37 - Services de construction; extraction minière; installation et réparation
Produits et services
Propulsion systems for aeronautical vehicles, as well as
their components and spare parts, including engines,
turbines, pumps, thrusters, reactors, nacelles, thrust
reversers, engine air intake, blower casing, exhaust cones;
compressors; heat exchangers (parts of machines);
transmission and propulsion couplings and components. Repair, overhaul, servicing and maintenance of propulsion
systems for aeronautical vehicles as well as their component
parts and spare parts.
2.
METHOD FOR PRODUCING A PREFORM FOR PART OF A BLADE OR PROPELLER BY WINDING A WEAVE OBTAINED BY WEAVING TO SHAPE
A method for manufacturing a fibrous preform for a blade or propeller part of a turbomachine, includes at least one fixing base extended by a mounting portion of an aerodynamic profile, the method including the winding of a fibrous texture, obtained by contour weaving, on a substrate of changing section having at least a first region of extra thickness in the shape of the fixing base and a second region in the shape of the mounting portion of the aerodynamic profile.
B29C 70/22 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins deux directions formant une structure bidimensionnelle
B29B 11/16 - Fabrication de préformes caractérisées par la structure ou la composition comprenant des charges ou des agents de renforcement
B29C 53/56 - Enroulement et assemblage, p.ex. enroulement en spirale
B29C 70/32 - Façonnage par empilage, c.à d. application de fibres, de bandes ou de feuilles larges sur un moule, un gabarit ou un noyau; Façonnage par pistolage, c.à d. pulvérisation de fibres sur un moule, un gabarit ou un noyau sur un moule, un gabarit ou un noyau rotatifs
B29L 31/08 - Pales pour rotors, stators, ventilateurs, turbines ou dispositifs analogues, p.ex. hélices
3.
METHOD AND SYSTEM FOR DETERMINING A FAULTY COMPONENT OF AN AIRCRAFT
The invention relates to a digital system that makes it possible to determine at least one faulty component of an aircraft. The system comprises: - a first module (MDS1) for determining at least a first symptom (S1ikk) collected in the vicinity of the aircraft; - a second module (MDS2) for determining at least a second symptom (S2ii) of the aircraft on the basis of degradation data (DCk) representative of an abnormal variation in flight data; - a module (MDPD) for determining probabilities of failure in a plurality of components of the aircraft on the basis of the at least first and second symptoms, this module comprising a Bayesian network (BR) trained using a knowledge base (BC) comprising data produced by a natural language processing module (MNLP) configured to process aircraft maintenance reports; - a module (MSAC) configured to flag at least one faulty component from among the plurality of components according to the probabilities.
The invention relates to a turbomachine (1), in particular an aircraft turbomachine, comprising: - a splitter nose (19) arranged in a primary duct (13) in which a primary flow circulates and intended to separate the primary flow into a radially inner flow (F11) and into a radially outer flow (F12); - stator equipment (26) installed in a channel (22) in which the radially outer flow (F12) circulates; and - first stator vanes (25) installed in the channel and intended to straighten the radially outer flow (F12) towards the equipment. According to the invention, the turbomachine comprises a diffuser (29) which is arranged downstream of the first stator vanes (25) and which is divergent between an inlet section and an outlet section, and second stator vanes (31) arranged downstream of the diffuser in order to deflect the radially outer flow at the outlet of the diffuser towards the equipment, the second stator vanes being situated upstream of the equipment.
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
F02K 3/02 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion
F02K 3/04 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
F02K 3/072 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comportant des rotors contra-rotatifs
F02K 3/077 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux l'ensemble fonctionnel étant du type multi-flux, c. à d. ayant au moins trois flux
F02K 3/08 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant avec réchauffage supplémentaire du fluide de travail; Leur commande
The invention relates to a turbine ring assembly (2) provided with a CMC sectorized turbine ring (4) and with a ring support structure (6), each sector (10) comprising a base (12) from which an upstream attachment tab (16) and a downstream attachment tab (14), which are axially spaced, extend radially outwards, the support structure (6) comprising an upstream radial flange (62) and a downstream radial flange (64) between which the attachment tabs (14, 16) are held. The assembly (2) further comprises, at each junction between two adjacent ring sectors (10) in the circumferential direction (Dc), a first cylindrical sealing element (42) extending mainly in the axial direction (DA) and arranged between the bases (12) of the two adjacent ring sectors (10), a second sealing element (124) extending radially along the downstream attachment tab (14) from the cylindrical seal (42), forming a sealed connection, and a third sealing element (126) extending radially along the downstream attachment tab (16) from the cylindrical seal (42), forming a sealed connection.
The invention relates to a method for weaving a texture (200) between warp yarns (210) and weft yarns (220) by means of a weaving loom (100) comprising at least one lance (130) which is able to pass through the loom (100) along a reference path (T130) and to draw a weft yarn (220), the method comprises at least displacing at least part of the warp yarns intended to exit from the texture (200) so that the yarns are arranged on one side of the reference path (T130) and the rest of the non-exiting warp yarns are arranged on another side of the reference path (T130), followed by the lance (130) passing through the loom (100) along the reference path (T130) without drawing any weft yarn.
The invention relates to a module (100) for a turbine engine, in particular for an aircraft turbine engine, comprising a shaft extending in an axial direction and an electric machine (110) comprising: - a rotor (120) rotatably coupled to the shaft of the module, the rotor comprising a disc (122) and magnetic elements (124) arranged regularly around the periphery of the disc; and - a stator (130) rigidly attached to a casing of the module comprising a ring and coils (134) distributed annularly inside the ring of the stator, the coils being arranged outside the magnetic elements of the rotor in a radial direction, characterised in that each coil is radially movable between a first end position and a second end position, each coil being held in the first position in normal operation and moved toward the second end position in the event of the rotor moving radially towards the coil of the stator.
Method for the automated weaving of a woven structure (30) by means of a weaving machine (1), comprising the insertion of a weft thread during the opening of a shed of warp threads, and prior steps of optimization, by an optimization tool (20), of an objective function having a first member evaluating the balance of said warp threads during the shed opening, said balance of the warp threads being evaluated by differences in elongation between the warp threads of an upper opening (10a) of said shed, and of a lower opening (10b) of said shed, and a second member evaluating constraints specific to the weaving machine (1). This optimization comprises the determination of a set of opening parameters that, at least locally, minimize said objective function, and this set of opening parameters is used to set the parameters of the weaving machine (1).
The invention relates to a turbine ring assembly (2) provided with a sectorized CMC turbine ring (4) and with a ring support structure (6), each sector (10) of the ring (4) comprising a base (12) from which an upstream attachment tab (16) and a downstream attachment tab (14), which are spaced axially, extend radially outwards, the support structure (6) comprising an upstream radial flange (62) and a downstream radial flange (64), between which the attachment tabs (14, 16) of each sector (10) are held, and the assembly (2) comprising, for each sector (10), at least one first pin (50) passing through the downstream attachment tab (14) and the downstream radial flange (64), and at least one second pin (40) passing through the upstream attachment tab (16) and the upstream radial flange (62). The ring assembly (2) further comprises an annular insert plate (66) mounted between the downstream attachment tab (14) of the ring sector (10) and the downstream radial flange (64) of the ring support structure (6).
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
The invention relates to a cooling-air injection casing (100) comprising a casing upstream end (30a), a casing downstream end (30b) and a casing main wall (30), the cooling-air injection casing further comprising: - an air mixing cavity (41); - an air intake cavity (42); - an air bleeding cavity (44); - a first sealing gasket (51); - and a second sealing gasket (52); the injection casing further having an air passage cavity (43) bounded by a substantially radial wall (33), the casing main wall (30) and the substantially axial wall (32), wherein the air passage cavity (43) is in fluid communication with the air mixing cavity (41) via openings (62) made in the substantially axial wall (32), and wherein the air passage cavity (43) is in fluid communication with the air bleeding cavity (44) by means of openings made in the substantially radial wall (63).
AA), the sealing gasket extends circumferentially around the axis A and the outer surface (51, 52) is arranged to face the sealing gasket, the sealing assembly being characterised in that it further comprises an air straightening member (200) comprising at least one row of radially extending vanes, the air straightening member being located upstream of the sealing gasket and the air straightening member being configured to straighten the air that passes through the sealing gasket.
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
The invention relates to a gasket configured to ensure a predefined clearance (j) between the gasket and an outer surface (500) and comprising a plurality of gasket sectors distributed circumferentially around the axis A, each gasket sector comprising an inner ring sector (13) connected to an outer ring sector (11) by a return member (12), the gasket being characterised in that the radially inner surface (Sint) of each inner ring sector comprises at least one pattern (31, 32) hollowed out from the radially inner surface of the inner ring sectors, the inner ring sector further comprising at least one channel (501) connecting a first opening (502) which opens onto the upstream face of the inner ring sector and a second opening (503) which opens into one of the patterns hollowed out from the radially inner surface of the inner ring sectors.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
13.
BLADED WHEEL FOR AN AIRCRAFT TURBOMACHINE, AND AIRCRAFT TURBOMACHINE
The invention relates to a bladed wheel (20) for an aircraft turbomachine, comprising a plurality of blades (26) with a shroud (32) extending circumferentially on either side of the blade tip (46) respectively in a first shroud portion (32a) and a second shroud portion (32b). Between a first blade and an adjacent second blade of the wheel, the latter comprises a member (50) for reconstituting a gas flow path, which member is arranged in a first material recess (54a) provided on the radially inner surface (36a) of the first shroud portion (32a) of the first blade, and in a second material recess (54b) provided on the radially inner surface (36a) of the second shroud portion (32b) of the second blade, the first and second recesses each having a bottom surface (56) intended to cooperate by friction with a radially outer friction surface (58a) of the member (50).
The invention relates to an assembly (100) for an aircraft turbine engine (1) having a longitudinal axis (X), the assembly (100) comprising: - a cylindrical rotor shroud (11) intended to be rotated about the longitudinal axis (X), - a nozzle (20) which has a ring of stator blades comprising a root (403) at the radially inner nozzle end supporting a seal (22), the seal being annular and configured to sealingly cooperate without contact with the cylindrical shroud (11) that is located radially below the nozzle, wherein the cylindrical shroud (11) comprises an annular layer (12) which is located radially across from the seal (22) and is made of a first material having a hardness greater than a hardness of a material of a radially inner end (23) of the seal (22) radially across from the annular layer (12).
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
The invention relates to an assembly for an aircraft turbine engine having a longitudinal axis (X) comprising a turbine nozzle that has a ring of stator vanes comprising a root (403) at the radially inner end of the turbine nozzle, the root carrying a seal (501), the seal (51) being a ring seal configured to co-operate as a non-contact seal with a cylindrical shroud of a rotor of the turbine engine arranged radially below the turbine nozzle, the seal (501) comprising a plurality of seal sectors distributed circumferentially around the longitudinal axis (X), each seal sector comprising a radially outer annular wall sector (502) and a radially inner annular wall sector (503) connected to one another by an elastically deformable member (504), wherein each radially outer annular wall sector (502) is attached to a seal support (505), the seal support (505) being a ring seal support rigidly attached to the root (403) of the turbine nozzle, and each radially outer annular wall sector (502) carries a coupling member (506) engaged circumferentially and retained radially in a circumferential groove of the seal support (505).
