A method for producing a plurality of blades having a first height includes providing a mold having an internal cavity bounded by an internal side face of the mold which is axisymmetric about a longitudinal axis, the side face extending along the longitudinal axis to a second height which is greater than or equal to first height, the side face defining a plurality of sections perpendicular to the longitudinal axis, which have a radius for each section. The ratio of the height of the side face to the radius of the section is less than or equal to 2.0. The method also includes rotating the mold about the longitudinal axis at a rotation speed of between 200 rpm and 800 rpm, and pouring a material into the internal cavity in order to form, by centrifugation, a tubular part along the longitudinal axis.
B22D 13/02 - Coulée par centrifugationCoulée utilisant la force centrifuge de pièces longues, pleines ou creuses, p. ex. de tuyaux, coulées dans des moules tournant autour de leur axe longitudinal
An assembly for an aircraft turbine engine is disclosed. The aircraft turbine engine has a longitudinal axis and a distributor which has a ring of stator vanes including a foot at the radially inner end of the distributor, foot having an annular seal configured to engage in a non-contact sealing manner with a cylindrical shroud of a rotor of the turbine engine where the shroud is arranged radially under the distributor. The assembly includes the seal having a plurality of seal segments distributed circumferentially about the longitudinal axis, each seal segment including a radially outer annular wall segment and a radially inner annular wall segment connected to each other by an elastically deformable member, wherein each radially outer annular wall segment is attached to a seal support.
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
An assembly for an aircraft turbine engine having a longitudinal axis includes a distributor having a stator vane ring with a foot at the radially inner end of the distributor, the foot carrying an annular seal configured to cooperate in a non-contact sealing manner with a cylindrical rotor shroud of the turbine engine, the turbine engine being arranged radially under the distributor. The seal includes a plurality of seal segments distributed circumferentially about the longitudinal axis, each seal segment including a radially outer annular wall segment and a radially inner annular wall segment which are connected to each other by an elastically deformable member, wherein each radially outer annular wall segment is fixed to a seal support.
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
A seal for an aircraft turbine engine comprising a plurality of seal segments distributed circumferentially about a longitudinal axis, each seal segment having a radially outer annular wall segment and a radially inner annular wall segment which are connected to each other by an elastically deformable member. The elastically deformable member includes an inner leg having a radially inner end fixed to the radially inner annular wall segment and an outer leg having a radially outer end fixed to the radially outer annular wall segment, the radially outer end of the inner leg and the radially inner end of the outer leg being connected to each other by a connecting wall, the inner leg and the outer leg extending radially such that the radially inner end of the outer leg is arranged radially inside the radially outer end of the inner leg.
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
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
5.
AC POWER GENERATION COMPRISING A NON-INDUCTIVE OVERVOLTAGE PROTECTION DEVICE
A device for protecting against fast overvoltages and slow overvoltages placed in parallel with an alternating-current generation line includes: a first current source including an input and an output, the input of the first current source being connected to a first terminal of the generation line, a second current source including an input and an output, the input of the second current source being connected to a second terminal of the generation line, the output of the second current source being connected to the output of the first current source.
The invention relates to a method for assisting in the diagnosis of an aeronautical system (20), the aeronautical system (20) comprising a set of equipment, the method comprising automatically processing a document (x) describing a failure of the aeronautical system, the method further comprising providing the document to a predictive model (11) suitable for providing a classification vector, each component of which associates a score with a class, and providing said vector to a conformal prediction module (12) suitable for determining a set of combinations of classes characterising the failure and for implementing steps of generating candidate combinations and selecting the combinations satisfying a comparison criterion, wherein the step of generating combinations comprises constructing a tree whose nodes correspond to the candidate combinations, the construction comprising comparing a lower bound of the non-conformity score for a node with the quantile, and generating the descendants of the node only if said lower bound is less than or equal to the quantile.
G06Q 10/04 - Prévision ou optimisation spécialement adaptées à des fins administratives ou de gestion, p. ex. programmation linéaire ou "problème d’optimisation des stocks"
G06Q 10/20 - Administration de la réparation ou de la maintenance des produits
G06Q 50/40 - Procédés d’affaires s’appliquant à l’industrie du transport
7.
METHOD FOR PRODUCING AN ANTI-WEAR COATING FOR A PART MADE OF ORGANIC MATRIX COMPOSITE MATERIAL
The invention relates to a method (100) for producing an anti-wear coating for a part made of organic matrix composite materials. The method (100) comprises: - carrying out physical vapour deposition (101) of an anti-wear functional layer such as titanium nitride or tungsten-doped hydrogenated carbon on a first face of a metal substrate; - preparing (103) the surface of a second face, opposite the first face, of the metal substrate; - preparing (109) the surface of a contact area of an outer surface of the part; - draping (111) a film adhesive layer over the contact area; - positioning (113), on the film adhesive layer, the metal substrate covered with the thin layer; and, - polymerising (115) the film adhesive layer.
C23C 14/06 - Revêtement par évaporation sous vide, pulvérisation cathodique ou implantation d'ions du matériau composant le revêtement caractérisé par le matériau de revêtement
C09J 5/02 - Procédés de collage en généralProcédés de collage non prévus ailleurs, p. ex. relatifs aux amorces comprenant un traitement préalable des surfaces à joindre
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
8.
AUTOMATED WEAVING METHOD WITH OPTIMIZED SHED OPENING
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 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 3/22 - Fabrication de pièces ou d'objets à partir de poudres métalliques, caractérisée par le mode de compactage ou de frittageAppareils 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 3/24 - Traitement ultérieur des pièces ou objets
B22F 9/08 - Fabrication des poudres métalliques ou de leurs suspensionsAppareils 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
C22C 19/07 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de cobalt
10.
MANAGEMENT OF THE DRAWING OF MECHANICAL POWER FROM A TWO-SPOOL OR THREE-SPOOL TURBINE ENGINE
The present invention relates to a method and system for managing the drawing of mechanical power from a two-spool or three-spool turbine engine (1) for an aircraft, in which at least two electric machines (3, 4) are suitable for recovering mechanical energy, one from a shaft driven by one of the turbines of the turbine engine (1), the other from a shaft driven by another turbine, in which the distribution of the drawing between one and the other of the two electric machines (3, 4) is controlled dynamically according to the phases of flight.
A seal for an aircraft turbine engine includes a plurality of seal segments that are circumferentially distributed about a longitudinal axis. Each seal segment has a radially outer annular wall segment and a radially inner annular wall segment which are connected to each other by an elastically deformable member. The radially outer annular wall segments form a monolithic outer shroud, and the radially inner annular wall segments are arranged circumferentially end-to-end. Each radially inner annular wall segment has a first circumferential edge and an opposite second circumferential edge which each has a slit. A tongue is mounted in part at the junction of two facing circumferential edges.
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
12.
TOOL FOR BONDING AN ANTI-WEAR STRIP TO A TURBOMACHINE VANE
A tool for bonding at least one anti-wear strip to a turbomachine blade comprising: —a support, —at least one first fastening member which is fastened to the support and is able to cooperate with the root of the blade in order to immobilize it with respect to the support and —at least one second member for holding said at least one anti-wear strip on the blade and for heating this anti-wear strip for the bonding thereof, this second member being fastened to the support and comprising, on the one hand, at least one expandable bag, and, on the other hand, at least one heating element able to heat said at least one anti-wear strip for the bonding thereof.
An oil tank for an aircraft turbomachine, including: an enclosure capable of containing the oil and having a main axis intended to be oriented vertically in the mounting position; an oil inlet at an upper part of the enclosure; an oil outlet at a lower part of the enclosure. The oil tank further includes: a wheel arranged in the lower part of the enclosure and capable of being driven in rotation by drive means so as to centrifuge the oil against a retaining wall of the enclosure and against the oil outlet, and thus maintain an oil supply to said oil outlet during flight phases of the aircraft in zero or negative gravity.
The invention relates to a tool (1) for bonding a thermoplastic polymer film (3) to a vane (2) of an aircraft turbine engine, the tool comprising a mould (6) that includes a cavity (8) configured to receive the vane (2) or blade at least partially covered with the thermoplastic polymer film (3), a draining fabric layer (11) intended to cover the thermoplastic polymer film (3), a vacuum bag (12) intended to cover at least the draining fabric layer (11) and configured to apply a vacuum within the cavity (8), a pressurising device (13') intended to apply a pressure to the vacuum bag (12), and a heating device configured to receive the mould (6).
