each of the two outlet conduits. At least one of the outlet ports is non-circular in cross-sectional shape.
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
11.
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 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.
H02H 7/08 - Circuits de protection de sécurité spécialement adaptés pour des machines ou appareils électriques de types particuliers ou pour la protection sectionnelle de systèmes de câble ou ligne, et effectuant une commutation automatique dans le cas d'un chan pour moteurs dynamo-électriques
B60W 10/08 - Commande conjuguée de sous-ensembles de véhicule, de fonction ou de type différents comprenant la commande des ensembles de propulsion comprenant la commande des unités de traction électrique, p.ex. des moteurs ou des générateurs
13.
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.
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
G01M 3/26 - Examen de l'étanchéité des structures ou ouvrages vis-à-vis d'un fluide par utilisation d'un fluide ou en faisant le vide par mesure du taux de perte ou de gain d'un fluide, p.ex. avec des dispositifs réagissant à la pression, avec des indicateurs de débit
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
G07C 5/00 - Enregistrement ou indication du fonctionnement de véhicules
The gas turbine engine can have an engine core; a core output shaft drivable by the engine core; a power output shaft; an auxiliary power shaft; and a reduction gearbox having gears, the gears drivingly connecting the core output shaft to the auxiliary power shaft. The gears can include an epicyclic gearing drivingly connecting the core output shaft and the auxiliary power shaft to the power output shaft. The gas turbine engine can further have a second auxiliary power shaft interconnected to the auxiliary power shaft, the power output shaft, and the core output shaft by the gears.
F02C 7/32 - Aménagement, montage ou entraînement des auxiliaires
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
A 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.
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
An integrally bladed rotor (IBR) for a gas turbine engine and method is provided. The IBR is configured for use in blade off testing and includes a hub, a plurality of rotor blades, a central passage, and first and second lateral cavities. The hub has forward and aft ends and a circumferentially extending exterior surface. The central passage is disposed in the hub radially below a test rotor blade, extending along a path between an inlet at or forward of the test blade leading edge and an outlet at or aft of the test blade trailing edge. The first and second lateral cavities are disposed in the hub, extending generally parallel to the central passage path, on opposite circumferential sides. The first lateral cavity is disposed a distance (MSD1) from the central passage and the second lateral cavity is disposed a distance (MSD2) from the central passage.
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
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
20.
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
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.
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.
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.
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 ({dot over (W)}HPT) of the high-pressure turbine, determine an estimated inlet temperature value for the high-pressure turbine using the estimated work ({dot over (W)}HPT), 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.
G01K 13/024 - Thermomètres spécialement adaptés à des fins spécifiques pour mesurer la température de fluides en mouvement ou de matériaux granulaires capables de s'écouler de gaz en mouvement
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
G01K 1/02 - Moyens d’indication ou d’enregistrement spécialement adaptés aux thermomètres
G01K 1/14 - Supports; Dispositifs de fixation; Dispositions pour le montage de thermomètres en des endroits particuliers
G01L 3/24 - Dispositifs pour déterminer la valeur de la puissance, p.ex. en mesurant et en multipliant simultanément les valeurs du couple par le nombre de tours par unité de temps, en multipliant les valeurs de la force de traction ou propulsive par la vitesse
26.
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.
B64D 33/04 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des sorties d'échappement ou des tuyères
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
B64D 29/00 - Nacelles, carénages ou capotages des groupes moteurs
F02K 1/36 - 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 comportant un éjecteur
Aircraft power plants and associated methods are provided. A method for driving a load on an aircraft includes: transferring motive power from an internal combustion (IC) engine to the load; discharging a flow of first exhaust gas from the IC engine when transferring motive power from the IC engine to the load; receiving the flow of first exhaust gas from the IC engine into a combustor; mixing fuel with the first exhaust gas in the combustor and igniting the fuel to generate a flow of second exhaust gas; receiving the flow of second exhaust gas at a turbine and driving the turbine with the flow of second exhaust gas from the combustor; and transferring motive power from the turbine to the load.
B64D 27/04 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à pistons
F02B 3/10 - Moteurs caractérisés par la compression d'air et l'addition subséquente de combustible avec allumage par compression avec introduction intermittente de combustible
F02B 37/00 - Moteurs caractérisés par l'utilisation de pompes entraînées au moins temporairement par les gaz d'échappement
F02B 53/10 - Alimentation en combustible; Introduction du combustible dans la chambre de combustion
F02B 53/14 - Adaptation des moteurs pour l'entraînement d'autres dispositifs ou combinaisons des moteurs avec ceux-ci
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.