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
The invention relates to a seal (51) for an aircraft turbine engine, comprising a plurality of seal sectors that are circumferentially distributed about a longitudinal axis (X), each seal sector having a radially inner annular wall sector (56) and a radially outer annular wall sector (55) which are connected to each other by an elastically deformable member (57), an annular flange (75) being applied to an upstream face (74) of the radially outer annular wall sectors (55) so as to axially cover each radially outer annular wall sector (55) and each elastically deformable member (57), and the annular flange (75) being supported by the radially outer annular wall sectors (55) and being arranged with an annular clearance relative to the radially inner annular wall sectors (56).
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
The invention relates to a seal (100) for an aircraft turbomachine, comprising a plurality of seal sectors distributed circumferentially about a longitudinal axis (X), each seal sector (102a, 102b) comprising a radially outer annular wall sector (106a, 106b) and a radially inner annular wall sector (110a, 110b), which are connected to each other by an elastically deformable member (105a, 105b), wherein the circumferentially adjacent pairs of seal sectors (102a, 102b) have their respective elastically deformable members formed monolithically as a common deformable elastic member (105), the common deformable elastic member (105) connecting together two radially inner annular wall sectors (110a, 110b) of circumferentially adjacent seal sectors, and wherein the radially outer annular wall sectors (106a, 106b) of the seal sectors form a monolithic outer shroud.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
F01D 11/04 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe utilisant un fluide d'obturation, p.ex. de la vapeur
F01D 11/16 - Régulation ou commande du jeu d'extrémité des aubes, c.à d. de la distance entre les extrémités d'aubes du rotor et le corps du stator par des moyens auto-réglables
The invention relates to a metal powder for an additive manufacturing process, the metal powder comprising an alloy comprising by weight between 20% and 24% chromium, between 20% and 24% nickel, between 13.00% and 16.00% tungsten, between 0.02% and 0.12% lanthanum, between 0.05% and 0.15% carbon, between 0.20% and 0.50% silicon, at most 1.25% manganese, at most 3.00% iron, at most 0.015% sulfur, at most 0.020% phosphorus, at most 0.0001% bismuth, at most 0.0010% silver, at most 0.0010% lead, at most 0.015% boron, at most 0.0250% oxygen, at most 0.0200% nitrogen and less than 0.050% other elements in total, the balance being cobalt.
C22C 1/04 - Fabrication des alliages non ferreux par métallurgie des poudres
B22F 1/052 - Poudres métalliques caractérisées par la dimension ou la surface spécifique des particules caractérisées par un mélange de particules de dimensions différentes ou par la distribution granulométrique des particules
B22F 9/08 - Fabrication des poudres métalliques ou de leurs suspensions; Appareils ou dispositifs spécialement adaptés à cet effet par des procédés physiques à partir d'un matériau liquide par coulée, p.ex. à travers de petits orifices ou dans l'eau, par atomisation ou pulvérisation
B22F 10/28 - Fusion sur lit de poudre, p.ex. fusion sélective par laser [FSL] ou fusion par faisceau d’électrons [EBM]
B33Y 70/00 - Matériaux spécialement adaptés à la fabrication additive
B22F 10/366 - Paramètres de balayage, p.ex. distance d’éclosion ou stratégie de balayage
B22F 10/64 - Traitement de pièces ou d'articles après leur formation par des moyens thermiques
The invention relates to a metal powder for an additive manufacturing method, the metal powder comprising an alloy comprising, by weight, between 23% and 24.5% of chromium, between 9% and 11% of nickel, between 6.5% and 7.5% of tungsten, between 3% and 4% of tantalum, between 0.55% and 0.65% of carbon, between 0.3% and 0.5% of zirconium, between 0.15% and 0.25% of titanium, at most 2% of iron, at most 0.3% of silicon, at most 0.1% of manganese, at most 0.1% of copper, at most 0.015% of sulfur, at most 0.015% of phosphorus, at most 0.01% of boron, at most 0.025% of oxygen, at most 0.020% of nitrogen and at most 0.010% of hydrogen and less than 0.050% of other elements in total, the remainder being cobalt.
C22C 1/04 - Fabrication des alliages non ferreux par métallurgie des poudres
B22F 1/05 - Poudres métalliques caractérisées par la dimension ou la surface spécifique des particules
B22F 9/08 - Fabrication des poudres métalliques ou de leurs suspensions; Appareils ou dispositifs spécialement adaptés à cet effet par des procédés physiques à partir d'un matériau liquide par coulée, p.ex. à travers de petits orifices ou dans l'eau, par atomisation ou pulvérisation
B33Y 40/20 - Posttraitement, p.ex. durcissement, revêtement ou polissage
B33Y 70/00 - Matériaux spécialement adaptés à la fabrication additive
C22C 19/07 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de cobalt
B22F 10/28 - Fusion sur lit de poudre, p.ex. fusion sélective par laser [FSL] ou fusion par faisceau d’électrons [EBM]
B22F 10/36 - Commande ou régulation des opérations des paramètres du faisceau d’énergie
B22F 10/64 - Traitement de pièces ou d'articles après leur formation par des moyens thermiques
B33Y 50/02 - Acquisition ou traitement de données pour la fabrication additive pour la commande ou la régulation de procédés de fabrication additive
B33Y 80/00 - Produits obtenus par fabrication additive
C22F 1/10 - Modification de la structure physique des métaux ou alliages non ferreux par traitement thermique ou par travail à chaud ou à froid du nickel ou du cobalt ou de leurs alliages
B22F 10/366 - Paramètres de balayage, p.ex. distance d’éclosion ou stratégie de balayage
B22F 5/00 - Fabrication de pièces ou d'objets à partir de poudres métalliques caractérisée par la forme particulière du produit à réaliser
20.
FUEL SYSTEM AND METHOD FOR SUPPLYING A COMBUSTION CHAMBER IN AN AIRCRAFT TURBOSHAFT ENGINE WITH FUEL
A system for supplying the combustion chamber in an aircraft turboshaft engine with fuel, comprising injectors comprising a first pilot stage, a second pilot stage, and a main stage. The pilot stages being configured to deliver a permanent flow rate of fuel and the main stage being configured to deliver an intermittent flow rate of fuel. A device for distributing (2) fuel into the injectors comprising a primary circuit, a secondary circuit, and a valve configured to distribute the fuel between the primary circuit and the secondary circuit. The primary circuit being connected to the first pilot stage and the main stage, the secondary circuit being connected to the second pilot stage, and a member for controlling the valve.
The invention relates to a device comprising a fixed actuator for a control system for the orientation of fan blades of a turbomachine, comprising: at least one actuator body housed inside an annular casing of the turbomachine forming a bearing support, the annular casing being centred on a longitudinal axis of the turbomachine an inclined relative thereto, the actuator body having an internal volume in which a piston is able to slide, separating the internal volume into two chambers isolated from one another; and a plurality of actuator rods passing through the annular casing forming a bearing support and each having one end fixed to the piston and an opposite end intended to be connected to an internal ring of a movement transfer bearing in order to drive it in translation, the internal volume of the actuator body being closed downstream by a chamber bottom which forms a single part with the annular casing forming a bearing support.
An assembly for a propeller of a turbine engine of an aircraft is provided. The assembly includes a variable-pitch vane having a blade connected to a root, the vane having a longitudinal axis aligned with a vane pitch axis which passes through the root, a base rigidly connected to the vane to rotate together with the vane about the axis and connected to a portion of a toothed wheel extending around the axis, and a counterweight device having a shaft which can rotate about an axis substantially perpendicular to the axis, the shaft being connected to at least one flyweight and to a pinion meshed with the toothed wheel portion. The toothed wheel portion can be attached to the base by at least one shear pin.
An aeronautical thruster having a longitudinal axis and having a hub, an upstream annular row and a downstream annular row which are spaced apart from each another along the longitudinal axis, the upstream annular row being rotatable around the longitudinal axis, and the downstream annular row having a first blade and a second blade each extending in a radial direction to define a radial dimension between the hub and a radially outer end of the corresponding blade. When observed angularly around the longitudinal axis, the second blade can be positioned closer to an angular position at 3 o'clock or an angular position at 9 o'clock than the first blade, and the radial dimension of the first blade can be greater than the radial dimension of the second blade.
The invention relates to a ball-and-socket connection system (801) for actuating a movable part of an aircraft turbine engine. The system (801) comprises a first part (803) comprising an opening (807) and holes (809), a shaft (811) inserted into the holes (809) and held in the holes (809) by attachment means (813) and bushings (815), a second part (805), and a ball (819) rigidly attached to the second part (805) for rotating the second part (805) about the shaft (811). Moreover, the shaft (811), the attachment means (813) and the bushings (815) are configured so that, when the system (801) is assembled, the ball (819) is clamped at its ends in the direction of the shaft (811) such that the ball (819) is prevented from translating along the shaft (811) and a clearance (821) allows the shaft (811) to translate in its direction.
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
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE BORDEAUX (France)
INSTITUT POLYTECHNIQUE DE BORDEAUX (France)
Inventeur(s)
Cavarroc, Marjorie, Christine
Mauvy, Fabrice, Claude, Michel
Poulon, Angélique, Nadine, Jeanne
Abrégé
Provided is an ammonia manufacturing method in which an element present in the reactive medium is coated with a protective coating of TaN tantalum nitride having a hexagonal crystallographic structure or doped TaN tantalum nitride having a hexagonal crystallographic structure.
C25B 11/04 - PROCÉDÉS ÉLECTROLYTIQUES OU ÉLECTROPHORÉTIQUES POUR LA PRODUCTION DE COMPOSÉS ORGANIQUES OU MINÉRAUX, OU DE NON-MÉTAUX; APPAREILLAGES À CET EFFET Électrodes; Leur fabrication non prévue ailleurs caractérisées par le matériau
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE BORDEAUX (France)
INSTITUT POLYTECHNIQUE DE BORDEAUX (France)
Inventeur(s)
Cavarroc, Marjorie, Christine
Mauvy, Fabrice, Claude, Michel
Poulon, Angélique, Nadine, Jeanne
Abrégé
Provided is an electrode material comprising a metal substrate (11) coated by a protective coating (12) of TaN tantalum nitride having a hexagonal crystallographic structure or doped hexagonal TaN tantalum nitride, wherein the protective coating is a semiconductor and the band gap thereof has a width of at most 50 meV.