The invention relates to a tool (1) for bonding a thermoplastic polymer film (3) to a vane (2) of an aircraft turbine engine, the tool comprising a mould (6) that includes a cavity (8) configured to receive the vane (2) or the blade at least partially covered with the thermoplastic polymer film (3), a heating device (9, 10) intended to cover at least part of a surface (4) of the vane (2) or of the blade opposite the bottom of the cavity (8), a draining fabric layer (11) intended to cover the heating device (9, 10), a vacuum bag (12) intended to cover at least the draining fabric layer (11) and configured to apply a vacuum within the cavity (8), and a pressurising device (13') intended to apply a pressure to the vacuum bag (12).
A turbomachine stator extending about a central axis and including an inner wall and an outer wall delimiting an annular flow passage, a plurality of variable-pitch vanes, each vane extending radially in the annular flow passage between a base movably fixed to the outer wall, and a free end, the inner wall including a main profile and at least one circular platform arranged opposite the free end of one of the vanes, the platform including a main surface and at least one longitudinal groove, the platform being movable in rotation relative to the main profile between a first position in which a main axis of the groove is substantially parallel to the central axis, and at least a second position in which the main axis is transverse to the central axis.
F01D 9/04 - InjecteursLogement des injecteursAubes de statorTuyères de guidage formant une couronne ou un secteur
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
An aircraft part includes a composite material structure including a fibrous reinforcement and matrix, the fibrous reinforcement being embedded into the matrix, and heating elements that heat the aircraft part, each of the heating elements being embedded into the matrix of the composite material structure and being distributed across the aircraft part with a density depending on a position in the aircraft part.
A method for achieving inertial localization of a carrier using a stochastic filter of recursive-Bayesian type implemented by an inertial localization device, the method including a plurality of cycles, each cycle including a propagation procedure and, when a measurement is available, an updating procedure, in which method, in the propagation procedure, the propagation is performed using a propagation model taking into account at least one component of a matrix of the spatial gradients of the gravity vector, which matrix is delivering using an enriched gravimetric model.
G01C 21/16 - NavigationInstruments de navigation non prévus dans les groupes en utilisant des mesures de la vitesse ou de l'accélération exécutées à bord de l'objet navigantNavigation à l'estime en intégrant l'accélération ou la vitesse, c.-à-d. navigation par inertie
The invention relates to a shell (1) for fabricating at least one single-crystal component by molding using a removable material, comprising a central casting pit (11) configured for introduction of molten metal therein, the shell (1) comprising at least one mold (12) arranged around the central pit (11), each mold (12) comprising an inner cavity (121) in fluidic communication with the central pit (11), the cavity (121) being configured so that the molten metal fills the cavity (121) and forms a component, the shell (1) comprising at least one generally annular heat shield (13) arranged around the central pit (11), the at least one heat shield (13) further comprising an outer wall (131) at least partially covered with an outer coating (135) having a surface roughness Ra less than 5 µm and an emissivity E greater than 0.6 when measured at a temperature between 1400 and 1600°C.
A serial data bus asynchronous communication system (100) includes a bus controller, BC, module (102) in signal communication with a bus network (101,103,105) and a remote terminal, RT, module (150) in signal communication with the bus network (101, 103, 105) to exchange data with the BC module (102). The RT module (150) includes a RT controller (152) implemented with a read/write serial peripheral interface, SPI, (160) and a timing controller (162) in signal communication (168) with the read/write SPI (160). The read/write serial peripheral SPI (160) manages transfer of the data. The timing controller (162) determines an active data access of the RT module (150) by the BC module (102) and outputs a trigger signal (166) indicating the active data access, and the read/write SPI (160) manages the transfer of the data based on the trigger signal (166).
G06F 13/12 - Commande par programme pour dispositifs périphériques utilisant des matériels indépendants du processeur central, p. ex. canal ou processeur périphérique
G06F 13/362 - Gestion de demandes d'interconnexion ou de transfert pour l'accès au bus ou au système à bus communs avec commande d'accès centralisée
H04L 12/417 - Réseaux à ligne bus avec commande décentralisée avec accès déterminé, p. ex. passage de jeton
H04Q 9/04 - Dispositions à fonctionnement synchrone
21.
AIRCRAFT PROPULSION SYSTEM WITH TURBINE ENGINE COUPLED TO AN ELECTRIC GENERATOR THE POWER OF WHICH IS LIMITED AS A FUNCTION OF ENGINE SPEED
An aircraft propulsion system including a gas turbine engine with a combustion chamber and an engine spool, the engine spool including a turbine downstream of the combustion chamber, a compressor upstream of the combustion chamber and a transmission shaft; a control system; a member for determining a current engine speed of the engine spool; and a rotary machine for generating electrical power removing mechanical power from the transmission shaft. The control system limits a current electrical power generated by the rotary machine, at least when the engine spool is in the acceleration phase, to a maximum authorized power, the maximum authorized power being a function of the current engine speed before acceleration, and the propulsion system includes a device for determining the current electrical power and the control system establishes the engine speed above an idle speed which is a function of the current electrical power.
A module for a turbine engine, in particular for an aircraft turbine engine, includes a shaft extending in an axial direction and an electric machine including a rotor rotatably coupled to the shaft of the module, the rotor comprising a disc and magnetic elements arranged regularly around the periphery of the disc; and a stator rigidly attached to a casing of the module including a ring and coils distributed annularly inside the ring of the stator, the coils being arranged outside the magnetic elements of the rotor in a radial direction, wherein 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.
H02K 1/22 - Parties tournantes du circuit magnétique
H02K 7/00 - Dispositions pour la mise en œuvre d'énergie mécanique associées structurellement aux machines dynamo-électriques, p. ex. association structurelle avec des moteurs mécaniques d'entraînement ou des machines dynamo-électriques auxiliaires
A seal for an aircraft turbine engine includes a plurality of seal segments distributed circumferentially about a longitudinal axis, each seal segment having a radially outer annular wall segment and a radially inner annular wall segment which are connected to each other by an elastically deformable member. Pairs of circumferentially adjacent seal segments have their respective elastically deformable member implemented monolithically as a common deformable elastic member, the common deformable elastic member connecting together two radially inner annular wall segments of circumferentially-adjacent seal segments such that the radially outer annular wall segments of the seal segments form a monolithic outer shroud.
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 a shell (1) for fabricating at least one single-crystal component by molding using a removable material, comprising a central casting pit (11) configured for introduction of molten metal therein, the shell (1) comprising at least one mold (12) arranged around the central pit (11), each mold (12) being configured such that the molten metal fills a cavity (121) in a main filling direction and forms a component, the shell (1) comprising at least one heat shield (13) arranged in the vicinity of the at least one mold (12) and extending substantially in the filling direction along the at least one mold (12), the at least one heat shield (13) further comprising an outer wall (131) having a surface roughness Ra less than 5 µm and an emissivity E greater than 0.6 when measured at a temperature between 1400 and 1600°C.
G05D 1/46 - Commande de la position ou du cap dans les trois dimensions
G05D 1/646 - Suivi d’une trajectoire prédéfinie, p. ex. d’une ligne marquée sur le sol ou d’une trajectoire de vol
G05D 1/695 - Commande coordonnée de la position ou du cap de plusieurs véhicules pour maintenir la position des véhicules les uns par rapport aux autres, p. ex. pour les voyages en convoi ou les vols en formation
The invention relates to a composition of a nickel superalloy consisting of: - an aluminum content by mass of between 0% and 0.5%, - a cobalt content by mass of between 0% and 12%, - a chromium content by mass of between 14% and 20%, - an iron content by mass of between 0% and 10%, - a molybdenum content by mass of between 0% and 2%, - a niobium content by mass of between 0% and 1%, - a tantalum content by mass of between 13% and 17%, - a titanium content by mass of between 0% and 1%, - a tungsten content by mass of between 0% and 3%, - a carbon content by mass of between 0% and 0.05%, - a manganese content by mass of between 0% and 0.5%, - a silicon content by mass of between 0% and 0.5%, as well as unavoidable impurities and the balance being nickel.
C22C 19/05 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de nickel avec du chrome
C22C 1/04 - Fabrication des alliages non ferreux par métallurgie des poudres
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
The invention relates to a turbomachine compressor casing comprising an inner annular wall (101) and an outer annular wall (102) delimiting between them an annular cavity (150) extending in length in an axial direction (DA) between a front bottom (151) and a rear bottom (152) and in height in a radial direction (DR) between an inner face (1021) of the outer annular wall and an outer face (1011) of the inner annular wall, the inner annular wall of the casing comprising a plurality of slots (110) which open into the annular cavity, are arranged one next to the other in a circumferential direction (DC) and extend in length in the axial direction between the front bottom and the rear bottom and in height in the radial direction, characterized in that, in the axial direction, the annular cavity comprises a first portion (155) situated above, in the radial direction, an upstream part of the slots and a second portion (156) situated downstream of the first portion, the height of the annular cavity in the first portion (H155) being greater than that of the second portion (H156).
