A method for controlling a hybrid turbomachine includes determining a sequence of deviations, each being a difference between a measurement of a sequence of measurements of a speed of a low-pressure shaft of the hybrid turbomachine and a setpoint of a sequence of setpoints of the speed of the low-pressure shaft, determining a variation of fuel flow rate and determining a variation of electrical torque, the electrical torque being configured to be applied to the low-pressure shaft, and controlling the hybrid turbomachine in order to vary a fuel flow rate according to the determined variation of fuel flow rate and in order to vary an electrical torque according to the determined variation of electrical torque.
F02C 9/28 - Systèmes de régulation sensibles aux paramètres ambiants ou à ceux de l'ensemble fonctionnel, p. ex. à la température, à la pression, à la vitesse du rotor
B64D 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
2.
AERO-PROPULSION SYSTEM WITH IMPROVED PROPULSION EFFICIENCY
A propulsion system includes a drive shaft movable in rotation about an axis of rotation, a fan shaft, a fan section comprising a ducted fan rotor driven in rotation by the fan shaft, the fan rotor including plural blades, and a reduction mechanism that couples the drive shaft and the fan shaft in order to drive the fan shaft at a rotation speed lower than the rotation speed of the drive shaft. A thrust density per blade of the fan rotor of the propulsion system being greater than or equal to 1.40×105 and less than or equal to 1.70×105 N/m2 where the thrust density per blade is defined based on a thrust generated by the fan rotor measured when the propulsion system is stationary in take-off rating and a diameter of the fan rotor.
A method for controlling a hybrid turbine engine includes determining a sequence of deviations, each being a difference between one measurement of a sequence of measurements of a speed of a drive shaft of the engine and one setpoint of a sequence of engine speed setpoints, determining a variation of a fuel flow rate and determining a variation of an electrical torque, the determining the variation of fuel flow rate and the determining the variation of electrical torque both use a difference between a last measurement and a penultimate measurement of the sequence of measurements, a last deviation of the sequence of deviations that is temporally associated with the last measurement, and a sum of the deviations of the sequence of deviations, and controlling the hybrid turbine engine in order to vary the fuel flow rate and the electrical torque according to the determined variations.
Controlling a turbomachine including: the detection of a speed of a drive shaft of the turbomachine below an idle speed; the detection of a saturation of the fuel setpoint; the selection of a combined torque and fuel regulation loop from among at least one single-variable engine regulation loop and, the determination, in a combined manner, of a value representative of a fuel setpoint and of a value representative of a torque setpoint such that the torque setpoint compensates for the saturation of the fuel setpoint to allow the maintenance of a steady-state speed of the turbomachine.
Valve for an aircraft turbomachine, including a seat defining an air inlet, a shutter movable relative to the seat between a closed position in which the shutter at least partly closes the air inlet, and an open position in which the shutter closes the air inlet less than in the closed position, a frame fixed relative to the seat, the frame defining a housing and a recovery chamber in communication with each other, and a plug located in the housing, the plug being fusible and configured, in the solid state, to hold the shutter in a first position among the closed position and the open position, and in the molten state, flow towards the recovery chamber.
F16K 31/00 - Moyens de fonctionnementDispositifs de retour à la position de repos
F01D 21/12 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs sensibles à la température
F16K 17/40 - Soupapes ou clapets de sûretéSoupapes ou clapets d'équilibrage avec élément de rupture, p. ex. un diaphragme de rupture, un joint fusible
6.
TURBOMACHINE COMPRISING AN ELECTRIC MACHINE PROVIDED WITH A FIXED WINDING, AN EXCITATION WINDING AND A ROTARY WINDING WHICH ARE ARRANGED RADIALLY
A turbomachine for an aircraft includes one part rotating about an axis of rotation and one part fixed in rotation about the axis of rotation relative to the rotating part. The fixed part includes a fan casing and one electric machine that operates in generator mode. The electric machine includes at least one fixed winding and one excitation winding, which are positioned on the fixed part of the turbomachine, and at least one rotating winding positioned on the rotating part of the turbomachine. The fixed winding, the excitation winding and the rotating winding are arranged radially about the axis of rotation.
The invention relates to a rotor element (48) for a turbine engine, comprising a metal disk (62) and a plurality of blades (49) mounted on the disk (62). At least one of said blades (49) consists of a hybrid blade (79) made of composite material and comprising a metal root (80) linked to the disk (62).