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
30.
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
31.
Systems and methods for controlling an air flow path for a propulsion system air intake
An air intake for an aircraft propulsion system includes an air inlet duct, a core flow duct, a bypass flow duct, a splitter, and a flow control device. The air inlet duct includes an intake inlet and a gas path floor. The core flow duct includes a core flow outlet. The bypass flow duct includes a bypass flow outlet. The bypass flow duct includes the gas path floor. The splitter separates the core flow duct and the bypass flow duct. The flow control device is disposed on a portion of the gas path floor. The flow control device is configured to be selectively positioned to control an air flow path for air flowing through the air inlet duct, the core flow duct, and the bypass flow duct.
F02C 7/05 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs pour empêcher la pénétration d'objets ou de particules endommageantes
F02C 7/042 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction à géométrie variable
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
A method of operating a reverse thrust system of an aircraft engine, the method comprising: receiving a status signal indicative that the aircraft is on-ground or in-flight; and upon detecting that the aircraft is on-ground, overriding a protection module such that the reverse thrust system is operable regardless of the protection module being in an active state or in a disabled state, the protection module causing, absent the overriding, the reverse thrust system to be inoperable when in the active state. An aircraft comprising: an engine including a reverse thrust system; a thrust control input device configured for generating an input signal indicative of a reverse thrust demand; at least one sensor configured for generating at least one status signal indicative of the aircraft being in-flight or on-ground; and a control system electronically connected with the at least one sensor, the thrust control input device and the reverse thrust system.
F02K 1/76 - Commande ou régulation des inverseurs de poussée
F02K 1/60 - Inversion du jet principal par blocage de l'échappement vers l'arrière à l'aide d'éléments pivotants ayant la forme de paupières ou de coquilles, p.ex. inverseurs du type se trouvant en aval de la sortie de la tuyère en position de fonctionnement
F02K 1/66 - Inversion du flux de la soufflante en inversant les aubes du ventilateur
F02K 1/70 - Inversion du flux de la soufflante utilisant des volets inverseurs de poussée ou des portes montées sur le carter de la soufflante
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
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.
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.
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.
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.
G01N 23/046 - Recherche ou analyse des matériaux par l'utilisation de rayonnement [ondes ou particules], p.ex. rayons X ou neutrons, non couvertes par les groupes , ou en transmettant la radiation à travers le matériau et formant des images des matériaux en utilisant la tomographie, p.ex. la tomographie informatisée
42.
ADAPTIVE COMPONENT OVERHAUL USING STRUCTURED LIGHT SCAN DATA
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.
B23K 31/12 - Procédés relevant de la présente sous-classe, spécialement adaptés à des objets ou des buts particuliers, mais non couverts par un seul des groupes principaux relatifs à la recherche des propriétés, p.ex. de soudabilité, des matériaux
43.
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 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 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.
B29C 64/393 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
B29C 64/153 - Procédés de fabrication additive n’utilisant que des matériaux solides utilisant des couches de poudre avec jonction sélective, p.ex. par frittage ou fusion laser sélectif
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.
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
55.
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
Methods and systems of operating a gas turbine engine in a low-power condition are provided. In one embodiment, the method includes supplying fuel to the combustor by supplying fuel to the first fuel manifold via a first flow divider valve and supplying fuel to the second fuel manifold via a second flow divider valve. While supplying fuel to the combustor by supplying fuel to the first fuel manifold, the method includes stopping supplying fuel to the second fuel manifold and supplying pressurized gas to the second fuel manifold via the second flow divider valve to flush fuel in the second fuel manifold into the combustor and hinder coking in the second fuel manifold and associated nozzles.