C25B 11/052 - PROCÉDÉS ÉLECTROLYTIQUES OU ÉLECTROPHORÉTIQUES POUR LA PRODUCTION DE COMPOSÉS ORGANIQUES OU MINÉRAUX, OU DE NON-MÉTAUX; APPAREILLAGES À CET EFFET Électrodes; Leur fabrication non prévue ailleurs caractérisées par le matériau Électrodes comportant des électro-catalyseurs sur un substrat ou un support Électrodes comportant un substrat et un ou plusieurs revêtements électro-catalytiques
C25B 11/067 - Composé inorganique, p.ex. oxyde d'indium-étain, silice ou oxydes de titane
C25B 11/075 - PROCÉDÉS ÉLECTROLYTIQUES OU ÉLECTROPHORÉTIQUES POUR LA PRODUCTION DE COMPOSÉS ORGANIQUES OU MINÉRAUX, OU DE NON-MÉTAUX; APPAREILLAGES À CET EFFET Électrodes; Leur fabrication non prévue ailleurs caractérisées par le matériau Électrodes comportant des électro-catalyseurs sur un substrat ou un support caractérisées par le matériau électro-catalytique formé d’un seul élément catalytique ou composé catalytique
The invention relates to a seal (100) which is configured to ensure a predefined clearance (j) between the seal and an outer surface (500) of a rotor which is rotatably mounted about an axis A arranged facing the seal, the seal extending circumferentially about the axis A and comprising a plurality of seal segments which are distributed circumferentially about the axis A, each seal segment comprising an inner ring segment (13) which is connected to an outer ring segment (11) by a return member (12), the seal being characterised in that the inner surface (Sint) of each inner ring segment comprises at least one row of patterns which are carved into the inner surface of the inner ring segments, each of the patterns having an elongate shape extending in an oblique direction with respect to the axial direction and being separated from another pattern by an uncarved portion (41, 42, 43) of the inner surface.
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
The invention relates to a gasket (100) comprising a plurality of gasket sectors, each comprising an inner ring sector (13) connected to an outer ring sector (11) by a return member (12), the return member comprising at least one return arm (200) comprising: - a radially extending outer attachment portion (201) connected to the outer ring sector; - a first circumferentially extending branch (203) connected to the outer attachment portion by an outer bend (202); - a second circumferentially extending branch (205) connected to the first branch by an intermediate bend (204); - a radially extending inner attachment portion (207) connected to the inner ring sector and connected to the second branch by an inner bend (206), and wherein the inner bend is preferentially arranged circumferentially between the outer bend and the intermediate bend.
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE SUPERIEURE DE PHYSIQUE ET DE CHIMIE INDUSTRIELLE DE LA VILLE DE PARIS (France)
SORBONNE UNIVERSITE (France)
Inventeur(s)
Carreira Rufato, Raul
Bussy, Emmanuel
Lebeyf, Thierry Michel André
Oussar, Yacine
Ditchi, Thierry
Abrégé
arcIsourcearcIsourcearcarc and the constants of the knowledge model at output; the method comprising continuously detecting the presence of an arc using a decision function depending on a cost function used in the learning phase.
G01R 31/12 - Test de la rigidité diélectrique ou de la tension disruptive
G06F 30/27 - Optimisation, vérification ou simulation de l’objet conçu utilisant l’apprentissage automatique, p.ex. l’intelligence artificielle, les réseaux neuronaux, les machines à support de vecteur [MSV] ou l’apprentissage d’un modèle
G06N 3/044 - Réseaux récurrents, p.ex. réseaux de Hopfield
G01R 19/25 - Dispositions pour procéder aux mesures de courant ou de tension ou pour en indiquer l'existence ou le signe utilisant une méthode de mesure numérique
G01R 31/00 - Dispositions pour tester les propriétés électriques; Dispositions pour la localisation des pannes électriques; Dispositions pour tests électriques caractérisées par ce qui est testé, non prévues ailleurs
The invention relates to a sealing gasket for an aircraft turbomachine comprising a plurality of sealing gasket sectors (20) distributed circumferentially around a longitudinal axis (X), each sealing gasket sector (20) comprising a radially inner annular wall sector (22) and a radially outer annular wall sector (24) connected to one another by an elastically deformable member (26), wherein each radially inner annular wall sector (22) is connected to an inner radial tab (28) and each radially outer annular wall sector (24) is connected to an outer radial tab (24), the inner tab (28) and the outer tab (30) being connected to one another by the elastically deformable member (26) which is formed of a stack of a plurality of contiguous elastic strips (32).
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/04 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe utilisant un fluide d'obturation, p.ex. de la vapeur
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
F01D 11/16 - Régulation ou commande du jeu d'extrémité des aubes, c.à d. de la distance entre les extrémités d'aubes du rotor et le corps du stator par des moyens auto-réglables
The invention relates to an assembly for an aircraft turbine engine having a longitudinal axis (X), comprising a nozzle which has a ring of stator blades (10) including a root (403) at the radially inner end of the nozzle, which root has a seal (400), the seal being annular and configured to engage in a non-contact sealing manner with a cylindrical shell of a rotor of the turbine engine, which shell is arranged radially under the nozzle, the seal comprising a plurality of seal segments distributed circumferentially about the longitudinal axis (X), each seal segment comprising a radially outer annular wall segment (424) and a radially inner annular wall segment (420) connected to each other by an elastically deformable member (400), wherein each radially outer annular wall segment (502) is attached to a seal support (505).
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/04 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe utilisant un fluide d'obturation, p.ex. de la vapeur
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
F01D 11/16 - Régulation ou commande du jeu d'extrémité des aubes, c.à d. de la distance entre les extrémités d'aubes du rotor et le corps du stator par des moyens auto-réglables
The invention relates to a seal (600) for an aircraft turbine engine, comprising a plurality of seal sectors that are circumferentially distributed about a longitudinal axis (X), each seal sector having a radially outer annular wall sector and a radially inner annular wall sector (610) which are connected to each other by an elastically deformable member (203), characterized in that the radially outer annular wall sectors form a monolithic outer shroud and the radially inner annular wall sectors are disposed circumferentially end-to-end, in that each radially inner annular wall sector (610) includes a first circumferential edge (612) and an opposite second circumferential edge which each have a slit (614), and in that a tongue (616) is mounted in part at the junction of two facing circumferential edges.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F02C 7/28 - Agencement des dispositifs d'étanchéité
The invention relates to a seal (201) for an aircraft turbine engine, comprising a plurality of seal segments distributed circumferentially about the axis (X), each seal segment comprising a radially outer annular wall segment (204) and a radially inner annular wall segment (202) connected to each other by an elastically deformable member (203), wherein the elastically deformable member (203) comprises an inner tab (206), a radially inner end (208) of which is attached to the radially inner annular wall segment (202, 205), and an outer tab (207), a radially outer end (209) of which is attached to the radially outer annular wall segment (204), the radially outer end (210) of the inner tab (206) and the radially inner end (211) of the outer tab (207) being connected to each other by a connecting wall (215), the inner tab (206) and the outer tab (207) extending radially so that the radially inner end (211) of the outer tab (207) is arranged radially inside the radially outer end (210) of the inner tab (206).
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/04 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe utilisant un fluide d'obturation, p.ex. de la vapeur
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
F01D 11/16 - Régulation ou commande du jeu d'extrémité des aubes, c.à d. de la distance entre les extrémités d'aubes du rotor et le corps du stator par des moyens auto-réglables
The invention relates to an assembly for an aircraft turbine engine having a longitudinal axis (X), which aircraft turbine engine comprises a turbine nozzle which has a stator blade ring comprising a root (403) at the radially inner end of the turbine nozzle, the root bearing a seal (419), the seal (419) being annular and configured to engage in a non-contact sealed manner with a cylindrical shell of a rotor of the turbomachine which is arranged radially inside the turbine nozzle, the seal (419) comprising a plurality of seal sectors which are distributed circumferentially around the longitudinal axis (X), each seal sector comprising a radially outer annular wall sector (401) and a radially inner annular wall sector (420) which are connected to one another by an elastically deformable member (407), wherein the seal support (402) comprises radial notches into each of which are inserted radial tongues (416) of the radially outer annular wall sectors (401) of the seal (419), the upstream annular tab (404) and the downstream annular tab (405) forming a U.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe
F01D 11/04 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages par obturation non contact, p.ex. du type labyrinthe utilisant un fluide d'obturation, p.ex. de la vapeur
F02C 7/28 - Agencement des dispositifs d'étanchéité
The invention relates to a metal powder for a metal injection moulding method, the metal powder being formed of a cobalt-based alloy comprising: - between 23.00 wt% and 24.25 wt% chromium; - between 9.00 wt% and 11.00 wt% nickel; - between 6.50 wt% and 7.50 wt% tungsten; - between 3.00 wt% and 4.00 wt% tantalum; - between 0.45 wt% and 0.60 wt% carbon; - between 0.30 wt% and 0.50 wt% zirconium; - between 0.15 wt% and 0.25 wt% titanium; - at most 2.00 wt% iron; - at most 0.30 wt% silicon; - at most 0.10 wt% manganese; - at most 0.10 wt% copper; - at most 0.015 wt% sulphur; - at most 0.015 wt% phosphorus; - at most 0.010 wt% boron; - at most 200 ppm oxygen; - at most 200 ppm nitrogen; - at most 100 ppm hydrogen.
B22F 1/05 - Poudres métalliques caractérisées par la dimension ou la surface spécifique des particules
C22C 1/04 - Fabrication des alliages non ferreux par métallurgie des poudres
B22F 3/22 - 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 pour la fabrication de pièces par coulée en moule poreux ou absorbant, c. à d. par coulée d'une suspension de poudre métallique dans un moule poreux, d'une façon analogue au coulage de la barbotine
B22F 5/00 - Fabrication de pièces ou d'objets à partir de poudres métalliques caractérisée par la forme particulière du produit à réaliser
C22C 19/07 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de cobalt
An aeronautical thruster having a longitudinal axis and having a hub and at least two annular rows of unducted blades including an upstream annular row and a downstream annular row which are spaced apart from one another along the longitudinal axis, the upstream annular row being rotatable around the longitudinal axis, and the downstream annular row comprising a series of blades including a first blade and a second blade each extending in a radial direction from the hub to define a radial dimension between the hub and a radially outer end of the corresponding blade, wherein the radial dimension of the first blade is greater than the radial dimension of the second blade.
The invention relates to a turbine ring assembly (2) provided with a sectorized CMC turbine ring (4) and with a ring support structure (6), each sector (10) of the ring (4) comprising a base (12) from which an upstream attachment tab (16) and a downstream attachment tab (14), which are spaced axially, extend radially outwards, the support structure (6) comprising an upstream radial flange (62) and a downstream radial flange (64), between which the attachment tabs (14, 16) of each sector (10) are held, and the assembly (2) comprising, for each sector (10), at least one pin (50) and at least one passage (140) passing axially through an attachment tab (14, 16) over an axial length (Ll) of the passage (140). At least one pin (50) comprises a planar bearing surface (506) facing an inner surface (142) of the passage (140) through which the pin (50) passes, the bearing surface (506) facing the inner surface (142) extending axially over a length (L2) less than the axial length (Ll) of the passage (140).
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
38.
SYSTEM FOR COUPLING BETWEEN THE INPUT SHAFT AND THE SUN GEAR OF AN EPICYCLIC GEAR TRAIN
The invention relates to a coupling system (50) comprising an input shaft (54) extending along a longitudinal axis (X), an output shaft (56) and a reduction gear (52) coupling the input shaft (54) to the output shaft (56), the reduction gear (52) comprising an input pinion (58) driven in rotation by the input shaft (54) and at least partially surrounding the input shaft (54), wherein the coupling system (50) further comprises a primary coupling device (60) for coupling the input shaft (54) to the input pinion (58), configured to allow a longitudinal movement of the input shaft (54) relative to the input pinion (58), the primary coupling device (60) comprising two primary flexible coupling members (62a,b) arranged axially on either side of the input pinion, each primary flexible coupling member (62a,b) forming a connection between the input pinion (58) and the input shaft (54), said connection allowing longitudinal translations and rotations about radial axes perpendicular to the longitudinal axis.