A turbomachine for an aircraft extends along an axis of rotation and includes one part rotating about the axis of rotation and one part fixed in rotation about the axis of rotation relative to the rotating part. The fixed part includes a fan casing. The turbomachine includes at least one permanent magnet synchronous electric machine that includes at least one permanent magnet positioned on the fixed part of the turbomachine and at least one winding positioned on the rotating part of the turbomachine. The at least one permanent magnet is arranged radially around the at least one winding.
H02K 21/28 - Moteurs synchrones à aimants permanentsGénératrices synchrones à aimants permanents avec des induits tournants et des aimants fixes avec des induits tournant à l'intérieur des aimants
9.
PITCH-CHANGE MECHANISM WITH CANTILEVERED PITCH-LOCKING DEVICE
The invention relates to a pitch-change mechanism (70), which comprises a frame (72), a control actuator (74) including a stationary portion (100) rigidly connected to the frame (72) and a movable portion (102) translatable along a longitudinal axis (X), a connection system (78) connecting the movable portion (102) to a variable-pitch blade (56) so as to convert the translation of the movable portion (102) into a rotation of the variable-pitch blade (56), and a pitch-locking device (160) capable of blocking the translation of the movable portion (102) relative to the stationary portion (100) in at least one direction. The pitch-locking device (160) is cantilevered longitudinally relative to the frame (72).
The invention relates to a pitch-change mechanism (70), which comprises a frame (72) and an actuating cylinder (74) including a movable portion (102) that is translatable relative to the frame (72). The mechanism also comprises a connection system (78) capable of converting a translation of the movable portion (102) in a first direction into a rotation of the variable-pitch blade (56) towards the flat position, and a translation of the movable portion (102) in a second direction into a rotation of the variable-pitch blade (56) towards the feathered position. This connection system (78) comprises a first hinge (144) rigidly 10 connected to the movable portion (102) and a second hinge (146) rigidly connected to the variable-pitch blade (56). The first hinge (144) is arranged upstream of the second hinge (146), the first direction going from upstream to downstream.
A propulsion system includes a drive shaft movable in rotation about an axis of rotation, a fan shaft, a fan section including an unducted fan rotor driven in rotation by the fan shaft, the fan rotor including plural blades, and a reduction mechanism coupling the drive shaft and the fan shaft in order to drive the fan shaft at a rotation speed lower than the rotation speed of the drive shaft. A thrust density per blade of the fan rotor of the propulsion system being greater than or equal to 5.0×104 N/m2 and less than or equal to 10.0×104 N/m2, where the thrust density per blade is defined based on a thrust generated by the fan rotor and measured when the propulsion system is stationary in take-off rating, a number of blades in the fan rotor, and a diameter of the fan rotor in meters.
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnelDisposition permettant la dilatation calorifique ou le déplacement
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
12.
AIRCRAFT TURBINE ENGINE COMPRISING AN ELECTRIC MOTOR COMPRISING A DEVICE FOR CORRECTING A PRECESSION MOVEMENT AND ASSOCIATED METHOD
An aircraft turbine engine including at least a propulsion shaft and an electric motor. A rotor is rigidly connected to the propulsion shaft. The electric motor includes at least three power channels, a control device designed to determine a control parameter for each power channel on the basis of a control command, and a correction device designed to correct a precession movement of the rotor relative to the stator. The correction device being designed to determine the control command of the control device from a setpoint command and a precession level. The control command being designed to generate a global magnetic force opposing the precession movement of the rotor so as to damp said movement.
H02P 9/00 - Dispositions pour la commande de génératrices électriques de façon à obtenir les caractéristiques désirées à la sortie
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
H02K 7/18 - Association structurelle de génératrices électriques à des moteurs mécaniques d'entraînement, p. ex. à des turbines
13.
AERONAUTICAL PROPULSION SYSTEM WITH IMPROVED PROPULSION EFFICIENCY
A propulsion system includes a drive turbine connected to a drive shaft movable in rotation about an axis of rotation, a fan section including a fan rotor connected to a fan shaft, and a reduction structure coupling the drive shaft and the fan shaft in order to drive the fan shaft at a rotation speed lower than the rotation speed of the drive shaft. In the propulsion system, a rotation speed of the drive shaft complies with a formula defining a relationship among the rotation speed of the drive shaft, an inlet temperature of the drive turbine when the propulsion system is stationary in a take-off rating in a standard atmosphere and at sea level, and an inlet section of the drive turbine, in square meters.