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
F04D 29/52 - Carters d'enveloppeTubulures pour le fluide énergétique pour pompes axiales
F04D 29/68 - Lutte contre la cavitation, les tourbillons, le bruit, les vibrations ou phénomènes analoguesÉquilibrage en agissant sur les couches limites
A method for detecting an anomaly in a plurality of predefined zones of a single aeronautical component includes supplying an image to be inspected to a detection system that includes a backbone neural network configured to supply an intermediate representation of the image to be inspected, and localisation neural networks that are respectively associated with the zones and each designed, on the basis of the intermediate representation, to search for one or more anomalies and to locate each anomaly found in the image to be inspected. The detection system has been previously trained, for each zone, by a step of supplying, to the backbone network, training images, and, on the basis of the training images of the relevant zone, an adaptive adjustment step during which parameters of the backbone network and of the localisation network that is associated with the relevant zone are adjusted.
G06V 10/25 - Détermination d’une région d’intérêt [ROI] ou d’un volume d’intérêt [VOI]
G06V 10/762 - Dispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant le regroupement, p. ex. de visages similaires sur les réseaux sociaux
G06V 10/764 - Dispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant la classification, p. ex. des objets vidéo
G06V 10/774 - Génération d'ensembles de motifs de formationTraitement des caractéristiques d’images ou de vidéos dans les espaces de caractéristiquesDispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant l’intégration et la réduction de données, p. ex. analyse en composantes principales [PCA] ou analyse en composantes indépendantes [ ICA] ou cartes auto-organisatrices [SOM]Séparation aveugle de source méthodes de Bootstrap, p. ex. "bagging” ou “boosting”
G06V 10/82 - Dispositions pour la reconnaissance ou la compréhension d’images ou de vidéos utilisant la reconnaissance de formes ou l’apprentissage automatique utilisant les réseaux neuronaux
29.
TURBOMACHINE ASSEMBLY COMPRISING PLATFORMS HAVING EDGES PROVIDED WITH COMPLEMENTARY PROTRUSIONS AND NOTCHES AND ASSOCIATED TURBOMACHINE
A bladed turbomachine assembly extends about an axis (AX) and includes disc carrying blades (21a-21d), wherein each blade (21a-21d) extends radially from a root by which it is rigidly connected to the disc to a heel including a platform (26a-26d). The assembly includes an alternation of blades (21a, 21c) prestressed towards upstream and blades (21b, 21d) prestressed towards downstream, each blade (21a, 21c) prestressed towards upstream having its heel (24c) tending to shift towards upstream, each blade (21b, 21d) prestressed towards downstream having its heel (24b) tending to shift towards downstream.
A bladed turbomachine assembly extends about an axis (AX) and includes disc carrying blades (21a-21d), wherein each blade (21a-21d) extends radially from a root by which it is rigidly connected to the disc to a heel including a platform (26a-26d). The assembly includes an alternation of blades (21a, 21c) prestressed towards upstream and blades (21b, 21d) prestressed towards downstream, each blade (21a, 21c) prestressed towards upstream having its heel (24c) tending to shift towards upstream, each blade (21b, 21d) prestressed towards downstream having its heel (24b) tending to shift towards downstream.
The pairs of circumferentially adjacent blades have platforms (26a-26d), contiguous circumferential ends of which (27b-27d, 28a-28c) include protrusions (29a-29c) and notches (31b-31d) fitting together.
A bladed turbomachine assembly extends about an axis (AX) and includes disc carrying blades (21a-21d), wherein each blade (21a-21d) extends radially from a root by which it is rigidly connected to the disc to a heel including a platform (26a-26d). The assembly includes an alternation of blades (21a, 21c) prestressed towards upstream and blades (21b, 21d) prestressed towards downstream, each blade (21a, 21c) prestressed towards upstream having its heel (24c) tending to shift towards upstream, each blade (21b, 21d) prestressed towards downstream having its heel (24b) tending to shift towards downstream.
The pairs of circumferentially adjacent blades have platforms (26a-26d), contiguous circumferential ends of which (27b-27d, 28a-28c) include protrusions (29a-29c) and notches (31b-31d) fitting together.
Each protrusion (29a-29c) is pressed into a notch (31b-31d) receiving it according to the axial direction (AX) by the prestresses of the blades that carry them.
The invention relates to a variable-pitch blade (7) for a turbine engine, the blade comprising: - a tubular metal sleeve (13) comprising an opening (17) extending through the sleeve (13) along a pitch axis (Y), the opening (17) being delimited by an internal wall of the sleeve (13); - a fibrous structure (20) comprising a blade root (22), the blade root (22) comprising a tubular attachment portion (23) inserted axially into the opening (17) of the sleeve (13), this attachment portion (23) comprising a housing (20a) centered on the pitch axis (Y); - a polymerized resin for solidifying the fibrous structure (20) and securing the fibrous structure (20) to the internal wall (18) of the sleeve (13), and - a permanent insert (224) centered on the pitch axis (Y) and inserted axially into the housing (20a) of the attachment portion (23), the permanent insert (224) being separated from the attachment portion (23) by a layer of adhesive (226).
The invention relates to a rotor blade (20) for an aircraft turbine engine, this blade comprising an airfoil (22) and a root connected to the airfoil by a platform, the airfoil extending from the platform to a free end (22a) comprising a recessed bathtub (36), the rotor blade further comprising an internal circuit (38) for ventilation air to pass therethrough, the internal circuit (38) extending from the root to the free end and opening out through at least one air outlet opening (42) formed in the bottom (44) of the bathtub, the at least one air outlet opening extending along an axis (X) that is not parallel to a longitudinal axis (A) of the airfoil, the at least one opening being oriented toward a wall (33) of a pressure-side face (32) of the airfoil and toward the trailing edge (30) of the airfoil when the opening opens out at the bottom of the bathtub.
The invention relates to a turbine engine (1) of an aircraft (2), comprising: - a gas generator (3) of longitudinal axis (X1) which comprises a power turbine (4); - an unducted fan (5) which is movable about a longitudinal axis (X2) which is parallel to the axis (X1) and offset radially with respect to the axis (X1), the fan (5) being driven by the power turbine (4) via a speed reducer (6) which is arranged radially above an air inlet (7) of the gas generator (3); - an unshrouded stator (8) which is rotationally fixed and arranged axially downstream of the unducted fan (5); characterized in that the stator (8) comprises a ring (9) which supports vanes (10, 11) and which jointly surrounds a casing (12) of the speed reducer (6) and the air inlet (7), the ring (9) being fastened to the casing (12) of the speed reducer (6).
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F02C 7/36 - Transmission de puissance entre les différents arbres de l'ensemble fonctionnel de turbine à gaz, ou entre ce dernier et l'utilisateur de puissance
33.
METHOD FOR MANUFACTURING A FIBROUS PREFORM FOR A COMPOSITE PART INTENDED TO BE ARTICULATED WITH OTHER PARTS
The present invention relates to a method for manufacturing a fibrous preform of a core for a composite part intended to be articulated with other parts, comprising forming a one-piece strip by three-dimensional weaving, shaping said strip to give it a generally H-shaped cross-section, and positioning a reinforcing counter-form with rounded corners in order to reduce local stress concentrations when the part is under load.
The invention relates to a fibrous web (10) formed by three-dimensional weaving of first and second yarns, the fibrous web (10) being elongate along a longitudinal axis and comprising a debonding zone (20) present between a first region (11) and a second region (12), the first region comprising woven first lateral portions (11a) connected by a first intermediate portion, the second region comprising second lateral portions (12a) connected by a second intermediate portion (12b) and woven with the first lateral portions, the first yarns extending transversely to the longitudinal axis and comprising a first set (412) extending in the second region and a second set (413) extending in the first region and overlying the debonding zone, the second set (413) crossing the first set (412) in a crossing zone (30) adjacent to the edges (20a) of the debonding zone, so as to extend into the second lateral portions.
An oil tank for a turbomachine, including an enclosure for the oil, having a main axis, and with walls delimiting a first compartment and a second compartment separated from the first compartment by a separation wall; a channel extending into the first compartment, from the separation wall; an oil inlet mixed with air, arranged on the second compartment; an oil outlet including a conduit extending and opening into the first compartment. The channel extends inside the conduit, so as to ensure oil filling of the conduit and thus maintain an oil supply to the conduit during flight phases in zero or negative gravity.
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
F01M 11/00 - Parties constitutives, détails ou accessoires non couverts par les groupes ou présentant un intérêt autre que celui visé par ces groupes
36.