F02C 7/30 - Prévention de la corrosion dans les espaces balayés par les gaz
F02C 6/00 - Ensembles fonctionnels multiples de turbines à gaz; Combinaisons d'ensembles fonctionnels de turbines à gaz avec d'autres appareils; Adaptations d'ensembles fonctionnels de turbines à gaz à des applications particulières
F02C 6/16 - Ensembles fonctionnels de turbines à gaz comportant des moyens pour emmagasiner l'énergie, p.ex. pour faire face à des pointes de charge pour emmagasiner de l'air comprimé
F02C 7/22 - Systèmes d'alimentation en combustible
F02C 7/232 - Soupapes pour combustible; Systèmes ou soupapes de drainage
F02C 7/236 - Systèmes d'alimentation en combustible comprenant au moins deux pompes
F02C 9/42 - Commande de l'alimentation en combustible spécialement adaptée à la commande simultanée d'au moins deux ensembles fonctionnels
A fuel injector for a gas turbine engine combustor is provided that includes a swirler, a mounting stage, and a distributor. The swirler has a shaft, a collar, a throat section, and first and second axial ends. The throat section includes an inner radial surface that defines a central passage that extends between the swirler inner bore and the collar. The collar includes a plurality of apertures extending therethrough disposed radially outside of the central passage. The mounting stage is disposed in the inner bore, and has an annular flange, a central hub, and at least one strut. The distributor has a stem attached to a head. The stem has a distal end opposite the head portion engaged with the central hub. The head portion has an end surface and a side surface. The distributor is selectively positionable relative to the throat section.
F23R 3/12 - Aménagements de l'entrée d'air pour l'air primaire créant un tourbillon
F23R 3/22 - Moyens de stabilisation de la flamme, p.ex. accroche-flamme de postcombustion d'ensembles fonctionnels à propulsion par réaction réglables, p.ex. autoréglables
F23R 3/28 - Chambres de combustion à combustion continue utilisant des combustibles liquides ou gazeux caractérisées par l'alimentation en combustible
59.
System and method for controlling fluid flow with a pressure relief valve
A fluid supply system and method is provided that includes a fluid pump, a pressure sensor, a pressure relief valve (PRV), and a fluid monitoring device. The fluid pump receives fluid from a first conduit and discharges fluid into a second conduit. The pressure sensor produces sensed fluid pressure signals. The PRV is in signal communication with the pressure sensor. The fluid monitoring device includes a control orifice in fluid communication with second and third conduits. The second conduit has a first inner diameter, the third conduit has a second inner diameter, and the control orifice has an orifice inner diameter, and the orifice inner diameter is less than the first and second inner diameters. The pressure sensor senses fluid pressure in the third conduit at a position in close proximity to the control orifice. The fluid monitoring device may be in a lead or a lag domain configuration.
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 method of operating a variable guide vane assembly of an aircraft engine, the variable guide vane assembly including guide vanes rotatable about respective spanwise axes and circumferentially distributed about a central axis, the method comprising: obtaining a target exit flow angle defined between a direction of a flow exiting the guide vanes and the central axis; predicting an exit flow angle as a function of at least a geometric angle, the exit flow angle defined between the direction of the flow exiting the guide vanes and the central axis, the geometric angle defined between the guide vanes and the central axis; and when a difference between the exit flow angle and the target exit flow angle is above a threshold, modulating the guide vanes to modify the geometric angle until the difference between the exit flow angle and the target exit flow angle is at or below the threshold.
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 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 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.
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.
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
64.
SYSTEMS AND METHODS FOR DETERMINING GAS TURBINE ENGINE OPERATING MARGINS
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.
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.
66.
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
69.
METHOD OF REPAIRING A COMBUSTOR LINER OF A GAS TURBINE ENGINE
Methods and systems for characterizing holes in a combustor liner of a gas turbine engine, and associated repair methods are provided. One method comprises receiving first measured data of the combustor liner in an uncoated state. The method includes determining a first location and a first orientation of a first hole and a first location and a first orientation of a second hole in the combustor liner using the first measured data. The method includes receiving second measured data of the combustor liner in a coated state where the second hole is at least partially obstructed by a coating and the first hole is substantially unobstructed by the coating. The method includes inferring a second location of the second hole of the combustor liner in the coated state using a known spacing between the first location of the first hole and the first location of the second hole. The characterization of the holes may be used to re-drill the obstructed second hole.
B29C 73/26 - Appareils ou accessoires non prévus ailleurs pour le prétraitement mécanique
G05B 19/402 - 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 des dispositions de commande pour le positionnement, p.ex. centrage d'un outil par rapport à un trou dans la pièce à usiner, moyens de détection additionnels pour corriger la position
B23B 35/00 - Méthodes d'alésage ou de perçage ou autres méthodes de travail impliquant l'utilisation de machines à aléser ou à percer; Utilisation d'équipements auxiliaires en relation avec ces méthodes
70.