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
Aeronautical thruster having a longitudinal axis includes an upstream rotor row and a downstream stator row of unducted blades. Two adjacent blades of the downstream stator row have an azimuthal spacing between them, about the longitudinal axis, and at least some of these spacings differ from one another such that at least some of the blades of the downstream stator have a heterogeneous distribution about the longitudinal axis.
An assembly for a turbomachine of longitudinal axis includes a nozzle carried by an annular part which is fixed to an outer casing. The nozzle has an outer annular platform from which extends, radially outwards, a radial annular wall which carries an annular attachment tab extending longitudinally upstream from the radial annular wall. The assembly has a cooling circuit with an inner annular chamber delimited radially by the outer annular platform and the annular part and delimited downstream by the radial annular wall. An intermediate annular chamber is formed radially externally to the annular attachment tab and in fluid communication with the outer casing. Means of fluid communication between the inner annular chamber and the intermediate annular chamber is formed in the annular attachment tab of the nozzle.
An inlet cone for an aircraft turbine engine is configured to be rotated about an axis (X). The inlet cone includes a frustoconical body and a tip made of elastically deformable material fixed to a smaller-diameter end of the body. The body includes through-housings distributed around the axis, and parts made of elastically deformable material located in the through-housings. Each of the parts made of elastically deformable material has at least one peripheral flange for retaining the part in the corresponding through-housing. The peripheral flange bears against an internal surface of the body.
F02C 3/04 - Ensembles fonctionnels de turbines à gaz caractérisés par l'utilisation de produits de combustion comme fluide de travail ayant une turbine entraînant un compresseur
A turbomachine module, including fan comprising variable pitch fan vanes, a speed reducer connected to a fan shaft which is guided in rotation by two first and a second rolling guide bearings, a pitch change system for the vanes including a controller including an annular body integral with the fan shaft and a movable body, a fluid transfer device intended to feed the control means. The speed reducer has a planetary gear train, and the transfer device is arranged upstream of the speed reducer and includes a stator portion and a rotor portion engaged in the stator portion, the rotor portion being integral in rotation with the annular body of the controller which is arranged upstream of the first and second guide bearings.
A turbine engine rotary assembly includes a disc and at least one blade including a stilt wall extending radially from a root situated in a recess of the disc including a first circumferential end surface, a platform which extends at the radially outer end of the stilt and is mounted circumferentially end-to-end with another stilt of another circumferentially adjacent blade, the platform or the stilt including a first coupling element, the assembly also including a sealing member mounted in a cavity including a second coupling element nested with that of the blade for axially retaining the sealing member in the cavity in both directions.
A bearing assembly includes a shaft having an axis of rotation and a hub mounted around the shaft. The bearing also includes a bearing having an inner bearing race attached to the shaft, an outer bearing race attached to the hub, rolling elements arranged between the inner bearing race and the outer bearing race, and a protective device suitable for protecting the bearing from overloads. The protective device includes a first ring and a second ring borne by a support from among the shaft and the hub, the first ring being arranged at a distance from the second ring when the bearing assembly is subjected to a load lower than a threshold load; the first ring making contact with the second ring when the bearing assembly is subjected to a load higher than said threshold load.
F16C 19/06 - Paliers à contact de roulement pour mouvement de rotation exclusivement avec roulements à billes essentiellement du même calibre, en une ou plusieurs rangées circulaires pour charges radiales principalement avec une seule rangée de billes
20.
Process for bonding a heating mat to a turbomachine blade
A process for bonding an elastomer multilayer with integrated heating system and a thermosetting elastomer surface layer to at least one of the surfaces of a composite and/or titanium-based alloy substrate by an epoxy adhesive film having a curing temperature enabling complete crosslinking, of less than or equal to 126° C.
A gas turbine engine including a turbomachine. The turbomachine including a compressor section, a combustion section, and a turbine section arranged in serial flow order. The turbomachine including an inlet duct defining an engine inlet, a fan duct defining a fan duct inlet, and a core duct defining a core duct inlet. The gas turbine engine including a fan cowl disposed circumferentially around the turbomachine and a heat exchanger defining a portion of the fan duct. At least a portion of the heat exchanger formed integrally with or coupled to the fan cowl.