PROPULSION ASSEMBLY COMPRISING AN EJECTION DUCT FOR A DISCHARGE VALVE
The present disclosure relates to a propulsion assembly (1) comprising: - a main channel (9) which extends between an air inlet (91) and a splitter (92) that divides the main channel (9) into a primary flow path (10) and a cooling flow path (20), - an ejection duct (30) which extends between an ejection inlet (31) that opens into the primary flow path (10) and an ejection outlet (32, 32', 32''), and - a discharge valve (40) which is configured to discharge an air flow into the ejection duct (30), the ejection duct (30) comprising a portion which passes through the cooling flow path (20) such that the ejection outlet (32, 32', 32'') opens out radially outside an external casing of the propulsion assembly (1), the ejection outlet (32, 32', 32'') being located upstream of the outlet of the cooling flow path (20).
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
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 3/075 - 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
37.
ASSEMBLY FOR AN AIRCRAFT TURBINE ENGINE, COMPRISING A SERVO-CONTROLLED AXIAL SEAL SYSTEM
The main subject matter of the invention is an assembly (100) for an aircraft turbomachine, comprising a first stator part (2), a second rotor part (3), and a servo-controlled axial seal system (1) comprising a body (7) sliding freely on the first part (2), the assembly defining a main leakage flow path (14) extending from a first enclosure (6) to a second enclosure (5), and the servo-controlled axial seal system comprising a main pressure drop zone (20), as well as a pressure drop initiation zone (30) comprising at least two radial fins (109) spaced axially apart from one another.
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
A turbine rotor assembly of a turbine engine includes a rotor disc centred on an X-axis and having recesses, movable blades each including 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.
A turbine ring assembly provided with a CMC-sectored turbine ring and a ring support structure, each sector of the ring containing a base from which an upstream hooking lug and a downstream hooking lug axially spaced from each other extend radially outwards, the support structure including an upstream radial clamp and a downstream radial clamp between which the hooking lugs of each sector are held, and the assembly including, for each sector, at least one pin and at least one passage passing axially through a hooking lug over an axial length of the passage. At least one pin containing a planar bearing surface facing an inner surface of the passage through which the pin passes, the bearing surface facing the inner surface extending axially over a length smaller than the axial length of the passage.
Methods, apparatus, systems and articles of manufacture are disclosed. An apparatus for mounting a gas turbine engine to a pylon, the gas turbine including an upstream section and a downstream section, the gas turbine defining a roll axis, a yaw axis, and a pitch axis, the apparatus including: a first mount to couple the upstream section of the gas turbine engine to the pylon; a second mount to couple the upstream section of the gas turbine engine to the pylon, the second mount downstream of the first mount; a thrust linkage to couple the upstream section to the pylon, wherein the downstream section is decouplable from the upstream section without decoupling the first mount, the second mount, and the thrust linkage.
The present invention relates to a method for manufacturing an annular casing (3) for a turbomachine, this casing comprising: - an envelope (5) extending around an axis (X) and made of a composite material; - a layer (4) made of abradable material extending at the inside of the envelope; and - support panels (6) extending around the axis (X) and interposed between the envelope and the abradable layer, the method comprising the following steps: (a) fastening the support panels to the annular envelope; (b) preparing a paste by mixing a polymer resin and a fibrous reinforcement; (c) depositing the paste on an inner annular surface, for example of the support panels; and (d) curing the paste to form the casing, the method being characterized in that, in step (b), the paste further comprises a curing accelerator, for example based on an amine resin.
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 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
42.
DEVICE FOR COOLING A TURBINE CASING FOR A TURBOMACHINE
One aspect of the invention relates to a cooling device (100) extending around an axis (X) and intended to cool a coaxial turbine casing for a turbomachine, the cooling device (100) comprising at least one cooling tube (120) extending circumferentially around the axis (X), the at least one tube (120) comprising a plurality of primary circulation orifices intended to cool the casing, the cooling device (100) being characterized in that the at least one tube (120) comprises a discharge orifice intended to discharge particles present in the at least one tube (120), the discharge orifice being located on a closure face closing the distal end (122) of the tube (120).
The invention relates to a method for impregnating a fibrous preform (110) comprising a three-dimensional weave, the method being characterised in that it comprises the following steps: - arranging a fibrous preform comprising a three-dimensional weave in the moulding cavity of a mould (100); - arranging a powder (121, 122) made of thermoplastic material having a particle size of between 1.0 µm and 2000 µm in the moulding cavity of the mould; - a step of compressing the elements present in the moulding cavity of the mould, in order to allow the powder made of thermoplastic material to impregnate the fibrous preform; and - a step of heating to a temperature greater than or equal to the glass transition temperature of the thermoplastic material.
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
B29C 70/46 - 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"
B29L 31/08 - Pales pour rotors, stators, ventilateurs, turbines ou dispositifs analogues, p. ex. hélices
44.
ASSEMBLY COMPRISING AN AIRCRAFT TURBINE ENGINE AND ITS MOUNTING PYLON
The invention relates to an assembly (201) comprising a turbine engine (203) of an aircraft (200) and a mounting pylon (205) for the turbine engine. The assembly (201) comprises a first suspension member (217) connected to an upstream structural casing (209), a second suspension member (219) connected to an intermediate structural casing (211), and a third suspension member (221) connected to a downstream structural casing (215). The third suspension member (221) comprises a cavity (223), which is secured to the pylon (205) on its internal part (205a) and comprises an opening (223a) in an internal surface (223b), a block of material (225), which is accommodated in the cavity (223), and a coupling device (227) which is inserted into the cavity (223) at the opening (223a) and passes through the block of material (225). The coupling device (227) has a plate (227a) bearing on an external surface (225a) of the block of material (225) such that, when the block of material (225) expands, the block of material (225) exerts a pressure force on the plate (227a), thereby causing a radial movement of the coupling device (227).
The invention relates to an assembly (201) comprising a turbine engine (203) of an aircraft (200) and a mounting pylon (205) for the turbine engine (203). The assembly (201) comprises a first suspension member (217) connected to an upstream structural casing (209), a second suspension member (219) connected to an intermediate structural casing (211), and a third suspension member (221) connected to a downstream structural casing (215). The third suspension member (221) comprises a coupling device (223) that extends radially from the downstream structural casing (215) and is connected to a pivot (225) of a sliding pivot connection (227), the sliding pivot connection (227) also comprising a connection pin (A) that is fastened to the pylon (205) and forms an acute angle (θ) with the axis (X) such that a modification of the axial position of the coupling device (223) brings about a modification of the radial position of the coupling device (223), as a result of the pivot (225) sliding along the connection pin (A).
The present disclosure relates to a sealing assembly for a turbine engine, the assembly comprising a stator part (1), a rotor part (2), an axial seal (3) that provides sealing between an outer cavity (4) and an inner cavity (5) and comprises an inner platform (32) suitable for sliding on a downstream axial end (11) of the stator part (1) as well as a radial face (31) mounted axially opposite an upstream face (21) of the rotor part (2), the sealing assembly further comprising at least two fins (7) situated between the radial face (31) and the upstream face (21), an additional fin (8), an outer platform (33) designed to extend opposite the additional fin (8), and an injection duct (35) passing through the radial face (31) of the axial seal (3) and opening radially between the two fins (7).
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 present disclosure relates to a propulsion assembly (1) comprising: - a main channel (9) which extends between an air inlet (91) and a splitter (92) that divides the main channel (9) into a primary flow path (10) and a cooling flow path (20), - an ejection duct (30) which extends between an ejection inlet (31) that opens into the primary flow path (10) and an ejection outlet (32), and - a discharge valve (40) which is configured to discharge an air flow into the ejection duct (30), the ejection duct (30) being arranged between the cooling flow path (20) and the primary flow path (10) such that the ejection outlet (32) opens out radially outside the external casing of the propulsion assembly (1), the ejection outlet (32) being located downstream of the cooling outlet (22).
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 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
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
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
An aeronautical propulsion system includes a ducted fan section with a fan shaft extending along a longitudinal axis, supported by a first guiding system. A drive shaft to be driven by a turbine and that is supported by a second guiding system includes thrust bearings. A reduction structure, including an epicyclic gear train with a ring gear, couples the drive shaft to the fan shaft such that the drive shaft rotates at a higher speed than the fan shaft. A support attaches the reduction structure to a stator. A coupling device connects the drive shaft to the reduction structure and compensates for misalignment therebetween. The coupling device includes shaft sections and radial flexibility elements connecting the shaft sections. Specific dimensional ratios between axial distances of the guiding systems, reduction structure, support, and coupling device elements relative to a reference plane are defined to optimize system performance.
The invention relates to a device for removably attaching a piece of equipment (E), comprising a base (200) and a frame (100) to which the piece of equipment is rigidly connected; the frame (100) having a first side (101) and a second side (102), opposite each other with respect to a first axis (A1) of the frame (100), and a base plate (103) extending between the sides (101, 102) in parallel with the first axis (A1) in order to rest on an upper surface (203) of the base (200). The device comprises at least a first abutment (104) arranged on the frame (100) in order to be engaged under a hook (204) rigidly connected to the base (200), and a second abutment (106) arranged on the frame (100) in order to engage with a tie rod (206) connected to the base (200).