EXHAUST ASSEMBLY FOR PURGING A NACELLE CAVITY OF A PROPULSION SYSTEM
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.
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.
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 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
F02C 7/06 - Aménagement des paliers; Lubrification
F02C 7/22 - Systèmes d'alimentation en combustible
F02C 7/232 - Soupapes pour combustible; Systèmes ou soupapes de drainage
72.
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 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.
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 27/24 - Aéronefs caractérisés par le type ou la position des groupes moteurs utilisant la vapeur, l'électricité ou l'énergie de ressorts
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
74.
ACOUSTICAL HEALTH MONITORING OF GAS TURBINE ENGINES
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.
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
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
A rotary engine, has: an outer body defining a rotor cavity; a rotor rotatable within the rotor cavity and in sealing engagement with walls of the outer body and defining at least one chamber of variable volume in the rotor cavity; a pilot subchamber defined by the outer body, the pilot subchamber having an outlet in fluid flow communication with the rotor cavity; and a fuel injector having a tip in communication with the rotor cavity at a location spaced apart from the outlet of the pilot subchamber, the tip of the fuel injector having: a first outlet in fluid communication with the rotor cavity independently of the pilot subchamber; and a second outlet in fluid communication with the rotor cavity through the pilot subchamber.
F02B 53/10 - Alimentation en combustible; Introduction du combustible dans la chambre de combustion
F01C 1/22 - "Machines" ou machines motrices à piston rotatif du type à axe interne, avec mouvement relatif des organes coopérants dans le même sens aux points d'engagement ou dont l'un des organes coopérants est stationnaire, l'organe interne ayant plus de dents ou de parties équivalentes de prise que l'organe
F02M 21/02 - Appareils pour alimenter les moteurs en combustibles non liquides, p.ex. en combustibles gazeux stockés sous forme liquide en combustibles gazeux
F02M 61/14 - Disposition des injecteurs par rapport aux moteurs; Montage des injecteurs
The torque transfer assembly can have a torque shaft rotatable around a rotation axis, the torque shaft having a first end and a second end opposite the first end along the rotation axis, a first reference feature at an intermediary location between the first end and the second end, and a sun gear integrated to the torque shaft, at the first end, the second end has a first external diameter, and the sun gear having a second external diameter greater than the first external diameter; and a reference tube having a fixed end secured to the torque shaft adjacent the second end, a free end having a second reference feature adjacent the first reference feature, the reference tube extending around the torque shaft, the reference tube having an internal diameter, the internal diameter being between the first external diameter and the second external diameter.
F16H 1/28 - Transmissions à engrenages pour transmettre un mouvement rotatif avec engrenages à mouvement orbital
F16H 57/04 - Caractéristiques relatives à la lubrification ou au refroidissement
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F16D 3/06 - Accouplements extensibles, c. à d. avec moyens permettant le mouvement entre parties accouplées durant leur entraînement adaptés à des fonctions particulières spécialement adaptés pour permettre un déplacement axial
77.
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
A method of controlling a multi-engine aircraft includes receiving input for commanded thrust and modifying the commanded thrust using a model of an incumbent powerplant to generate a modified commanded thrust for matching aircraft performance with a new powerplant to the aircraft performance with the incumbent powerplant. The method includes applying the modified commanded thrust to the new powerplant.
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.
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.
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
B64D 37/00 - Aménagements relatifs à l'alimentation des groupes moteurs en carburant
85.
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 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
86.
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.
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
89.
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
90.
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.
A fuel system can include a first fuel circuit, a second fuel circuit, and an inert gas purge system operatively connected to both the first fuel circuit and the second fuel circuit to purge at least a portion of either or both of the first and/or second fuel circuit. The first fuel can be a liquid fuel and the second fuel can be a gaseous fuel. The first fuel circuit can include a first fuel manifold configured to fluidly communicate a first fuel supply with at least one dual fuel nozzles downstream of the first fuel manifold.
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 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
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
94.
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.
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 7/042 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction à géométrie variable
95.
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.
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p.ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F02C 9/18 - Commande du débit du fluide de travail par prélèvement, par bipasse ou par action sur des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre leurs étages
96.
AIRCRAFT INTAKE DUCT WITH ACTIVELY MOVABLE FLOW RESTRICTOR
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
97.
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
98.
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.
99.
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.