F02C 7/14 - Refroidissement des ensembles fonctionnels des fluides dans l'ensemble fonctionnel
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnelDisposition permettant la dilatation calorifique ou le déplacement
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
A seal configured to be interposed between a first and a second element of a propulsion assembly in an assembled configuration, the seal having a segment with an irregular reinforcement extending over at least a part of the length of the seal and which has a contour comprising a first contour part provided with a first reinforcement part, and a second contour part in which said first reinforcement part is absent, such that the second contour part has an increased susceptibility to buckling by comparison with the first contour part.
Turbine for a turbomachine comprising an annular casing (1), upstream (3) and downstream (4) ring supports and a sealing ring (2) configured to extend around and radially face a blade assembly (6) of a rotor of the turbine,
the supports are mounted against a radially inner surface of the casing (1),
the sealing ring (2) being configured to be disposed radially between the blade assembly (6) and the casing (1), and being mounted against the upstream support and the downstream support,
the sealing ring (2) having a maximum radius with respect to the main axis (X) of the turbine, and
the casing (1) being dimensioned such that the maximum radius of the sealing ring (2) is strictly smaller than the radius of the upstream end of the casing (1) so as to allow a mounting by axial insertion, of the sealing ring (2) and of the upstream ring support (3), against the downstream ring support (4).
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
A turbine ring assembly provided with a segmented turbine ring made of CMC and a ring support structure, each segment including a base from which extend, radially outward, an upstream attaching lug and a downstream attaching lug axially spaced apart from one another, the support structure including an upstream radial flange and a downstream radial flange between which are retained the upstream attaching lugs. The assembly further includes, at each join between two adjacent ring segments along the circumferential direction a first cylindrical sealing element extending mainly along the axial direction and disposed between the bases of the two adjacent ring segments, a second sealing element extending radially along the downstream attaching lug from the cylindrical seal forming a sealed connection, and a third sealing element extending radially along the downstream attaching lug from the cylindrical seal forming a sealed connection.
A variable-pitch vane made of a composite material for an unducted fan of an aircraft, the vane including a fibrous reinforcement formed as a single piece of three-dimensional fabric obtained by three-dimensional weaving including an airfoil portion and a root portion forming a retaining bulb and intended to be connected to a variable-pitch mechanism, the three-dimensional weaving being carried out continuously between the root portion and the airfoil portion, the fibrous reinforcement defining an internal shaping cavity formed by a separation extending inside the root portion and the blade portion, wherein a matrix densifies the fibrous reinforcement and reinforcing metal shells are arranged around the root portion.
B32B 5/02 - Produits stratifiés caractérisés par l'hétérogénéité ou la structure physique d'une des couches caractérisés par les caractéristiques de structure d'une couche comprenant des fibres ou des filaments
B32B 15/14 - Produits stratifiés composés essentiellement de métal adjacent à une couche fibreuse ou filamenteuse
B32B 27/38 - Produits stratifiés composés essentiellement de résine synthétique comprenant des résines époxy
A seal configured to provide a predefined clearance between the seal and an outer surface of a rotor mounted rotatably about an axis A disposed opposite the seal, the seal being characterized in that the inner surface of each inner ring segment includes a plurality of patterns hollowed out from the inner surface of the inner ring segment, the plurality of patterns including at least:
a first pattern having a longilineal shape extending along a diagonal direction with respect to the axial direction; and
a second pattern including a first part extending along the same diagonal direction as the first pattern, and a second part extending along the circumferential direction.
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages par obturation non contact, p. ex. du type labyrinthe
The invention relates to a cooling-air injection casing comprising a casing upstream end, a casing downstream end, and a casing main wall and further including: an air mixing cavity; an air intake cavity; an air bleeding cavity; a first sealing gasket; a second sealing gasket; the injection casing having further an air passage cavity bounded by a substantially radial wall, the casing main wall and the substantially axial wall, the air passage cavity being in fluid communication with the air mixing cavity via apertures made in the substantially axial wall, and in that the air passage cavity is in fluid communication with the air bleeding cavity by means of apertures made in the substantially radial wall.