The invention relates to an assembly for a turbine engine having a longitudinal axis X, the assembly being configured to be arranged between a lubrication chamber of a speed reducer and a lubricant-tight zone, the assembly comprising: a casing arm configured to be provided in an air circulation duct intended to at least partially supply a gas generator of the turbine engine, the casing arm being formed by a wall delimiting a hollow interior space configured to be passed through by a cable, the hollow interior space being closed radially towards the interior by a lubricant-tight bottom partition, the bottom partition having a cable passage opening; a ring dimensioned so as to sealingly surround a portion of the cable, the ring being housed in the cable passage opening and being fastened to the bottom partition, the ring comprising a main body having a shape substantially complementary to the opening; sealing means provided at a junction interface between the ring and the opening.
The invention relates to a clip (30) for retaining a harness (24) for an aircraft engine testing machine (10), the clip (30) being formed as a single piece with two lateral legs (32, 34) connected by a bridge (36) extending between the legs (32, 34), a first of the lateral legs (32) comprising two branches (38, 40) which extend along an axis (A) perpendicular to the bridge (36) and which define between them a space (42) through which a harness (24) can pass, the branches (38, 40) having first axial ends (38a, 40a) connected together and to the bridge (36) and second axial ends (38b, 40b) which are opposite the first axial ends (38a, 40a) and which define a narrowed section (44) of said space (42).
F16B 2/22 - Clips, c.-à-d. dispositifs de fixation à serrage effectué uniquement par la résistance à la déformation inhérente au matériau dont est fait le dispositif en matériau élastique, p. ex. un matériau caoutchouteux
52.
DETECTION OF ELECTRICAL ARCS IN AN AERONAUTICAL SYSTEM BASED ON ELECTRICAL CURRENT AND VOLTAGE MEASUREMENTS
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (CNRS) (France)
ECOLE SUPERIEURE DE PHYSIQUE ET DE CHIMIE INDUSTRIELLES DE LA VILLE DE PARIS (France)
Inventeur(s)
Carreira Rufato, Raul
Pniak, Lucas
Lebey, Thierry Michel André
Abrégé
The invention relates to a detection device (4) for detecting the presence of an electrical arc on an electrical line (3) of an electrical circuit of an aeronautical system, comprising at least: - a measurement module adapted to obtain measurements of at least one current on said electrical line (3) and of an electrical voltage relative to said electrical line; - a computation module adapted to extract descriptors from said measurements; and the detection of the presence of an electrical arc, by applying a decision function, previously provided by a learning device (4'), to said descriptors.
G01R 31/00 - Dispositions pour tester les propriétés électriquesDispositions pour la localisation des pannes électriquesDispositions pour tests électriques caractérisées par ce qui est testé, non prévues ailleurs
G01R 31/12 - Test de la rigidité diélectrique ou de la tension disruptive
The invention relates to a rotor blade (20) for an aircraft turbine engine, this blade (20) comprising a blade airfoil (22) and a root (24) connected to the blade airfoil (22) by a platform (26), the blade airfoil (22) extending from the platform (26) to a free end (22a) comprising a recessed pocket (36), the blade airfoil (22) having a leading edge (28) and a trailing edge (30), and pressure-side and suction-side faces (32, 34) which extend from the leading edge (28) to the trailing edge (30), the blade (20) also having an internal circuit (38) for the passage of ventilation air, the internal circuit (38) extending from the root (24) to the free end (22a), the internal circuit (38) comprising an internal cavity (50) which is located at the free end (22a), just below the pocket (36), and which is connected to the bottom (44) of the pocket (36) by outlet openings (42), this cavity (50) having an elongate shape from the leading edge (28) to the trailing edge (30) and being surrounded by parts of the pressure-side and suction-side faces (32, 34).
The invention relates to a system for adjusting the pitch and de-icing the blades (14) of a propeller (12) of an aircraft, the system comprising a rotary drive element (20); a gear reduction device (24); first and second inverter devices (40) configured to be connected to an electrical power source delivering a DC voltage; a first low-radius rotary transformer (34) connected to the first inverter device (40) and delivering a first secondary AC voltage (VsAC1) for powering an electric motor (80) for adjusting the pitch of the blades of the propeller; a second low-radius rotary transformer (50) connected to the second inverter device (56) and delivering a second secondary AC voltage (VsAC2); a de-icing device (60) carried by the rotary drive element and powered by the second secondary AC voltage; and a regulation unit configured to maintain the DC voltage delivered by the electrical power source at a constant value.
Turbomachine (1) comprising a cylindrical exhaust case (8) having an axis (AX) extending along an axial direction (X) and defining an exhaust volume (Ve) through which an exhaust flow (Fe) circulates. An exhaust case arm (20) connects a frame (10) to the exhaust case (8). This exhaust case arm (20) comprises a fairing (22) which delimits an internal volume (V20) and which comprises an intake port (26) for a cooling airflow (FR) and an outlet port (52, 62) connecting the internal volume (V20) to the exhaust volume (Ve). A flow management device (70) is arranged to allow the flow of the cooling flow (FR) and to oppose a reversed flow (Finv).
The invention relates to a method for estimating a state of health of an aircraft engine, comprising a training phase during which an agent is trained by reinforcement to estimate hyperparameters for a Bayesian estimator, and a deployment phase during which the Bayesian estimator estimates a state of health of the aircraft engine using hyperparameters provided by the agent.
G06N 7/01 - Modèles graphiques probabilistes, p. ex. réseaux probabilistes
G06N 3/006 - Vie artificielle, c.-à-d. agencements informatiques simulant la vie fondés sur des formes de vie individuelles ou collectives simulées et virtuelles, p. ex. simulations sociales ou optimisation par essaims particulaires [PSO]
A method for controlling the speed of an engine assembly of an aircraft, including updating the speed setpoint of the engine assembly as a function of the current state-of-charge of an onboard energy storage device, and controlling the engine assembly according to the speed setpoint of the engine assembly, as well as a control device suitable for implementing this method and an engine assembly and an aircraft incorporating such a control device.
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
An aeronautical propulsion system includes a stator, an unducted fan section with a fan shaft extending along a longitudinal axis, and a first guiding system supporting and rotationally guiding the fan shaft. A drive shaft extends along the longitudinal axis and is driven by a turbine. A second guiding system supports and rotationally guides the drive shaft. A reduction structure couples the drive shaft to the fan shaft such that rotation of the drive shaft at a first speed drives rotation of the fan shaft at a second speed less than the first speed. A coupling device connects the drive shaft to the reduction structure and compensates for misalignment therebetween. The coupling device includes shaft sections and radial flexibility elements connecting the shaft sections. A ratio of a difference between axial distances of the radial flexibility elements from a reference plane to one of the axial distances is between 0.55 and 2.0.
A vaned impeller for aircraft turbomachine, comprising several vanes with a heel circumferentially extending on either side of the blade tip respectively into a first heel portion, and a second heel portion. Between adjacent first vane and second vane of the impeller, the latter comprises a member for reconstituting a gas circulation path, arranged in a first recess provided on the radially inner surface of the first heel portion of the first vane, and in a second recess provided on the radially inner surface of the second heel portion of the second vane, the first and second recesses each having a bottom surface for cooperating by friction with a radially outer friction surface of the member.
The invention relates to a device (300) for measuring parameters of an aerodynamic flow in a duct of a turbomachine, the device comprising: - a guide housing (200) defining an inner cavity (230), - a comb (100) comprising a comb body (120) which extends between a comb head (110) and a lower end (123) of the comb (100), the comb head (110) being slidably mounted in the inner cavity (230), the comb body (120) extending outside the guide housing (200) through the opening present in the lower part (220) of the guide housing (200), at least one spring element (130) interposed between the comb head (110) and the upper part (210) of the guide housing (200), the comb head (110) being movable in the inner cavity (230) between a first position and a second position, the comb body further comprising means (121) for measuring parameters of an aerodynamic flow.
G01M 9/06 - Dispositions pour la mesure spécialement adaptées aux tests aérodynamiques
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
G01M 15/14 - Test des moteurs à turbine à gaz ou des moteurs de propulsion par réaction
61.
TOOLING FOR LOCALIZED HEAT TREATMENT OF A MECHANICALLY WELDED COMPONENT AND ASSOCIATED METHOD FOR RELIEVING WELDING STRESSES BY MEANS OF THIS TOOLING
The invention relates to a tooling (1) for localized heat treatment of a metal component comprising a weld bead that is to be relieved of stress. The tooling comprises a clamp (2) which has a stationary jaw (4), a movable jaw (5) which can move toward or away from the stationary jaw, and an assembly (6) for guiding this movement. The stationary jaw and the movable jaw each have a shoe (15) and a clamping element (19). The two shoes are mounted facing each other and can be moved toward or away from each other, by means of their respective clamping element which pushes them toward each other when it is actuated. Each of the shoes is equipped with an abutment plate (25) and a heating element that heats the abutment plate, the two abutment plates being arranged facing each other.