F16D 3/72 - Accouplements extensibles, c. à d. avec moyens permettant le mouvement entre parties accouplées durant leur entraînement avec pièces d'accouplement reliées par un ou plusieurs organes intermédiaires avec des fixations aux pièces d'accouplement axialement espacées
F16D 3/79 - Accouplements extensibles, c. à d. avec moyens permettant le mouvement entre parties accouplées durant leur entraînement avec pièces d'accouplement reliées par un ou plusieurs organes intermédiaires ayant la forme d'un disque élastique ou d'une bague plate disposés perpendiculairement à l'axe des pièces d'accouplement, différents jeux de trous dans le disque ou la bague étant reliés à chaque pièce d'accouplement, p.ex. joints Hardy le disque ou la bague étant métalliques
The invention relates to an assembly for a turbine engine (1) consisting of an intermediate casing (10) that comprises a primary duct portion (91), a tertiary duct portion (93), an intermediate hub (11), an inter-duct shroud (13), an outer shroud (14), structural arms (12) connecting the intermediate hub (11) and the outer shroud (14), and an exhaust duct (95) extending between the primary duct (91) and the tertiary duct (93), and bypass valves (80) capable of drawing fluid from the primary duct portion (91) in order to direct it toward the exhaust duct (95). The intermediate casing (10) is formed as a single piece and has a cavity (20) located between the intermediate hub (11) and the inter-duct shroud (13) that opens onto the upstream face (150) of the intermediate casing (10) via an opening (21), wherein the bypass valves (80) are able to be inserted into the cavity (20) via the opening (21).
F02C 3/13 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur ayant des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre des étages de différents rotors
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
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux commande du rapport des débits des différents flux
F02K 3/077 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux l'ensemble fonctionnel étant du type multi-flux, c. à d. ayant au moins trois flux
40.
METHOD FOR MANUFACTURING A PREFORM FOR A TURBOMACHINE COMPONENT AND INSTALLATION FOR WEAVING SUCH A PREFORM
A method for manufacturing a preform for a turbomachine component with a weaving installation including at least a first delivery roll and a storage roll. The method includes the following steps: —feeding the weaving installation with weft and warp filaments, —weaving the weft and warp filaments together to form the preform, —driving the preform in a direction of travel towards the storage roll, applying a predetermined tension on the preform, —winding the preform onto and storing it on the storage roll. The method also includes a step of immobilizing the preform on the first delivery roll positioned upstream of the storage roll in the direction of travel of the preform.
B29C 70/48 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.à d. d'objets distincts utilisant des moules opposables, p.ex. pour déformer des préimprégnés [SMC] ou des "prepregs" avec une imprégnation des renforcements dans le moule fermé, p.ex. moulage par transfert de résine [RTM]
B29B 11/16 - Fabrication de préformes caractérisées par la structure ou la composition comprenant des charges ou des agents de renforcement
The invention relates to an aeronautical propulsion system comprising: - a ducted fan section (2) comprising a fan (22), the fan having a fan rotor (9) comprising blades (14); - a gas generator (3) comprising a drive shaft (10, 11) configured to directly or indirectly rotate the fan rotor (9), wherein: a. 0 ≤< P.D.10-4≤ 160 and b. 2.7. (P. D.10-4)0,6≤ 100.D2PD.D.10-4)0.6 where d is a diameter of the fan rotor in metres that is measured in a plane normal to an axis (X) of rotation of the fan rotor (9) at the intersection (24) between a tip edge (21) and a leading edge (23) of the blades of the fan rotor (9); P is a thrust generated by the fan (22) in Newtons (N) when the propulsion system is stationary at takeoff speed in a standard atmosphere and at sea level; and BPR is a bypass ratio of the flow passing through the propulsion system defined as a ratio of a mass flow rate of a secondary air flow (F2) flowing through the propulsion system around the gas generator (3) to a primary air flow (F1) flowing through the gas generator (3) measured when the propulsion system (1) is stationary and inactive at a speed corresponding to a takeoff speed of an aircraft comprising the aeronautical propulsion system in a standard atmosphere and at sea level.
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
The invention relates to a guide device (70) for an aircraft turbine engine, the guide device comprising: - a metal body (72) provided with at least one attachment tab (76) comprising an opening (74) configured to receive a screw or the like; - a guide pad (74) made of a plastic or composite material, this pad (74) provided with a first surface (74b) bonded to a first surface (72a) of the body (72) and a second opposing surface (74a) which is free and which is configured to form a sliding guide surface, characterised in that the metal body (72) comprises through-holes (84) which open onto the first surface (72a) of the body (72) and in that the guide pad (74) comprises studs (86) which project from the first surface (84b) of the pad (84) and which are respectively engaged and crimped in the holes (84) of the body (72).
F02C 3/13 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur ayant des raccordements variables du fluide de travail entre des turbines ou des compresseurs ou entre des étages de différents rotors
F02K 3/075 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux commande du rapport des débits des différents flux
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
The invention relates to a method (100) for controlling the hybridisation rate of a hybrid turbine engine, the control method (100) comprising at least: - a step of determining electric torque control saturation (102), in particular on the basis of a first static saturation law and a first dynamic saturation law, the dynamic saturation law being in particular defined on the basis of a target hybridisation rate; - a step of determining fuel-flow control saturation (103); - a step of determining electric torque control and fuel-flow control (105) using a corrector that takes into account at least - a difference between the engine speed to be reached and the current engine speed, - an electric torque control saturation, and - a fuel-flow control saturation, so as to obtain the target hybridisation rate.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 9/26 - Commande de l'alimentation en combustible
F02C 9/48 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel
F02C 9/56 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel avec la commande de la transmission de puissance
F02K 5/00 - Ensembles fonctionnels comportant un moteur, autre qu'une turbine à gaz, entraînant un compresseur ou un ventilateur soufflant
44.
METHOD FOR COMPENSATING THE ELECTRICAL POWER OF A HYBRID TURBINE ENGINE
The invention relates to a method (100) for compensating the electrical power of a hybrid turbine engine comprising an electrical network of a propulsion assembly connected to an external electrical network, which method comprises: o a step (101) of determining external electrical power; o a step of calculating the propulsion electrical power (102) during which additional power to be supplied to, or to be dissipated by, the electrical network of the propulsion assembly is calculated; o a step (103) of calculating compensation electrical power during which compensation electrical power needed to supply the electrical power required by the external electrical network during the transient state is calculated; and o a step (104) of drawing or injecting power during which the compensation electrical power is drawn or injected by an electrical compensation member so as to supply the electrical power required by the external electrical network.
The present application relates to an assembly for a turbine engine (1), comprising a first rotor (2) which is rotatable about a longitudinal axis (X-X) of the turbine engine (1), the first rotor (2) comprising a first arm (26); a second rotor (3) which is rotatable about the longitudinal axis (X-X) and comprises a second arm (36); a first sealing ring (4) which is centred on the longitudinal axis (X-X), is arranged radially outside the first arm (26) and comprises a first radial flange (40) fixedly mounted between the first arm (26) and the second arm (36); and a second sealing ring (5) which is separate from the first sealing ring (4), is centred on the longitudinal axis (X-X) and is arranged radially outside the second arm (36), the second sealing ring (5) comprising a first part (51) which is designed to come into contact with the second rotor (3), and a second part (52) which is separate from the first part (51) and is designed to come into contact with the first sealing ring (4).
A method for manufacturing a blade made of a composite material for a turbine engine, in particular a turbine engine of an aircraft, wherein a reinforcing shield is attached to one edge of an airfoil, the shield having a generally elongated shape along an elongation axis of the airfoil. The shield includes at least one excess dimension so as to be able to adjust the moment weight of the blade along the axis according to the position of the shield on the edge of the axis, and the method includes a step for removing a surplus portion of the shield so s to balance the blade.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE NORMALE SUPERIEURE PARIS- SACLAY (France)
SAFRAN (France)
SAFRAN AIRCRAFT ENGINES (France)
Inventeur(s)
Schneider-Die-Gross, Julien Paul
Mendoza Quispe, Arturo
Roux, Stéphane
Abrégé
A method for characterizing a mechanical part making it possible to evaluate the residual stresses in the part, as well as a method for constructing a predictive model and a non-destructive testing method making it possible to easily test such a part, the characterizing method including the following steps: measuring geometrical information of the part in a first state, physically transforming the part between the first state and a second state, measuring geometrical information of the part in its second state, determining the displacement field between the first state and the second state of the part by a digital image correlation method and obtaining the deformation field between the first state and the second state of the part, determining the stress field in the second state of the part by a finite element simulation method.
A tool for forming an abradable coating in an aircraft turbine engine module, this module having an annular stator casing and a central rotor which can move inside the casing about an axis, wherein it has a first portion which is attached to the rotor, a second portion which includes at least one first trans-verse arm extending radially outwards and carrying spreading rollers, and a flywheel centered on the axis and configured to be rotated manually by an operator in order to rotate the tool and the rotor within the casing, so that the spreading rollers are capable of spreading a resin in order to form the coating.
B05C 7/06 - Appareillages spécialement agencés pour appliquer un liquide ou autre matériau fluide à l'intérieur d'une pièce creuse par des dispositifs se déplaçant en restant en contact avec l'ouvrage
B05C 7/00 - Appareillages spécialement agencés pour appliquer un liquide ou autre matériau fluide à l'intérieur d'une pièce creuse
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
49.
HOUSING FOR AN AIRCRAFT TURBOMACHINE AND METHOD FOR HOUSING MANUFACTURE
A housing for an aircraft turbomachine, including an annular casing extending around an axis A and having an internal annular surface, the housing also including an annular abradable support cartridge that is fixed against the internal annular surface, the abradable support cartridge has a reinforced coating including a fibrous texture reinforcement embedded in a resin matrix, the fibrous texture reinforcement including a stack of fibrous texture plies, wherein the stack of plies includes at least one ply made of Kevlar® or of glass fibres.
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
F01D 21/04 - Arrêt des "machines" ou machines motrices, p.ex. dispositifs d'urgence; Dispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à une position incorrecte du rotor par rapport au stator, p.ex. indiquant cette position
50.
METHOD FOR BALANCING FAN VANES WITH TRAILING EDGE MACHINING
A method for manufacturing a blade of a fan for an aircraft turbomachine extending around a longitudinal axis, the blade including an aerofoil made of a composite material having a leading edge and a trailing edge connected by a pressure side and a suction side, the aerofoil extending between a base and a free end, the method including each of the following steps:—pairing and bonding a metal shield to the leading edge,—obtaining a first measurement of a radial moment weight of the blade before or after the metal shield has been bonded to the leading edge and determining a machining kerf,—machining the free end and/or the trailing edge along the machining kerf so as to adjust the mass and radial moment weight of the blade,—obtaining a second measurement of the radial moment weight of the blade.
Leaks in an oil lubrication circuit, in particular through a heat exchanger communicating with another circuit, are measured by the movements of a piston that is sensitive to variations in flow rates and pressure drops at two calibrated ports on either side of the predicted location of the leak. In one intended application, the second circuit contains overpressurised fuel which is used to constantly cool the oil.