A sealing assembly for a turbomachine including a sealing gasket configured to ensure a predefined clearance between the sealing gasket and an external surface of a rotor mounted in rotation around an axis A, the axis A defining an axial direction, the sealing gasket extending circumferentially around the axis A and the external surface being arranged facing the sealing gasket, the sealing assembly being characterized in that it further includes an air straightening member including at least one row of blades which extend radially, the air straightening member being located upstream of the sealing gasket, the air straightening member being configured to straighten the air intended to pass through the sealing gasket.
F01D 11/02 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages par obturation non contact, p. ex. du type labyrinthe
29.
METHOD AND MODULE FOR DETECTING THE STATE OF A COUPLING DEVICE, AND ASSOCIATED TURBINE ENGINE AND AIRCRAFT
This method for controlling the operation of a coupling device for an aircraft turbomachine comprises the steps of:
a first comparison of the determined rotational speed of an electric machine with a time evolution profile, the first comparison being carried out to determine an operating deviation;
a second comparison of the operating deviation with a detection threshold; and
according to the result of the second comparison step, identifying the existence or absence of a failure of the coupling device;
the second comparison step comprising determining the value of the detection threshold from at least one control parameter, the operating deviation and the rotational speed of the electric machine.
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
B64D 45/00 - Indicateurs ou dispositifs de protection d'aéronefs, non prévus ailleurs
F01D 15/00 - Adaptations des "machines" ou machines motrices à des usages particuliersCombinaisons des machines motrices avec les dispositifs qu'elles entraînent
A gasket configured to ensure a predefined clearance between the gasket and an outer surface, which includes a plurality of gasket sectors distributed circumferentially around the axis A, each gasket sector having an inner ring sector connected to an outer ring sector by a return member, where the radially inner surface of each inner ring sector includes at least one pattern hollowed out from the radially inner surface of the inner ring sectors, the inner ring sector further including at least one channel connecting a first opening which opens onto the upstream face of the inner ring sector and a second opening which opens into one of the patterns hollowed out from the radially inner surface of the inner ring sectors.
F01D 11/08 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages pour obturations de l'espace entre extrémités d'aubes du rotor et stator
B64D 27/10 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à turbine à gaz
A sealing device including a ring having a main sealing lip and a secondary sealing lip, where the main sealing lip presents an inner surface and the secondary sealing lip presents a radial surface, the sealing device also including a casing having a cylindrical portion and a radial portion, where the radial portion presents a contact surface, the ring being biased so as to apply the radial surface of the secondary sealing lip against the contact surface of the casing, where the sealing device presents at least one leakage path extending between the secondary sealing lip and the casing.
F01D 11/00 - Prévention ou réduction des pertes internes du fluide énergétique, p. ex. entre étages
F01D 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
F16J 15/16 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre
F16J 15/30 - Joints d'étanchéité entre deux surfaces mobiles l'une par rapport à l'autre par presse-étoupes pour bagues d'étanchéité rigides les bagues d'étanchéité étant en graphite
A tooling for manufacturing a blade made of composite material for a turbomachine, in particular of an aircraft, including a mould and a counter-mould which define between them a cavity configured to receive a woven preform, the cavity having a first part configured to receive a shield and at least one edge of the preform, and a second part configured to receive at least part of the remainder of the preform, and elements for managing the temperature of the cavity, which elements are configured to heat the first and second parts of the cavity at different temperatures during at least one step of a method for manufacturing the blade.
B29C 70/48 - Façonnage ou imprégnation par compression pour la fabrication d'objets de longueur définie, c.-à-d. d'objets distincts utilisant des moules opposables, p. ex. pour déformer des préimprégnés [SMC] ou des "prepregs" avec une imprégnation des renforcements dans le moule fermé, p. ex. moulage par transfert de résine [RTM]
B29C 33/02 - Moules ou noyauxLeurs détails ou accessoires comportant des moyens incorporés de chauffage ou de refroidissement
B29L 31/08 - Pales pour rotors, stators, ventilateurs, turbines ou dispositifs analogues, p. ex. hélices
33.