C21D 8/00 - Modification des propriétés physiques par déformation en combinaison avec, ou suivie par, un traitement thermique
C21D 9/50 - Traitement thermique, p. ex. recuit, durcissement, trempe ou revenu, adapté à des objets particuliersFours à cet effet pour joints de soudure
62.
METHOD AND UNIT FOR CONTROLLING PUMPS ARRANGED IN SERIES
The invention relates to a method and an electronic control unit for controlling a set of rotodynamic pumps (1,2,3) arranged in series. This set comprises at least one upstream rotodynamic pump arranged directly upstream of a downstream rotodynamic pump. According to this method, a rotational speed of the upstream rotodynamic pump and a degree of opening of a regulating valve of the upstream rotodynamic pump are controlled as a function of a saturation margin of the downstream rotodynamic pump and of a flow rate coefficient of the upstream rotodynamic pump. The saturation margin of the downstream rotodynamic pump is a difference between the current pressure and the saturation pressure of the liquid at the inlet of the downstream rotodynamic pump. The flow rate coefficient of the upstream rotodynamic pump is the volumetric flow rate of the upstream rotodynamic pump divided by the rotational speed of the upstream rotodynamic pump.
F04D 15/00 - Commande, p. ex. régulation de pompes, d'installations ou de systèmes de pompage
F02K 9/46 - Alimentation en propergols utilisant des pompes
F02C 7/236 - Systèmes d'alimentation en combustible comprenant au moins deux pompes
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
A method for controlling the speed of an engine assembly of an aircraft. The engine assembly includes at least a combustion engine and an electric machine mechanically coupled to the combustion engine and electrically connected to an onboard energy storage device of the aircraft. The method includes steps of comparing an electrical reserve required for engine assistance with an available electrical reserve, and controlling the speed of the engine assembly as a function of the difference between the electrical power reserve required for engine assistance and the available electrical power. A control device suitable for implementing this method, as well as the engine assembly and aircraft incorporating the control device.
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64D 27/357 - Aménagements pour la production, la distribution, la récupération ou le stockage d'énergie électrique à bord utilisant des piles
B64D 31/06 - Dispositifs amorçant la mise en œuvre actionnés automatiquement
64.
POWER SHARING IN AN AIRCRAFT USING CENTRALIZED CONTROL
A system for exchanging power in an aircraft, the system including a voltage bus; a low-pressure electromechanical converter; a high-pressure electromechanical converter; and for each electromechanical converter a voltage control module and a control module for ensuring that the exchange setpoint is observed. The system further includes a power adjustment module configured, for at least one of the electromechanical converters: to determine a voltage correction; and to apply, in the voltage control module associated with the relevant electromechanical converter, the voltage correction to a voltage setpoint so that the voltage control module adjusts the bus voltage to the corrected voltage setpoint.
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
A device for setting the pitch of a vane for a turbomachine, the device including a vane that is pivotable in a first direction and includes a leading edge, a trailing edge and a base which has an opening; a fin that includes a leading edge, a trailing edge and a portion accommodated in the opening; a controller that rotates the vane 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; 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
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
66.
NAVIGATION METHOD WITH INVARIANT EXTENDED KALMAN FILTER AND VEHICLE EQUIPPED TO IMPLEMENT SAME
In a method for navigating a vehicle provided with a LiDAR and an inertial measurement unit, both connected to an electronic navigation unit that implements a first simultaneous localization and mapping algorithm and a second data merging algorithm, the first algorithm receiving, as an input, first location data from the LiDAR in order to calculate a first displacement from an initialisation position, the second algorithm receiving, as an input, the first displacement and second location data from the inertial measurement unit and outputting a merged second position of the vehicle, the second algorithm implements invariant extended Kalman filtering and in that the merged second position is introduced into a feedback loop which feeds into the first algorithm as an input in order to calculate a new position, which becomes the initialisation position.
G01S 17/931 - Systèmes lidar, spécialement adaptés pour des applications spécifiques pour prévenir les collisions de véhicules terrestres
B64U 20/80 - Disposition de l’électronique à bord, p. ex. systèmes ou câblage d’avionique
G01C 21/16 - NavigationInstruments de navigation non prévus dans les groupes en utilisant des mesures de la vitesse ou de l'accélération exécutées à bord de l'objet navigantNavigation à l'estime en intégrant l'accélération ou la vitesse, c.-à-d. navigation par inertie
G01S 17/933 - Systèmes lidar, spécialement adaptés pour des applications spécifiques pour prévenir les collisions d’aéronefs ou d'engins spatiaux
G05D 1/242 - Moyens basés sur la réflexion des ondes générées par le véhicule
G05D 1/245 - Dispositions pour déterminer la position ou l’orientation utilisant la navigation à l’estime
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 an 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 last 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.
A method for adjusting a turbine engine part to be placed in a gas flow path of the turbine engine, the part having previously undergone an operation of adding material to fill a defect in at least one zone of its surface, resulting in an excess thickness in the at least one zone of its surface, including the steps of controlling formation of a three-dimensional image of the part, identifying the at least one zone of the part surface including the excess thickness, and identifying the excess thickness, in the three-dimensional image of the part, determining a trajectory of an adjustment tool, according to the geometry of the identified excess thickness, and controlling adjustment of the part according to the determined trajectory.
An aircraft turbomachine comprising at least one propulsion shaft, an electric machine (1) connected to the propulsion shaft by a transmission line, a control device (3) comprising a correction module (4) configured to determine a correction torque (TRQcorr) to correct a torsion mode of the transmission line and a regulation module (5) configured to determine a setpoint torque (TRQcons) from a setpoint distribution voltage (Vdc_cons), a measured distribution voltage (Vdc_mes) and an angular velocity (W) of the electric machine (1) so as to ensure the quality of the electrical network (6).
H02P 21/05 - Dispositions ou procédés pour la commande de machines électriques par commande par vecteur, p. ex. par commande de l’orientation du champ spécialement adaptés pour amortir les oscillations des moteurs, p. ex. pour la réduction du pompage
H02P 23/04 - Dispositions ou procédés pour la commande de moteurs à courant alternatif caractérisés par un procédé de commande autre que la commande par vecteur spécialement adaptés pour amortir les oscillations des moteurs, p. ex. pour la réduction du pompage
70.
NOZZLE FOR A TURBINE ENGINE COMPRISING AN INNER RING PROVIDED WITH A COLLAR IN WHICH ORIFICES ARE FORMED
This nozzle (36, 36a) for a turbine comprises an inner ring (40) and a bladed ring (42) provided with a plurality of blades (46), with an inner platform (48) and with a flange (50) extending radially inwards from the inner platform (48), the inner ring (40) being provided with a platform (52), with an upstream flange (56) and with a downstream flange (58), the upstream and downstream flanges (56, 58) extending radially outwards from the platform (52) of the inner ring (40) while being attached to the flange (50) of the bladed ring (42), the inner ring (40) comprising at least one collar (62) extending at least radially outwards from the platform (52) of the inner ring (40) while remaining at a distance from the upstream and downstream flanges (56, 58), a plurality of orifices (64) being formed in the collar (62).
B29C 70/54 - Parties constitutives, détails ou accessoiresOpérations auxiliaires
B29D 99/00 - Matière non prévue dans les autres groupes de la présente sous-classe
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 70/30 - Façonnage par empilage, c.-à-d. application de fibres, de bandes ou de feuilles larges sur un moule, un gabarit ou un noyauFaçonnage par pistolage, c.-à-d. pulvérisation de fibres sur un moule, un gabarit ou un noyau
C04B 35/80 - Fibres, filaments, "whiskers", paillettes ou analogues
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
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
B29L 31/08 - Pales pour rotors, stators, ventilateurs, turbines ou dispositifs analogues, p. ex. hélices
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]
72.
BLADE MADE OF COMPOSITE MATERIAL WITH INCREASED TORSIONAL STIFFNESS AND METHOD FOR MANUFACTURING SAME
The invention relates to a blade (10) comprising a blade core (20) forming a blade root, a shank and an airfoil body portion (13), the blade core being made of composite material comprising a fibrous reinforcement having a three-dimensional weave between a first plurality of warp yarns and a first plurality of weft yarns oriented respectively in the longitudinal direction and in the transverse direction. The blade (10) further comprises skins (30, 31) made of composite material comprising a fibrous reinforcement having a three-dimensional weave between a second plurality of warp yarns oriented mainly in a first direction (D1) forming an angle with the longitudinal direction (DL) of between ± 10° and ± 80° and a second plurality of weft yarns oriented mainly in a second direction (D2) forming an angle with the transverse direction (DT) of between ± 10° and ± 80°.