F16N 29/00 - Dispositifs particuliers dans les installations ou systèmes de lubrification indiquant ou détectant des conditions indésirables; Utilisation des dispositifs sensibles à ces conditions dans les installations ou systèmes de lubrification
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F02C 7/16 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur
F28F 27/00 - Commandes ou dispositifs de sécurité spécialement adaptés pour les appareils d'échange ou de transfert de chaleur
52.
ASSEMBLY COMPRISING A VANE AND A VANE PITCH SETTING SYSTEM
An assembly for an aircraft turbine engine is provided. The assembly generally includes an un-ducted propeller blade of which the root has a bulbous portion; a system for setting the pitch of the blade about a pitch setting axis, the pitch setting system including a cup configured to be placed inside the outer casing, has a flared open top end for axial insertion of the root and in which the root of the blade can be axially affixed by a locking member; and at least one radial retention prong in a radial space reserved between the wall of the cup and the root.
A method for preparing a root of a blade for mounting a turbomachine blade, made of composite material, in a root support, the method including positioning, on a side wall of the root of the blade, an anti-wear layer including synthetic fibres coated or impregnated with a bonding agent; polymerising the bonding agent such that the anti-wear layer adheres to the side wall of the root of the blade; and machining a free outer surface of the anti-wear layer intended to be in contact with the root support, so as to obtain a precision-ground outer surface.
F01D 5/28 - Emploi de matériaux spécifiés; Mesures contre l'érosion ou la corrosion
F01D 5/00 - Aubes; Organes de support des aubes; Dispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
F01D 5/30 - Fixation des aubes au rotor; Pieds de pales
The invention relates to a method (P) for treating a component (1), in particular of a turbine engine, comprising the following steps: - (S1) providing a surface stress (F), intended to be applied to a contact surface (5) of the component (1), and at least one radius of curvature (R); - (S2) determining a volume stress exerted in the material of the component (1) below the surface, comprising calculating an analytical expression of the volume stress as a function of the surface stress and the radius of curvature; - (S3) for each treatment of a set of thermochemical treatments, determining a hardness and a residual stress of the component (1); - (S4) identifying in the set acceptable treatments for which the hardness and the residual stress correspond to a mechanical strength of the component (1) that is greater than a threshold strength taking into account the volume stress, and a treatment duration is less than or equal to a threshold duration; and - (S6) applying one of the acceptable treatments to the component (1).
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
G01M 99/00 - Matière non prévue dans les autres groupes de la présente sous-classe
G01N 3/40 - Recherche de la dureté ou de la dureté au rebondissement
G01N 3/56 - Recherche de la résistance à l'usure ou à l'abrasion
F01D 5/28 - Emploi de matériaux spécifiés; Mesures contre l'érosion ou la corrosion
F01D 5/30 - Fixation des aubes au rotor; Pieds de pales
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
F01D 21/14 - Arrêt des "machines" ou machines motrices, p.ex. dispositifs d'urgence; Dispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à d'autres conditions spécifiques
55.
PROCESS FOR MANUFACTURING A GAS TURBINE ENGINE BLADE
The invention relates to a process for manufacturing a turbomachine blade, wherein: —a part (4) is manufactured comprising a foot (4), a heel (6) and an air stream zone (10) extending between the foot and the heel, the air stream zone comprising at least one protuberance (20, 24, 26) projecting from a main face (12) of the zone, the manufacturing being performed by injecting a mixture comprising a binder and a powder, the powder comprising at least a metal or a ceramic: —debinding is performed on the part so as to eliminate a greater quantity of the binder from the part: —heat treatment is performed on the part; and—the or each protuberance is eliminated from the air stream zone.
An assembly for an aircraft turbomachine (20), comprising an intermediate casing hub (30), an equipment support (40) having an upstream face (42) located downstream of the intermediate casing hub (30), and a transfer gear box (50) driving a power transmission shaft (54) configured to transmit mechanical power, taken off an engine shaft (17) of the turbomachine, to at least one upstream equipment item (46) mounted on the upstream face (42) of the equipment support (40), the upstream equipment item (46) comprising a structural element (462) attaching the equipment support (40) to the intermediate casing hub (30).
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE POITIERS (France)
ECOLE NATIONALE SUPERIEURE D'ARTS ET METIERS (France)
Inventeur(s)
Rame, Jérémy
Cormier, Jonathan
Mataveli Suave, Lorena
Menou, Edern
Abrégé
The invention relates to a nickel-based superalloy comprising in weight percentages: 5.0 to 6.0% of aluminium, 6.0 to 9.0% of tantalum, 1.0 to 5.0% of cobalt, 5.0 to 7.5% of chromium, 0 to 1.0% of molybdenum, 4.0 to 6.0% of tungsten, 0.50 to 3.5% of rhenium, 1.0 to 4.0% of platinum, 0.05 to 0.25% of hafnium, 0 to 0.15% of silicon, the remainder consisting of nickel and unavoidable impurities. The invention also relates to a single-crystal blade comprising such an alloy and to a turbomachine comprising such a blade.
The invention relates to a navigation method of a vehicle (1) provided with a lidar (5) and an inertial measurement unit (6) both of which are connected to an electronic navigation unit (2) implementing a first simultaneous location and mapping algorithm and a second data fusion algorithm, the first algorithm receiving as input first location data originating from the lidar (5) in order to calculate a first movement from a start position, the second algorithm receiving as input the first movement and the second location data originating from the inertial measurement unit (6) and providing as output a second fused position of the vehicle (1), characterised in that the second algorithm implements invariant extended Kalman filtering and in that the second fused position is introduced into a feedback loop supplying the first algorithm as input in order to calculate a new position which becomes the start position.
G01C 21/16 - Navigation; Instruments de navigation non prévus dans les groupes en utilisant des mesures de la vitesse ou de l'accélération exécutées à bord de l'objet navigant; Navigation à l'estime en intégrant l'accélération ou la vitesse, c. à d. navigation par inertie
59.
HEAT EXCHANGE SYSTEM FOR AN AIRCRAFT TURBINE ENGINE
The invention relates to a heat exchange system (20), in particular for an aircraft turbine engine (1), the system comprising: - two annular walls which extend one around the other and about the same axis (X), and which define between them a flow duct (V2) for a gas flow (F2); - a first annular heat exchange device (21) which is carried by one of the walls (22), and which is located in the duct (V2), this first device (21) occupying at most 75%, or even 50%, of a height (H) of the duct (V2); and - a second annular heat exchange device (30) which is carried by the other of the walls (23), and which is located in the duct (V2), this second device (30) occupying at most 75%, or even 50%, of a height (H') of the duct (V2).
The invention relates to a heat exchange system (20), in particular for an aircraft turbine engine (1), the system comprising: - an outer annular wall (22) and an inner annular wall (23), respectively; - a first annular heat exchange device (21) which is carried by one of the walls (22), and which comprises a heat exchange matrix (26) positioned between this wall (22) and a cover (27), characterised in that it further comprises: - a second annular heat exchange device (30) which comprises a heat exchange matrix (31) which is axially spaced apart from the heat exchange matrix (26) of the first device (21) and which is positioned between the cover (27) of the first device (21) and the opposite wall (23) or another cover (32) of the second device (30).
The invention relates to an aircraft turbine engine (1), the turbine engine (1) having a longitudinal axis (X) and comprising: - a primary oil circuit (C1) for cooling a mechanical device; - a secondary oil circuit (C2) for cooling an electronic device (16); and - a heat exchange system (20) located in a duct (V2) to be swept by a gas flow (F2), this heat exchange system (20) comprising a first cooling circuit (C21) connected to the secondary oil circuit (C2) and a second cooling circuit (C11) connected to the primary oil circuit (C1).
The invention relates to a method for controlling an aircraft turbomachine, the turbomachine comprising a high-pressure spool and a low-pressure spool, the method comprising the following steps: a)- detecting an acceleration request, said acceleration request generating an increase in a fuel setpoint in the turbomachine and an excursion of the speed of the high-pressure spool relative to the low-pressure spool, b)- comparing the speed of the high-pressure spool and a first predetermined threshold (S1), and if the speed of the high-pressure spool reaches the first threshold, diverting torque from the high-pressure spool, c)- if the speed of the high-pressure spool reaches a second predetermined threshold greater than the first threshold, reducing the fuel setpoint, and d)- continuing to divert torque from the high-pressure spool in order to return the speed of the high-pressure spool to below the second threshold.
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 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
F02C 9/56 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel avec la commande de la transmission de puissance
63.
METHOD FOR THE NON-DESTRUCTIVE TESTING OF PARTS MADE OF COMPOSITE MATERIAL
The present invention relates to a method for the non-destructive testing of a woven composite part, a tomographic volume of which has previously been acquired and divided into visualisation zones, characterised in that it comprises the following steps: dividing the visualisation zones of the tomographic volume into three-dimensional sub-volumes; extracting a descriptor vector and projecting the extracted descriptor vector so as to condense the descriptor vector into a reduced descriptor vector; converting the reduced descriptor vector into a scalar value representing a likelihood of the sub-volume belonging to a healthy class, aggregating the scalar values in each of the visualisation zones and estimating the membership of the healthy class; and checking for the presence of defects in the woven composite part according to the estimates of the membership of the healthy class of the thus-obtained visualisation zones of the tomographic volume.
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
G01N 33/00 - Recherche ou analyse des matériaux par des méthodes spécifiques non couvertes par les groupes
A method for shaping a fiber blank including a warp tracer thread and at least one weft tracer thread, includes the projection of a warp visual reference onto the fiber blank corresponding to a reference location of the warp tracer thread, the deformation of the blank from the bottom of the airfoil blank to the top of the airfoil blank such as to make the warp tracer thread correspond with its visual reference, the projection of a weft visual reference on the fiber blank corresponding to a reference location of the weft tracer thread, and the deformation of the blank from the warp tracer thread to the first edge and to the second edge of the blank such as to make the weft tracer thread correspond with its visual reference.
B29C 70/54 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes opération de façonnage des matières composites comprenant uniquement des renforcements - Parties constitutives, détails ou accessoires; Opérations auxiliaires
B29C 70/48 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.à d. d'objets distincts utilisant des moules opposables, p.ex. pour déformer des préimprégnés [SMC] ou des "prepregs" avec une imprégnation des renforcements dans le moule fermé, p.ex. moulage par transfert de résine [RTM]
B29C 70/88 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts caractérisées principalement par des propriétés spécifiques, p.ex. électriquement conductrices ou renforcées localement
B29L 31/08 - Pales pour rotors, stators, ventilateurs, turbines ou dispositifs analogues, p.ex. hélices
A device for additive manufacturing by powder bed fusion includes an enclosure having, in a lower portion of the enclosure, a manufacturing platform and a powder spreading means adapted to travel over the manufacturing platform in order to spread powder previously deposited on the manufacturing platform. The enclosure further includes, in an upper portion, a window facing the manufacturing platform and allowing the passage of laser beams. The device has at least one effector arranged in the enclosure and connected to at least two actuators, which are configured to move the effector between the upper portion of the enclosure and a target position on the manufacturing platform.