METHOD AND DEVICE FOR DETECTING AN ANOMALY IN THE OPERATION OF AN AIRCRAFT
Method for detecting an operating anomaly of an aircraft component associated with at least one status indicator, said method—for at least one acquired time sequence (Seq1, Seq2) comprising acquired values of said at least one indicator—comprising steps of:
determining (E20) an approximated time sequence (Seq1*, Seq2*) of approximated values of said acquired values, by an approximation module (AE1, AE2);
determining (E30) a global anomaly score (A1, A2) for said acquired time sequence (Seq1, Seq2) from differences between said acquired values (Seq1, Seq2) and approximated values; and
detecting (E40) an operating anomaly of said component as a function of a comparison of said global anomaly score with a first threshold,
said approximation module being configured to minimize approximation errors between time sequences of reference values and time sequences of approximated values.
B64F 5/60 - Test ou inspection des composants ou des systèmes d'aéronefs
B64D 33/00 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs
34.
PROPULSION ASSEMBLY FOR AN AIRCRAFT PROVIDED WITH A PROPELLER AND A CYCLIC PITCH DEVICE FOR THE BLADES OF THE PROPELLER AND METHOD FOR REGULATING THE CYCLIC PITCH OF THE BLADES OF THE PROPELLER
A propulsion assembly for an aircraft extends along an axis and includes a propulsion module having a propeller provided with blades, an outlet guide vane, and a propeller shaft to rotate the propeller. The blades of the propeller are entirely or partially made of composite material. The propulsion assembly further includes a gas generator having a drive shaft; a speed-reduction device rotationally coupling the drive shaft and the propeller shaft to drive the propeller shaft at a rotational speed lower than the rotational speed of the drive shaft; and a cyclic pitch device for the blades of the propeller.
A turbomachine includes a first separation lip capable of separating an incoming air flow into a radially internal air flow and a secondary air flow; a second separation lip capable of separating the radially internal air flow into a primary flow and a tertiary flow traveling through a tertiary flow vein radially external to a primary flow vein traveled by the primary flow; a heat exchanger arranged in the tertiary flow vein; and an internal casing. The exchanger includes a body and a flange extending radially internally and projecting from the body, the flange being fixed to the internal casing, the exchanger including downstream of the flange, a downstream part to which is attached a fire wall forming a heat shield.
F02K 3/115 - Chauffage du flux dérivé à l'aide d'un échange indirect de chaleur
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
A manual deburring tool including a consumable carbide plate having deburring edges, and a casing which is configured to be grasped manually and has an internal cavity for housing the plate, the casing including a body and a cover removably attached to the body so as to allow the insertion of the plate into the cavity and the removal of said plate from the cavity, the housing including at least one opening that opens into the cavity and exposes a deburring edge.
B24D 15/02 - Outils à main ou autres dispositifs pour meuler, polir ou affûter, sans mouvement rotatif rigidesOutils à main ou autres dispositifs pour meuler, polir ou affûter, sans mouvement rotatif avec des surfaces de travail montées de façon rigide
37.
METHOD FOR PRODUCING AN ABRADABLE COATING, ABRADABLE COATING AND COATED PART
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE PAUL SABATIER TOULOUSE III (France)
Inventeur(s)
Rousselle, Mélanie Laura Josépha
Fradet, Guillaume
Estournes, Claude Gilbert Jean-Pierre
Ansart, Florence
Herisson De Beauvoir, Thomas
Abrégé
A method for producing an abradable ceramic composite coating on a substrate, the method including: obtaining a pulverulent composition including a matrix powder and a ceramic filler hydrated precursor powder having a lamellar crystallographic structure, wherein the ceramic filler powder represents from 5 to 40% of the combined volume of the matrix powder and the ceramic filler powder, compressing the prepared pulverulent composition at a pressure greater than 150 MPa, and a step of reactive sintering the obtained pulverulent composition, during which the compression is maintained, at a temperature of less than 550° C., and the particles of the matrix powder in the sintered pulverulent composition have an aspect ratio of 2 or greater.
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
C23C 24/08 - Revêtement à partir de poudres inorganiques en utilisant la chaleur ou une pression et la chaleur
F01D 25/00 - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
38.
SECTOR OF A TURBINE STATOR FOR A TURBINE OF AN AIRCRAFT TURBOMACHINE
A sector of a turbine stator for a turbine of an aircraft turbomachine, the sector having an outer platform and an inner platform which are coaxial along an axis. The sector further includes at least one blade which connects the outer and inner platforms together, with the blade being variable-pitch about an axis of rotation of the blade. The blade can include an aerodynamic body delimited radially by a head and a root, the head being placed with a first functional clearance in an opening of the outer platform, the root being placed with a second functional clearance in a cavity of the inner platform, the axis of rotation of the blade being located downstream of the head and the root.