One aspect of the invention relates to a fastening collar (8) for fastening a cooling tube (2) of a cooling system of a turbine engine casing, the fastening collar (8) comprising a main body (10) having: - a circular portion (16) configured to surround the cooling tube over at least part of its circumference, and - two planar portions (19, 20) extending in the same plane (L, T), each arranged at one end of the circular portion. The fastening collar comprises a retaining ring (40) made of a graphite material and inserted into the main body, and a means for preventing the retaining ring from moving longitudinally in the main body. A second aspect of the invention relates to an assembly (4) for fastening a cooling tube, the assembly (4) comprising a fastening support (6) secured to the turbine engine casing, a fastening collar according to the invention, and a clamping device (60) for clamping the collar to the support.
The invention relates to a method for manufacturing a turbine engine part, which comprises: - producing a fibrous blank (100) of a structure with an aerodynamic profile by three-dimensional weaving of yarns, the blank comprising an internal housing (106), - inserting a shaping piece (400) into the internal housing (106) of the fibrous blank so as to form a preform, - densifying the fibrous preform with a matrix. The shaping piece (400) consists of a plurality of unit blocks (401-411) introduced one after another into the internal housing (106).
The invention relates to a method for manufacturing a blade or a propeller, the method comprising: - producing a fibrous blank (100) of a structure with an aerodynamic profile by three-dimensional weaving of yarns, the blank comprising an internal housing (106), - inserting a shaping piece (400) into the internal housing (104a) of the fibrous blank so as to form a preform, - densifying the fibrous preform with a matrix. The foam shaping piece (400) extends in the longitudinal direction in the aerodynamic profile portion over a height (H400) of between 20% and 80% of the height (H110) of the aerodynamic profile portion. The foam shaping piece further comprises recesses (410, 420a, 420b) each provided on at least one of the first or second outer faces (400a, 400b) of the foam shaping piece (400), the one or more recesses being filled with a material having a higher density than that of the foam of the shaping piece.
B29C 70/86 - Incorporation dans des couches de renforcement imprégnées cohérentes
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
The invention relates to a method for controlling an accelerometric sensor, the accelerometric sensor having electrodes forming capacitors of variable capacitance depending on a distance between the electrodes, a control unit being arranged to carry out a capacitance measurement operation and a control operation which selectively comprises: a fine control phase in which a first voltage is applied between one of the fixed electrodes and the movable electrode, the other of the fixed electrodes being at the same potential as the movable electrode; and an extended control phase in which a second voltage is applied between one of the fixed electrodes and the movable electrode, the other of the fixed electrodes being at the same potential as the movable electrode and the second voltage being greater in absolute value than the first voltage. The control unit is arranged to implement self-calibration. The invention also relates to a method using such a sensor.
G01P 21/00 - Essai ou étalonnage d'appareils ou de dispositifs couverts par les autres groupes de la présente sous-classe
G01P 15/13 - Mesure de l'accélérationMesure de la décélérationMesure des chocs, c.-à-d. d'une variation brusque de l'accélération en ayant recours aux forces d'inertie avec conversion en valeurs électriques ou magnétiques en mesurant la force nécessaire pour remettre à sa position de repos une masse d'épreuve soumise aux forces d'inertie
G01P 15/125 - Mesure de l'accélérationMesure de la décélérationMesure des chocs, c.-à-d. d'une variation brusque de l'accélération en ayant recours aux forces d'inertie avec conversion en valeurs électriques ou magnétiques au moyen de capteurs à capacité
78.
ASSEMBLY BY METAL SINTERING WITH CONTROL OF THE CHIP HEIGHT IN ORDER TO CONTROL FLATNESS AND ATTITUDE
The invention relates to a method (100) for assembling an electronic component (1) on a substrate (2), the method comprising the following steps - depositing a sintering material (7) on one face of a substrate (2), so as to form at least three columns of the sintering material; - stacking the electronic component (1) on the columns by means of a gripping tool (22) designed to place the electronic component on the columns at a predetermined height (h) from the face of the substrate (2); - heating the stack formed at least by the substrate (2), the columns and the electronic component (1) in order to sinter the columns, the at least three columns solidifying to form at least three sintered studs, the bond line thickness of which is related to the predetermined height (h).
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 baking (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.
A system for exchanging power in an aircraft includes voltage bus, a low-pressure and high-pressure electromechanical converters, and for each electromechanical converter, a voltage control module. The system further includes a computer providing a sharing setpoint between the exchanged powers, the sharing setpoint varying over time; and for at least one of the electromechanical converters: a module that ensures the sharing setpoint is observed. The module is configured to adjust the power that is exchanged by the relevant electromechanical converter in order to observe the received sharing setpoint, taking into account a received evaluation of the power exchanged by the other of the electromechanical converters.
B64D 31/18 - Systèmes de commande des groupes moteursAménagement de systèmes de commande des groupes moteurs sur aéronefs pour les groupes moteurs électriques pour les groupes moteurs hybrides-électriques
B64D 35/022 - Transmission de la puissance des groupes moteurs aux hélices ou aux rotorsAménagements des transmissions spécialement adaptés à des groupes moteurs spécifiques aux groupes moteurs électriques du type électrique-hybride
H02J 1/10 - Fonctionnement de sources à courant continu en parallèle
A system for measuring a clearance between a rotating blade and a casing of a turbomachine includes a removable fastening device provided with a locking mechanism to maintain the mounted position of the measurement system on the casing, a retractable stop mechanism provided with a guide and a finger that hold the rotating blade in a measurement position, a remote acquisition device that measures at least one clearance between the rotating blade and the casing when the removable fastening device is in the mounted position and the retractable stop mechanism is in the deployed position, and a plate to which the removable fastening device, the retractable stop mechanism and the remote acquisition device are attached.
G01B 11/14 - Dispositions pour la mesure caractérisées par l'utilisation de techniques optiques pour mesurer la distance ou la marge entre des objets ou des ouvertures espacés
82.
AERONAUTICAL PROPULSION SYSTEM COMPRISING AN OPTIMIZED FAN SECTION
A fan section of a propulsion system includes a fan rotor including twenty-two blades and having a solidity greater than or equal to 0.9 and less than or equal to 1.3, where the solidity is equal to a ratio between a chord at the tip of a blade and an inter-blade pitch at the tip of the blade. The fan section has a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5, and a peripheral speed at the tip of the blade greater than or equal to 270 m/s and less than or equal to 380 m/s. The pressure ratio and the peripheral speed are measured at cruising power.
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
This pitch-change mechanism (70) comprises 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 biasing 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.
This pitch-change mechanism (70) comprises 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 biasing 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.
An aeronautical propulsion system includes a ducted fan section including a fan, the fan having a fan rotor including blades, and a gas generator including a drive shaft that rotationally drives, directly or indirectly, the fan rotor. The aeronautical propulsion system is dimensioned such as to verify two relationships governing a diameter of the fan rotor, a maximum thrust generated by the fan, and a bypass ratio of a flow through the aeronautical propulsion system.
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
B64D 27/18 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à réaction à l'intérieur des ailes ou fixés à celles-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
The invention relates to a part (9) for an aircraft turbine engine, the part (9) comprising: - a body (10) having a surface (10a) intended to be swept by an airflow, and - a multilayer coating (11) arranged on the surface (10a) of the body (10), characterized in that the coating (11) comprises: - a first electrically insulating layer (12), - a second layer (13) arranged on the first layer (12), the second layer (13) comprising an electrically conductive circuit, and - a third layer (14) arranged on the second layer (13) and comprising a polymeric matrix and piezoelectric fillers, the piezoelectric fillers being deformable under the effect of an electric current transmitted by the second layer (13).
B64D 15/16 - Dégivrage ou antigivre des surfaces externes des aéronefs par dispositifs mécaniques, p. ex. des gaines ou des bourrelets fixés ou incorporés à la surface et soumis à des pulsations
H10N 30/85 - Matériaux actifs piézo-électriques ou électrostrictifs
87.
HOLLOW BLADE MADE OF A METAL ALLOY AND HAVING A MONOCRYSTALLINE STRUCTURE FOR A TURBINE ENGINE, AND METHOD FOR MANUFACTURING SAME
The invention relates to a hollow blade (2) made of a metal alloy and having a monocrystalline structure for a turbine engine, the blade (2) comprising: - an outer wall (210) comprising a leading edge (21), a trailing edge (22), and a pressure-side face and a suction-side face which connect the leading edge (21) to the trailing edge (22), - an inner wall (212) delimiting an inner cavity (200) which comprises at least one cooling ramp extending from a root of the blade to its tip (26), and - a series of cooling slots (220) located on the trailing edge (22), each cooling slot (220) being in fluidic communication with at least a portion of the inner cavity (200), wherein the hollow blade (2) further comprises micro-grooves (230) on at least one portion of the inner wall (212), these micro-grooves (230) having a depth of between 0.01 mm and 0.1 mm.