B22F 12/60 - Dispositifs de planarisation; Dispositifs de compression
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
20a20b20b) of the rear bearing (20b) onto the axis of rotation (X), along an axis of application of the forces of the rear bearing (20b), in metres (m); R is an average radius of the fan bearings (20a, 20b), in metres (m); Dg is the diameter of the fan rotor (9) in metres (m); and A = - 1.27 m-1 and B = 4.4.
F02C 7/06 - Aménagement des paliers; Lubrification
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
67.
VARIABLE-PITCH VANE MADE OF COMPOSITE MATERIAL FOR AN UNDUCTED FAN OF AN AIRCRAFT
The present invention relates to a variable-pitch vane (100) made of a composite material for an unducted fan of an aircraft, the vane comprising: a fibrous reinforcement formed as a single piece of three-dimensional fabric comprising an airfoil portion (102) and a root portion (104) forming a retaining bulb and intended to be connected to a variable-pitch mechanism (20), the three-dimensional weaving being carried out continuously between the root portion and the airfoil portion, the fibrous reinforcement defining an internal shaping cavity (111) formed by a separation (D1) extending inside the root portion and the blade portion, wherein a matrix densifies the fibrous reinforcement and reinforcing metal shells (105) are arranged around the root portion.
B32B 5/02 - Produits stratifiés caractérisés par l'hétérogénéité ou la structure physique d'une des couches caractérisés par les caractéristiques de structure d'une couche comprenant des fibres ou des filaments
B29C 70/24 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des renforcements, p.ex. matières plastiques auto-renforçantes des renforcements fibreux uniquement caractérisées par la structure des renforcements fibreux utilisant des fibres de grande longueur, ou des fibres continues orientées dans au moins trois directions formant une structure tridimensionnelle
B29D 99/00 - Matière non prévue dans les autres groupes de la présente sous-classe
A cooling jacket for a hollow airfoil of a turbine nozzle of a turbomachine, includes a main body including a central intake duct central defining a first ventilation air circulation area and connected to suction and pressure faces including at least two rows of drill holes by two separating walls defining second and third ventilation air circulation areas, an outer plate including first, second and third holes to allow the ventilation air respectively into the first, second and third ventilation air circulation areas, and an inner plate including a central opening to expel air from the first ventilation air circulation area, the outer and inner plates being secured by respectively soldering to the main body to form a one-piece unit with three ventilation air circulation areas, independent and airtight with respect to one another, before its installation in the hollow airfoil of the nozzle.
A support assembly for supporting a main accessory gearbox of an aircraft turbine engine, the accessory gearbox including gears and supporting at least one item of equipment driven by the gears. The support assembly can include a structure for connecting and supporting the turbine engine to a pylon of the aircraft including: an intermediate axial portion for attachment to the pylon, having an upper end that defines a linking interface with the pylon, the intermediate axial portion bearing suspension rods, which are intended to be connected to the turbine engine; a front axial portion extending forward of the intermediate portion and having at least one suspension member configured to be connected to the turbine engine; and a rear axial portion extending rearwards of the intermediate portion and supporting the main accessory gearbox.
A fibrous preform for manufacturing an annular housing made of composite material for a turbine engine includes at least one layer that has a fibrous texture, has a three-dimensional or multilayer weave, and extends about a longitudinal axis. At least one mat includes a thermoplastic material filled with carbon nanotubes and extending about the axis. At least one multiaxial fibrous sheet extends about the axis. The at least one mat is inserted between the fibrous sheet and the at least one layer having a fibrous texture.
B29C 53/00 - Façonnage par cintrage, pliage, torsion, redressage ou aplatissage; Appareils à cet effet
B29C 53/56 - Enroulement et assemblage, p.ex. enroulement en spirale
B29C 53/80 - Façonnage par cintrage, pliage, torsion, redressage ou aplatissage; Appareils à cet effet - Eléments constitutifs, détails ou accessoires; Opérations auxiliaires
B29K 105/08 - Présentation, forme ou état de la matière moulée contenant des agents de renforcement, charges ou inserts de grande longueur, p.ex. ficelles, mèches, mats, tissus ou fils
An unducted rectifier for a turbomachine is provided. The rectifier includes: (1) stator vanes each having a root and a blade radially projecting from the root, the root having two half-platforms; (2) pivots, each associated with a vane mounted so as to pivot about a pitch axis and configured to be connected connect to a system for changing the pitch of the blades, each pivot comprising having an upper platform and a cylindrical lower portion; and (3) attachment systems for attaching the half-platforms of the root of the vane to the associated pivot, each having two plates having a substantially identical shape to each half-platform and a clamping system for clamping the half-platforms between the plates of the attachment system and the upper platform of the pivot.
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
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
72.
SEALING ARRANGEMENT FOR A ROLLING BEARING SUBJECTED TO CENTRIFUGAL FORCES
A sealing device composed of a lip seal and a spring or another resilient seal which radially and axially compresses the first seal, forming two sealing lines closed over substantially planar surfaces which are opposed in the direction of the protected rolling bearing axis. The device is small and lightweight and withstands the centrifugal forces in the lateral direction of the rolling bearing. It is possible to dispense with any further parts for retaining it. The space requirement is reduced, and assembly is easy. The sealing device is suitable for arrangements of variable-pitch propeller blade roots.
The invention relates to a female electrical connector configured to be connected to a male electrical connector, the female electrical connector including a female connection pin, an insulating member, and an electrically conductive peripheral shell, forming an overall recess configured to receive a male connection pin of the male electrical connector so as to allow: in a connected position, electrical conduction between a first female surface and a first male surface, with a third electrically conductive female surface being in contact with a second electrically non-conductive male surface so as to prevent any change to a reference electrical potential of the peripheral shell; and in a partially disconnected position, electrical conduction between the first female surface and the first male surface, with the third electrically conductive female surface being in contact with the first electrically conductive male surface so as to change the potential of the peripheral shell.
H01R 13/703 - Association structurelle avec des composants électriques incorporés avec interrupteur incorporé actionné par l'engagement ou le retrait des pièces de couplage
H01R 13/53 - Socles ou boîtiers pour dures conditions de service; Socles ou boîtiers avec des moyens pour éviter l'effet couronne ou l'amorçage d'un arc
H01H 83/10 - Interrupteurs de protection, p.ex. disjoncteur ou relais de protection actionné par des conditions électriques anormales autres que seulement les courants excessifs actionnés par une tension excessive, p.ex. pour la protection contre la foudre
H01R 4/60 - Connexions entre ou avec des conducteurs tubulaires
H01R 13/66 - Association structurelle avec des composants électriques incorporés
H01R 24/20 - Pièces de couplage portant des douilles, des pinces ou des contacts analogues, assujetties uniquement à un fil ou un câble
H01R 24/28 - Pièces de couplage portant des broches, des lames ou des contacts analogues, assujetties uniquement à un fil ou un câble
H02H 3/20 - Circuits de protection de sécurité pour déconnexion automatique due directement à un changement indésirable des conditions électriques normales de travail avec ou sans reconnexion sensibles à un excès de tension
74.
ROTOR ASSEMBLY WITH INDEPENDENT PLATFORMS AND METHOD FOR ASSEMBLING SUCH A ROTOR
The invention relates to a turbine rotor assembly of a turbine engine including: - a rotor disc centred on an X-axis and having recesses, - movable blades each comprising a root mounted in a recess, - a plurality of platforms for holding the movable blades in the recesses of the disc, and - a flange for holding the movable blades in the recesses, each platform including: - a duct wall defining a gas flow duct, - an L-shaped downstream hook with a vertical portion extending radially from the duct wall and a horizontal portion housed in a holding hook positioned at the top of a tooth of the disc, and - an upstream hook extending substantially radially from an upstream area of the duct wall and held axially between the flange and the disc.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE PAUL SABATIER - TOULOUSE III (France)
INSTITUT NATIONAL POLYTECHNIQUE DE TOULOUSE (France)
Inventeur(s)
Rousselle, Mélanie
Ansart, Florence
Estournes, Claude Gilbert Jean-Pierre
Fradet, Guillaume
Fregeac, Arnaud Louis Gabriel
Selezneff, Serge Georges Vladimir
Abrégé
A method for manufacturing an abradable layer and a substrate coated with this layer, may include: preparing a powder composition including at least ceramic particles and an inorganic filler having a lamellar crystallographic structure, the volume content of the inorganic filler in the powder composition being in a range of from 1 to 75%; compressing the powder composition; and sintering the powder composition thus compressed in order to obtain the abradable layer.
C04B 35/48 - Produits céramiques mis en forme, caractérisés par leur composition; Compositions céramiques; Traitement de poudres de composés inorganiques préalablement à la fabrication de produits céramiques à base d'oxydes à base d'oxydes de zirconium ou d'hafnium ou de zirconates ou d'hafnates
C04B 35/50 - Produits céramiques mis en forme, caractérisés par leur composition; Compositions céramiques; Traitement de poudres de composés inorganiques préalablement à la fabrication de produits céramiques à base de composés de terres rares
C04B 35/626 - Préparation ou traitement des poudres individuellement ou par fournées
An abradable coating includes a tubular cell structure, wherein the the tubular cell structure includes a fibrous reinforcement of discontinuous short fibers which is densified by a ceramic matrix.
C04B 35/657 - Procédés comportant une étape de fusion pour la fabrication de réfractaires
C04B 38/00 - Mortiers, béton, pierre artificielle ou articles de céramiques poreux; Leur préparation
F01D 11/12 - Prévention ou réduction des pertes internes du fluide énergétique, p.ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator utilisant un élément de friction allongé, p.ex. un élément d'usure, déformable ou contraint de façon élastique
77.
TURBINE BLADE OF A TURBINE ENGINE WITH SELF-GENERATED INTERLOCK CONTACT FORCE IN OPERATION
Blade of a turbomachine turbine including a airfoil extending along a first direction between an internal end and an external end, a base secured to said internal end, and a root secured to said external end and extending, along a second direction perpendicular to said first direction, between an upstream edge, and a downstream edge and, along a third direction orthogonal to the first and second directions, between first and second lateral edges, the first side edge having, in a plane including the second and third directions, a shape complementary to that of the second lateral edge. The first lateral edge includes a first portion extending from the upstream edge, and a second portion externally protruding along the third direction at an acute angle with the first portion.
The invention relates to a device for setting the pitch of a vane (1) for a turbomachine, the device comprising: - a vane that is pivotable in a first direction and comprises a leading edge, a trailing edge and a base (6) which has an opening (7); - a fin (3) that comprises a leading edge, a trailing edge and a portion (12) accommodated in the opening; - a control means that rotates the vane (4) and moves the portion of the fin; - so as to keep a variation in a C/D ratio below 10%, the ratio being between a projection in a direction of projection normal to the first direction of a: - length C between the leading edge and the leading edge of the fin (3); and - length D between the leading edge and the trailing edge of the vane.
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
The invention relates to a pitch-change mechanism (70) comprising a frame (72), a moving part (102) which is translatably moveable along a longitudinal axis (X), and a locking device (160) for immobilising the moving part (102) relative to the frame (72). The locking device (160) comprises a blocking surface (162), a guide surface (164) and a blocking member (166) interposed therebetween. It further comprises a return member (200) which urges the guide surface (164) and the blocking member (166) toward one configuration of one relative to the other and a holding device (202) for holding the guide surface (164) and the blocking member (166) in another configuration under certain predetermined conditions.