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
39.
ASSEMBLY COMPRISING A RING AND A PIVOTING SUPPORT SLEEVE FOR VARIABLE PITCH VANE ROOTS, TURBINE ENGINE PROVIDED WITH SUCH AN ASSEMBLY AND METHOD FOR DISMANTLING SUCH AN ASSEMBLY
An assembly for a turbine engine includes a ring with a rotationally symmetrical axis. The assembly further includes a plurality of support sleeves, each having a rotationally symmetrical axis and configured to receive a stud of a variable pitch blade. Each support sleeve includes a bore passing right through the support sleeve along the rotationally symmetrical axis. Each support sleeve is mounted on the ring so as to pivot about a pivot axis transverse to the radial axis and to the rotationally symmetrical axis.
The disclosed fan blade is composed at least in part of a structure (50) made of composite material comprising a fiber reinforcement (62) and a matrix (64). The fiber reinforcement (62) includes a plurality of plies (66, 68) that are each formed of warp yarns (71) and weft yarns (72). The warp yarns (71) and weft yarns (72) include incomplete yarns (75, 76) that each have a terminal end (78) on the pressure side (40) or the suction side (42). The pressure side (40) and the suction side (42) each have at least one ply drop line connecting the terminal ends (78) of the incomplete yarns (75, 76) of the same ply (66, 68). The number of ply drop lines on the pressure side (40) is greater than the number of ply drop lines on the suction side (42) in a low region of the blade and/or in an extended region of the blade.
A device for guiding a shaft of an aircraft turbine engine, the device including a rolling bearing;—an annular bearing support and—an oil film compression damping system including two annular sealing segments; and—a first oil supply circuit this first circuit being connected to an annular supply groove wherein the damping system further—includes a second oil supply circuit, this second circuit being connected to at least one annular temperature control groove which is independent of the supply groove and which runs around one of the sealing segments.
A turbomachine assembly includes a high-pressure compressor driven by a high-pressure shaft, a low-pressure compressor driven by a low-pressure shaft, the low-pressure shaft being driven at a lower speed than the high-pressure shaft, a first power converter driven by the high-pressure shaft, an air circulation system including a first air bleed port positioned in the high-pressure compressor and a second air bleed port placed upstream of the first air bleed port, a device that determines an operating condition of the turbomachine assembly, and a controller that sends transfer instructions to the first power converter for transferring a power from the high-pressure shaft to the low-pressure shaft or a power from the low-pressure shaft to the high-pressure shaft depending on the operating condition.
F01D 21/00 - Arrêt des "machines" ou machines motrices, p. ex. dispositifs d'urgenceDispositifs de régulation, de commande ou de sécurité non prévus ailleurs
An assembly screw including at least a threaded part and a non-threaded part, the non-threaded part being equipped with a strain gage sensitive to the instantaneous elongation of the non-threaded part and with an RFID tag connected to the strain gage by a wired link, the assembly allowing the wireless transmission of an electrical resistance value of the strain gage representative of the instantaneous elongation.
F16B 31/02 - Assemblages à vis spécialement modifiés en vue de résister à une charge de tractionBoulons de rupture signalant ou limitant la charge de traction
44.
BLADE FOR A TURBOMACHINE TURBINE, AND CORRESPONDING ROTOR, TURBINE AND TURBOMACHINE
A blade intended to be mounted on a turbomachine rotor with an axis, the blade including a root configured to be mounted in a groove that opens to the outer periphery of a disk of the turbomachine rotor; an airfoil extending the root, in a radial direction with respect to the axis, and having an aerodynamic profile; a radially inner platform separating the airfoil from the root; a heel which extends as a continuation of the airfoil at a free radial end of the blade, at the end of the airfoil radially opposite the root, the heel including a radially outer platform, wherein the heel includes means for locking the radially outer platform with respect to another circumferentially adjacent radially outer platform.
A propeller for an aircraft turbine engine having a hub, vanes, bearings for guiding the roots of the vanes into orifices in the hub, and systems for retaining the roots of the vanes in the orifices in the hub along the pitch axis, this system for retaining the root including a ring configured to be fitted in a groove in the root, and a screw-nut system mounted on the root between the ring and one of the bearings.