A system and method for determining a landing zone for a drone, comprising obtaining images representing at least a search perimeter for a landing zone; obtaining 3D information representing at least said search perimeter; selecting a potential landing zone within said search perimeter by analysing the obtained images and 3D information; obtaining data relating to the moisture content and/or hardness of the ground in the potential landing zone, acquired by means of at least one contact sensor; and determining a measure of suitability of the ground in the potential landing zone for landing the drone, based on the data relating to the moisture content and/or hardness of the ground, and optionally the images and/or 3D information.
G05D 1/241 - Moyens pour détecter un contact physique, p. ex. capteurs tactiles ou capteurs de choc
G05D 1/242 - Moyens basés sur la réflexion des ondes générées par le véhicule
G05D 1/243 - Moyens de capture de signaux provenant naturellement de l’environnement, p. ex. signaux optiques, acoustiques, gravitationnels ou magnétiques ambiants
The present disclosure relates to a device for holding an air jet cooling tube on a turbomachine casing, the device comprising: - an external support adapted to be arranged around a turbomachine casing, - a clamp adapted to surround a tube and hook it to the support, and - a graphite ring intended to be positioned around the tube, between the clamp and the tube, the ring comprising two half-rings each intended to surround a portion of the tube.
The present disclosure relates to a device for securing an air-jet cooling tube to a turbine engine casing, the device comprising: - an external support designed to be arranged around a turbine engine casing, - a collar designed to surround a tube and fasten it to the support, and - a graphite ring intended to be positioned around the tube, between the collar and the tube, the ring comprising three ring portions each intended to surround a portion of the tube, the ring surrounding an angular portion of between 240° inclusive and 300° inclusive of the circumference of the tube.
Tooling (100) for fixing a metal reinforcement (110) to the leading edge of a turbomachine blade (140), the metal reinforcement comprising a nose (111) intended to form the leading edge of the blade and flanges (112, 113) extending from the nose of the metal reinforcement and intended to extend from the leading edge towards the pressure-side and suction-side faces of the blade, characterised in that the tooling comprises a reinforcement support (120) intended to receive the metal reinforcement, while leaving the nose free (160) and being in contact with the flanges of the metal reinforcement, said reinforcement support comprising at least one notch (151) intended to receive a strap (150).
F01D 5/28 - Emploi de matériaux spécifiésMesures contre l'érosion ou la corrosion
B23P 15/04 - Fabrication d'objets déterminés par des opérations non couvertes par une seule autre sous-classe ou un groupe de la présente sous-classe d'aubes de turbine ou d'organes équivalents, en plusieurs pièces
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
B29D 99/00 - Matière non prévue dans les autres groupes de la présente sous-classe
92.
FIBROUS TEXTURE FOR A ROTATING PART MADE OF A COMPOSITE MATERIAL WITH INCREASED TORSIONAL STIFFNESS
The invention relates to a fibrous texture (100) having a three-dimensional weave formed by a plurality of layers of warp threads extending in a longitudinal direction (DL) and a plurality of layers of weft threads extending in the lateral direction (DT), the fibrous texture comprising an end portion (T4) that extends over a given length (LT4) in the longitudinal direction, between an intermediate position that lies between proximal and distal ends (103, 104), and the distal end (104). The portion comprises a plurality of non-interlinked regions (160, 161, 162, 163, 164, 165) each forming an inner cavity (170, 171, 172, 173, 174, 175) in the fibrous texture, each non-interlinked region extending between the first and second lateral edges (101, 102) of the fibrous texture in a direction that forms an angle of between ±40° and ±50° with the longitudinal direction (DL).
A method for manufacturing a turbomachine part comprises - forming a fibrous blank (100) of an aerodynamic profile structure by three-dimensional weaving of yarns, said blank comprising an internal cavity (106), - inserting a shaping part (400) into the internal cavity (106) of the fibrous blank so as to form a preform, - densifying the fibrous preform with a matrix. The shaping part (400) further comprises recesses (410, 420, 430, 440, 450) each extending between first and second external faces of said foam shaping part, the recess or recesses being filled with the matrix.
One aspect of the invention relates to a fastening collar (8) for fastening a cooling tube (2) of a cooling system of a turbomachine casing, the fastening collar (8) comprising a main body (10) having: a circular portion (16) configured to surround the cooling tube over at least part of its circumference, and two planar portions (19, 20) extending in the same plane (L, T), each arranged at one end of the circular portion. The fastening collar comprises a retaining ring (32) made of a refractory material, the retaining ring (32) being inserted into the main body and bearing against an internal surface of the circular portion. A second aspect of the invention relates to an assembly (4) for fastening a cooling tube, the assembly (4) comprising: a fastening support (6) secured to the turbomachine casing, a fastening collar (8) according to the invention, and a clamping device (46) for clamping the collar to the support.
G05D 1/695 - Commande coordonnée de la position ou du cap de plusieurs véhicules pour maintenir la position des véhicules les uns par rapport aux autres, p. ex. pour les voyages en convoi ou les vols en formation
A system for powder bed fusion additive manufacturing includes build chamber extending along a vertical axis and a removable container capable of being inserted into the build chamber in a manufacturing configuration and of being removed from the build chamber in a depowdering configuration. The removable container includes a side wall extending along the vertical axis and a bottom wall capable of movement along the vertical axis inside the side wall and which includes a perforated first wall comprising first orifices configured to be closed off in the manufacturing configuration and unblocked in the depowdering configuration.
An aeronautical propulsion system includes a fan rotor to be driven by a fan shaft, a drive turbine that drives the fan rotor through a drive shaft about an axis of rotation, a reduction structure that couples the drive shaft and the fan shaft to drive the fan shaft at a rotation speed less than a rotation speed of the drive shaft, and bearings that center the drive shaft with respect to the axis of rotation and that include one front bearing extending upstream of a combustion chamber of the propulsion system and two rear bearings extending downstream of the combustion chamber. The position of the bearings in the propulsion system obeys a formula based on various parameters related to the bearings, the drive turbine, and the drive shaft.
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
98.
TOOLING AND METHOD FOR VISUALLY INSPECTING AN AIRCRAFT TURBOMACHINE PART
Tooling is configured for visually inspecting an aircraft turbomachine part that has an annular shape about an axis (X) and an annular attachment flange oriented in a radial direction with respect to the axis (X). The tooling includes an integrally formed body and a U-shaped hook configured to be mounted on the flange. The hook has two lateral branches configured to be located on both sides of the flange, respectively, and a frame connected to the hook and configured to cover a section of the part to be inspected. The frame defines a window through which an operator can view a portion of the section.
Navigation device including an inertial measurement unit and an electronic navigation calculation unit that are connected to each other by a data link, the inertial measurement unit having inertial sensors for producing location signals, and the electronic navigation calculation unit arranged to calculate a navigation from the location signals. The inertial measurement unit has an electronic processing circuit connected to the inertial sensors and with a memory containing an error model of the inertial measurement unit. The electronic processing circuit is arranged to transmit to the electronic navigation unit, firstly, at a first rate, said location signals and, secondly, at a second rate, correction signals having first correction data, which are representative of an impact of the sensor errors on the location signals over a given period and which are determined by the electronic processing circuit on the basis of the error model, and second correction data, which are representative of an effect of the random noises and errors of the inertial measurement unit over the given period. The electronic navigation calculation unit implements a hybrid navigation algorithm arranged to supply a hybrid location that is based on inertial location data extracted from the location signals and from the external location data and that is readjusted on the basis of the correction data.
G01C 21/28 - NavigationInstruments de navigation non prévus dans les groupes spécialement adaptés pour la navigation dans un réseau routier avec corrélation de données de plusieurs instruments de navigation
A control system for a battery-powered aircraft, the control system including a landing gear assembly moveable between a retracted position in which the landing gear assembly is housed within the aircraft and an extended position in which the landing gear assembly is deployed, the landing gear assembly including a main landing gear; and a nose landing gear; at least one local hydraulic pressure source configured to receive power from at least one battery; a landing gear controller configured to receive and deliver control inputs to the hydraulic pressure source; and at least one actuator in fluid communication with the hydraulic pressure source and configured to move the landing gear assembly between the retracted position and the extended position in response to inputs received by the landing gear controller; wherein at least one of the main landing gear and nose landing gear is biased in the retracted position by hydraulic pressure.