The invention relates to a pitch-change mechanism (70) comprising a frame (72), a movable part (102) that is translatable along a longitudinal axis (X), and a locking device (160) for immobilising the movable part (102) with respect to the frame (72). The locking device (160) comprises a movable member (210) that is translatable with respect to a locking member (164) between a retracted position in which it leaves the locking member (164) free to be in an unlocked configuration away from a surface (162) and a deployed position in which it forces the locking member (164) into a locked configuration engaged with the surface (162). It also comprises a return member (220) which urges the movable member (210) toward its deployed position and a holding device (222) for holding the movable member (210) in its retracted position.
A splitter between a primary flow and a secondary flow of a dual flow turbomachine includes a single-piece structure including an outer annular wall, an inner annular wall, a radial annular wall and an inner annular baffle, defining a first cavity between the outer annular wall and the inner annular baffle, and a second cavity between the inner annular wall, the radial annular wall and the inner annular baffle.
Mold intended to be used for the manufacture of a gas turbine fan casing made of composite material including: a mandrel on which a fibrous preform of a fan casing is intended to be wound said preform comprising an annular wall the profile of the outer surface of which corresponds to that of the inner surface of the casing to be manufactured and two side flasks the profiles of which correspond to those of outer flanges of the casing to be manufactured; a plurality of angular counter-mold sectors assembled in a sealed manner on the mandrel and intended to close the mold and to compact a fibrous preform wound on the mandrel; and at least one compacting wedge, the wedge being placed between the fibrous preform wound on the mandrel and the associated angular counter-mold sector.
B29C 70/48 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.à d. d'objets distincts utilisant des moules opposables, p.ex. pour déformer des préimprégnés [SMC] ou des "prepregs" avec une imprégnation des renforcements dans le moule fermé, p.ex. moulage par transfert de résine [RTM]
The invention relates to a method (100) for producing a part made of composite material having an organic matrix, the method comprising the following steps: - arranging (110) a fibrous preform of the part to be produced in a mould comprising an impregnation chamber by resting a first face of the fibrous preform on a support surface of the impregnation chamber, the impregnation chamber being closed by a flexible membrane placed facing a second face of the fibrous preform, the flexible membrane separating the impregnation chamber from a compaction chamber; - injecting (120) a compaction fluid into the compaction chamber so as to apply a compaction pressure (Pcompaction) to the flexible membrane; - injecting (130) a resin into the impregnation chamber from a lateral face to the first face of the fibrous preform in a direction parallel to the surface of the membrane, so that the resin impregnates the fibrous preform and polymerises to form an organic matrix within the fibrous preform, the compaction pressure being maintained on the flexible membrane before and during the injection and polymerisation of the resin.
B29C 70/48 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.à d. d'objets distincts utilisant des moules opposables, p.ex. pour déformer des préimprégnés [SMC] ou des "prepregs" avec une imprégnation des renforcements dans le moule fermé, p.ex. moulage par transfert de résine [RTM]
B29C 70/44 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.à d. d'objets distincts utilisant une pression isostatique, p.ex. moulage par différence de pression, avec un sac à vide, dans un autoclave ou avec un caoutchouc expansible
B29C 33/40 - Matière plastique, p.ex. mousse ou caoutchouc
84.
POWER SHARING IN AN AIRCRAFT USING DISTRIBUTED CONTROL
The invention relates to a system (100) for exchanging power in an aircraft, the system comprising: - a voltage bus (160); - a low-pressure electromechanical converter (150 BP); - a high-pressure electromechanical converter (150 HP); and - for each electromechanical converter (150 BP, 150 HP), a voltage control module (125 BP, 125 HP). The system (100) further comprises: - a computer (106) designed to provide a sharing setpoint (S) between the exchanged powers (PHP, PBP), this sharing setpoint (S) varying over time; and - for at least one of the electromechanical converters (150 BP, 150 HP): a module (120 BP, 120 HP) for ensuring that the sharing setpoint (S) is observed, which module is designed to adjust the power (PHP, PBP) that is exchanged by the relevant electromechanical converter (150 BP, 150 HP) in order to observe the received sharing setpoint (S), taking into account a received evaluation (PBP°, PHP°) of the power (PBP, PHP) exchanged by the other of the electromechanical converters (150 BP, 150 HP).
B64D 31/00 - Commande des groupes moteurs; Leur disposition
B64D 41/00 - Installations génératrices de puissance pour servitudes auxiliaires
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 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
H02J 1/10 - Fonctionnement de sources à courant continu en parallèle
H02J 1/12 - Fonctionnement de générateurs à courant continu en parallèle avec des convertisseurs, p.ex. avec un redresseur à arc de mercure
H02J 3/46 - Dispositions pour l’alimentation en parallèle d’un seul réseau, par plusieurs générateurs, convertisseurs ou transformateurs contrôlant la répartition de puissance entre les générateurs, convertisseurs ou transformateurs
H02J 4/00 - Circuits pour réseaux principaux ou de distribution, la nature alternative ou continue du courant n'étant pas précisée
H02J 13/00 - Circuits pour pourvoir à l'indication à distance des conditions d'un réseau, p.ex. un enregistrement instantané des conditions d'ouverture ou de fermeture de chaque sectionneur du réseau; Circuits pour pourvoir à la commande à distance des moyens de commutation dans un réseau de distribution d'énergie, p.ex. mise en ou hors circuit de consommateurs de courant par l'utilisation de signaux d'impulsion codés transmis par le réseau
85.
POWER SHARING IN AN AIRCRAFT USING CENTRALISED CONTROL
The invention relates to a system (100) for exchanging power in an aircraft, the system comprising: - a voltage bus (160); - a low-pressure electromechanical converter (150BP); - a high-pressure electromechanical converter (150HP); and - for each electromechanical converter (150BP, 150HP): a voltage control module (130BP, 130HP) and a control module (140BP, 140HP) for ensuring that the exchange setpoint (GBP*, GHP*) is observed. The system (100) further comprises a power adjustment module (106) configured, for at least one of the electromechanical converters (150BP, 150HP): to determine a voltage correction (δVBP, δVHP); and to apply, in the voltage control module (130BP, 130HP) associated with the relevant electromechanical converter (150BP, 150HP), the voltage correction (δVBP, δVHP) to a voltage setpoint (VDC*) so that the voltage control module (130BP, 130HP) adjusts the bus voltage (VDC) to the corrected voltage setpoint (VDC*).
B64D 41/00 - Installations génératrices de puissance pour servitudes auxiliaires
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
F02C 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
H02J 1/10 - Fonctionnement de sources à courant continu en parallèle
H02J 1/12 - Fonctionnement de générateurs à courant continu en parallèle avec des convertisseurs, p.ex. avec un redresseur à arc de mercure
H02J 3/46 - Dispositions pour l’alimentation en parallèle d’un seul réseau, par plusieurs générateurs, convertisseurs ou transformateurs contrôlant la répartition de puissance entre les générateurs, convertisseurs ou transformateurs
H02J 4/00 - Circuits pour réseaux principaux ou de distribution, la nature alternative ou continue du courant n'étant pas précisée
H02J 13/00 - Circuits pour pourvoir à l'indication à distance des conditions d'un réseau, p.ex. un enregistrement instantané des conditions d'ouverture ou de fermeture de chaque sectionneur du réseau; Circuits pour pourvoir à la commande à distance des moyens de commutation dans un réseau de distribution d'énergie, p.ex. mise en ou hors circuit de consommateurs de courant par l'utilisation de signaux d'impulsion codés transmis par le réseau
B64D 31/00 - Commande des groupes moteurs; Leur disposition
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventeur(s)
Youssef, Toni
Escriva, Arnaud
Azzopardi, Stéphane Joseph
Khazaka, Rabih
Bley, Vincent
Valdez-Nava, Zarel
Le, Trong Trung
Diaham, Sombel
Abrégé
Another aspect of the invention relates to methods (200, 300) for manufacturing the power module (100) wherein the electronic component (110) is coated in a particle-loaded resin (140), an electric field being applied (220, 370) to the electronic component (110) in order to migrate the particles (142) from the loaded resin (140) and generate a particle concentration gradient around a predefined area (115) of the electronic component.
H05K 1/03 - Emploi de matériaux pour réaliser le substrat
B29C 70/62 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des matières de remplissage les matières de remplissage étant orientées pendant le moulage
H01L 23/29 - Capsulations, p.ex. couches de capsulation, revêtements caractérisées par le matériau
B29C 70/58 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts comprenant uniquement des matières de remplissage
B29C 70/68 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts en incorporant ou en surmoulant des parties préformées, p.ex. des inserts ou des couches
B29C 70/88 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts caractérisées principalement par des propriétés spécifiques, p.ex. électriquement conductrices ou renforcées localement
H01L 23/31 - Capsulations, p.ex. couches de capsulation, revêtements caractérisées par leur disposition
H01L 23/538 - Dispositions pour conduire le courant électrique à l'intérieur du dispositif pendant son fonctionnement, d'un composant à un autre la structure d'interconnexion entre une pluralité de puces semi-conductrices se trouvant au-dessus ou à l'intérieur de substrats isolants
H01L 25/07 - Ensembles consistant en une pluralité de dispositifs à semi-conducteurs ou d'autres dispositifs à l'état solide les dispositifs étant tous d'un type prévu dans le même sous-groupe des groupes , ou dans une seule sous-classe de , , p.ex. ensembles de diodes redresseuses les dispositifs n'ayant pas de conteneurs séparés les dispositifs étant d'un type prévu dans le groupe
H05K 1/18 - Circuits imprimés associés structurellement à des composants électriques non imprimés
The invention relates to a shim (2) for a moving blade (3), the moving blade (3) being intended to have a blade root (30) mounted in a slot (90) which opens onto the outer periphery of a turbomachine rotor disc (1) having a longitudinal axis (X), the shim (2) comprising a base (20) which is configured to be mounted in the slot (90) and which is configured to be mounted in contact with the blade root (30). According to the invention, the base (20) has at least one portion, referred to as an elastic portion, which is corrugated so as to apply a radial force towards the outside on the blade root (30) when the blade root (30) is resting on the base (20) mounted in the slot (90). Additionally, the base (20) has an upstream longitudinal end and a downstream longitudinal end configured to bear axially against one of the disc or the blade root (30).
The invention relates to a turbomachine rotor having an axis (X) comprising at least two circumferentially adjacent sets of blades (3). Each set of blades comprises: - a root (31) configured to be mounted in a groove which opens onto the outer periphery of a disc (2) of the turbomachine rotor; - a blade (33) extending the root in a radial direction (R) with respect to the axis and having an aerodynamic profile; - a radially inner platform (32) separating the blade from the root; - a shroud (34) which extends in the extension of the blade at a free radial end of the set of blades at the end of the blade that is radially opposite the root (31), the shroud (34) comprising a radially outer platform (40). The shroud comprises means for locking the radially outer platform with respect to another circumferentially adjacent radially outer platform. The locking means comprise two locking flanges (41a, 41b) of the radially outer platform which are configured each to receive a circumferential end of a locking plate (42). The rotor further comprises a locking plate, one of the two locking flanges of the first of the two sets of blades and one of the two locking flanges of the second of the two sets of blades each accepting a circumferential end of the locking plate. The locking flanges of the radially outer platform extend obliquely and are inclined in a direction opposite to the radial direction. The locking plate comprises bevelled circumferential ends which are configured to come into contact with the two locking flanges of the radially outer platform. The invention also relates to a turbomachine (1) for an aircraft comprising such a turbine. The invention also relates to a turbomachine turbine comprising at least one such rotor.
One aspect of the invention relates to a computer-implemented method (10) for discriminating icing conditions in an environment of an aircraft, the aircraft having an on-board ice accumulation sensor (20), the method (10) comprising, in each current measurement cycle of a plurality of measurement cycles: - acquiring (13) data from the ice accumulation sensor (20) for the current measurement cycle; - classifying (14) the data acquired for the current measurement cycle, by a machine learning algorithm, into one of a plurality of classes comprising at least: - a first class representative of a first icing condition in the environment of the aircraft and - a second class representative of a second icing condition in the environment of the aircraft.
This method (200) for producing a composite spacer ring for a turbine engine fan comprises preparing (210) an upstream skin preform, preparing (220) a downstream skin preform, preparing (230) an insert preform, inserting (251) the insert preform between the upstream skin preform and the downstream skin preform so as to form a rotating spacer ring preform, and firing (260) the rotating spacer ring preform. The preparation (210) of the upstream skin preform and/or the preparation (220) of the downstream skin preform comprises drape forming (212, 222) prepreg tows on a skin mold by means of automated fiber placement.
B29C 70/38 - Empilage automatisé, p.ex. utilisant des robots, par application de filaments selon des modèles prédéterminés
B29B 11/16 - Fabrication de préformes caractérisées par la structure ou la composition comprenant des charges ou des agents de renforcement
B32B 38/00 - Opérations auxiliaires liées aux procédés de stratification
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
F01D 5/02 - Organes de support des aubes, p.ex. rotors
F01D 5/06 - Rotors à plus d'un étage axial, p.ex. du type à tambour ou à disques multiples; Leurs parties constitutives, p.ex. arbres, connections des arbres
The invention relates to a mold (10) for drying a fibrous preform, comprising at least one part (11) having inner faces (13.1-13.5) defining a cavity (14) of the mold (10), against which the fibrous preform, wetted by a wetting liquid, is intended to be brought into contact, said part (11) having an openwork structure allowing the wetting liquid to pass through this part of the mold (10), the openwork part (11) consisting of a plurality of intersecting filaments (17) defining through-openings (18) allowing the passage of the wetting liquid.
B29C 33/38 - Moules ou noyaux; Leurs détails ou accessoires caractérisés par la matière ou le procédé de fabrication
B29C 70/00 - Façonnage de matières composites, c. à d. de matières plastiques comprenant des renforcements, des matières de remplissage ou des parties préformées, p.ex. des inserts
B29C 33/42 - Moules ou noyaux; Leurs détails ou accessoires caractérisés par la forme de la surface de moulage, p.ex. par des nervures ou des rainures
92.
INNER SHROUD SECTOR FOR AN AIRCRAFT TURBINE ENGINE
The invention relates to an inner shroud sector (1) configured to be mounted in an aircraft turbine engine, the inner shroud being a rotating part having a longitudinal axis (X) oriented from upstream to downstream, the inner shroud sector (1) comprising a main body (2) having an outer surface (Sext) configured to internally delimit a secondary flow path, the main body (2) comprising an upstream end (3) configured to be attached to a hub of an intermediate casing of the aircraft turbine engine, the main body (2) being made of a composite material comprising a plurality of reinforcing fibres in a matrix, the inner shroud sector (1) comprising at least one hollow column (13) mounted on the upstream end (3) of the main body (2), each hollow column (13) extending radially and being configured to accommodate an attachment element for attaching the inner shroud sector (1) to the hub of the intermediate casing.
The invention relates to an inner shroud sector (1) configured to internally delimit a secondary duct (41) of an aircraft turbomachine (50), the inner shroud sector (1) comprising a downstream end (2) which comprises an upstream longitudinal portion (3) and a downstream longitudinal portion (4) which are connected by a middle portion (5) and are configured to extend, respectively, in the upstream extension and internally to a nacelle inner casing (31), the inner shroud sector (1) comprising an inner layer and an outer layer of composite material, the middle portion (5) of the downstream end (2) comprising at least one volume of intumescent material sandwiched between the inner layer and the outer layer of the composite material so as to form a local thickened portion in the longitudinal direction.
F01D 25/26 - Carcasses d'enveloppe doubles; Mesures contre les tensions thermiques dans les carcasses d'enveloppe
F01D 5/28 - Emploi de matériaux spécifiés; Mesures contre l'érosion ou la corrosion
F02K 3/06 - Ensembles fonctionnels comportant une turbine à gaz entraînant un compresseur ou un ventilateur soufflant dans lesquels une partie du fluide énergétique passe en dehors de la turbine et de la chambre de combustion l'ensemble fonctionnel comprenant des soufflantes carénées, c. à d. des soufflantes à fort débit volumétrique sous basse pression pour augmenter la poussée, p.ex. du type à double flux comprenant une soufflante avant
F01D 25/24 - Carcasses d'enveloppe; Eléments de la carcasse, p.ex. diaphragmes, fixations
F01D 9/04 - Injecteurs; Logement des injecteurs; Aubes de stator; Tuyères de guidage formant une couronne ou un secteur
94.
POWER MODULE WITH ENCAPSULATED COMPONENT AND METHOD FOR MANUFACTURING THIS MODULE
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventeur(s)
Youssef, Toni
Escriva, Arnaud
Azzopardi, Stéphane Joseph
Khazaka, Rabih
Bley, Vincent
Valdez-Nava,, Zarel
Le, Trong Trung
Diaham, Sombel
Abrégé
One aspect of the invention relates to a power module (100, 100') comprising at least one power electronic component (110) buried in a printed circuit (120) comprising: first and second conductive substrates (121, 122), and an encapsulation resin (130) housed between the first and the second conductive substrates, the encapsulation resin (130) comprising at least two separate permittivities (ɛ1, ɛ2), the encapsulation resin extending away from the component comprising a first permittivity ɛ1 and the encapsulation resin extending near the electronic component (110) comprising a second permittivity ɛ2, greater than the first permittivity. Another aspect of the invention relates to a method (200) for manufacturing the power module (100, 100') wherein a first inner pre-impregnated sheet is cut to form cavities suitable for receiving, on one side, the electronic component, and on the other side, other pre-impregnated sheets having different permittivities.
The invention relates to a method for adjusting a turbine engine part to be placed in a gas stream of the turbine engine, the part having previously undergone a material addition operation in order to fill in a defect in at least one area of the surface thereof, resulting in an extra thickness in at least one area of the surface thereof, characterised in that the method comprises steps of: - ordering (E1) the formation of a three-dimensional image of the part; - identifying (E2) the at least one area of the surface of the part that includes the extra thickness, and identifying the extra thickness, in the three-dimensional image of the part; - determining (E3) a trajectory of an adjustment tool, as a function of the geometry of the extra thickness identified; and - ordering (E4) the adjustment of the part according to the trajectory determined.
G05B 19/4097 - Commande numérique (CN), c.à d. machines fonctionnant automatiquement, en particulier machines-outils, p.ex. dans un milieu de fabrication industriel, afin d'effectuer un positionnement, un mouvement ou des actions coordonnées au moyen de données d'u caractérisée par l'utilisation de données de conception pour commander des machines à commande numérique [CN], p.ex. conception et fabrication assistées par ordinateur CFAO
G05B 19/42 - Systèmes d'enregistrement et de reproduction, c. à d. dans lesquels le programme est enregistré à partir d'un cycle d'opérations, p.ex. le cycle d'opérations étant commandé à la main, après quoi cet enregistrement est reproduit sur la même machine
B23P 6/00 - Remise en état ou réparation des objets
F01D 5/00 - Aubes; Organes de support des aubes; Dispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
The invention relates to a fan section (2) of an aeronautical propulsion system comprising a fan rotor (9) provided with twenty-two blades (14) and having a solidity greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 1.0 and less than or equal to 1.3, where the solidity is equal to a ratio of a chord at the tip (21) of the blade (14) to an inter-blade pitch (14) at the tip of the blade, the fan section (2) having: - a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5; and - a tip speed at the tip (21) of the blade (14) greater than or equal to 260 m/s and 400 m/s; the pressure ratio and the tip speed being measured at cruising power.
A stator ring for an aircraft turbine engine includes an inner shroud coaxial with an outer shroud. The shrouds are connected to each other by vanes that are each fully integral with the shrouds. The outer shroud has an outer annular surface connected to at least one catch for attaching the stator ring. The inner shroud has an inner annular surface connected to a support member with an abradable coating. At least one of the inner and outer surfaces includes recesses that are situated in line with the vanes and are configured such that the vanes are connected to the corresponding shroud.
An air inlet of an aircraft propulsion unit comprising thrust reversal means configured to modify a bypass air flow circulating from upstream to downstream in a bypass duct into a reverse air flow circulating from downstream to upstream. The air inlet comprising an interior wall defining openings and, in each opening, a mobile member that comprises a covering wall and a deflecting wall and is mounted so as to pivot between a covering position, in which the covering wall obstructs the opening so as to guide the bypass air flow and promote a thrust phase, and a deflecting position, in which the deflecting wall obstructs the opening so as to separate the reverse air flow and promote a thrust reversal phase.
F02K 1/72 - Inversion du flux de la soufflante utilisant des volets inverseurs de poussée ou des portes montées sur le carter de la soufflante la partie arrière du carter de la soufflante étant mobile pour découvrir des ouvertures d'inversion de poussée dans le carter de la soufflante
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
99.
ELECTROHYDRAULIC PITCH SETTING WITH REVERSIBLE PUMP
The present invention relates to a mechanism for changing the pitch of a propeller of a turbine engine, the mechanism comprising an electrohydraulic actuator that comprises: —an electric machine comprising a rotatable actuator shaft; —a pumping assembly comprising two hydraulic axial piston pumps, each comprising: o a barrel housed in a cavity filled with hydraulic fluid; o a set of cylinders formed in the barrel, each cylinder housing a translatable piston and comprising an inlet port and a delivery port; o and an inclined plate, the plate of a first pump comprising a crescent-shaped inlet port for circulating the hydraulic fluid when the plate is rotated in a first direction, and the plate of a second pump comprising a crescent-shaped inlet port for circulating the hydraulic fluid when the plate is rotated in a second direction.
The present invention relates to a pitch-change mechanism for a turbine engine propeller, said turbine engine comprising a stator portion and a rotor portion, said pitch-change mechanism comprising an electrohydraulic actuator comprising: —a cylinder for actuating the propeller, comprising a first chamber and a second chamber; —a hydraulic pump suitable for selectively supplying the first chamber or the second chamber of the cylinder with pressurized hydraulic fluid; —a hydraulic fluid reservoir configured to store a pressurized hydraulic fluid; —a valve controlled between a first position in which the valve places the first chamber of the cylinder in fluid communication with the hydraulic pump, and a second position in which the valve places the first chamber of the cylinder in fluid communication with the hydraulic fluid reservoir.