A hydraulic control circuit of a steering actuator having a first and a second chamber, for orienting the blades of an aircraft engine propeller, this circuit comprising a high-pressure line and a low-pressure line, a first protection valve being able to assume a deactivated position or an activated position for placing the first chamber in communication with the high-pressure line, a second protection valve being able to assume a deactivated position or an activated position for placing the second chamber in communication with the low-pressure line, each protection valve comprising a hydraulic activation inlet and a hydraulic deactivation inlet which can be pressurised in order to activate or deactivate these valves. The circuit includes a solenoid protection valve controlling an amplification valve connected to the inlets of the protection valves in order to activate them upon the activation of the solenoid protection valve.
F15B 11/08 - Systèmes de servomoteurs dépourvus d'asservissement avec un seul servomoteur
B64C 11/38 - Mécanismes de changement de pas des pales par fluide, p. ex. hydrauliques
F02K 1/66 - Inversion du flux de la soufflante en inversant les aubes du ventilateur
F02K 1/76 - Commande ou régulation des inverseurs de poussée
F15B 20/00 - Dispositions propres à la sécurité pour systèmes de manœuvre utilisant les fluidesUtilisation des dispositifs de sécurité dans les systèmes de manœuvre utilisant des fluidesMesures d'urgence pour les systèmes de manœuvre utilisant des fluides
B64C 11/30 - Mécanismes de changement de pas des pales
47.
Aeronautical casing preform comprising a fibrous web with two deployable portions
A fibrous web extending between a proximal edge and a distal edge, between a first side edge and a second side edge, and between an inner face and an outer face, the fibrous web including a distal portion present between an intermediate part and the distal edge in the longitudinal direction, the distal portion including, in the direction of its thickness, an inner portion and an outer portion, wherein the outer portion is connected to the inner portion by at least one linking portion, the outer portion comprising a first deployable portion separated from the inner portion by a first non-interlinked area, and a second deployable portion separated from the inner portion by a second non-interlinked area.
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
B29K 105/08 - Présentation, forme ou état de la matière moulée contenant des agents de renforcement, charges ou inserts de grande longueur, p. ex. ficelles, mèches, mats, tissus ou fils
A fibrous texture having a three-dimensional weave between a plurality of layers of warp strands and a plurality of layers of weft strands. The fibrous texture includes first to fifth portions. The first and fifth portions extend in the lateral direction from the first and second lateral edges, respectively. The first and fifth portions each include warp strands composed of a first type of fiber corresponding to carbon fibers having a Young's modulus greater than 290 GPA and an elongation at break comprised between 1.2% and 2% and warp strands composed of a second type of fiber having a Young's modulus between 150 GPA and 250 GPA and an elongation at break between 4% and 6%.
A turbomachine includes an oil-recovery device having, between mutually rotatable first and second members, a tubular tab of the first member surrounded by a tubular extension of a shroud of the second member. The tab defines, with a shroud of the first member, an interface zone. The tubular extension includes an outer groove opposite the tab, and the extension axially surrounds the tubular tab beyond an axial half-length of the tab. The extension has an inner diameter equal to an inner diameter of the outer shroud of the second member and has an outer diameter smaller than or equal to an outer diameter of the outer shroud. The second member has an annular groove with a rectangular cross-section.
The invention relates to a pitch-change mechanism (70) which comprises a frame (72), a moving part (102) which can be translated along a longitudinal axis (X), and a locking device (160). The locking device (160) comprises a support member (162) which is translatable along the longitudinal axis (X) between an operating position and a locked position, and a screw-nut system (164). The screw-nut system (164) is formed by a system of satellite roller screws (195) comprising a screw (176) which is translatably secured to the support member (162) and rotatably mounted about the longitudinal axis (X), and a nut (178) which is secured to the moving part (102). The screw (176) has an abutment surface (186) which is separated from the frame (72) when the support member (162) is in the operating position and bears against the frame (72) when the support member (162) is in the locked position.
An oil tank for an aircraft turbomachine, including: a main enclosure capable of containing the oil, an oil inlet mixed with air, an oil outlet, an auxiliary enclosure arranged inside the main enclosure, including an enclosing wall, and two end walls adjacent to the enclosing wall, at least one of the two end walls including at least one oil passage with the interior of the main enclosure, at least one auxiliary inlet opening tangentially to the enclosing wall and passing through the main enclosure, an auxiliary outlet, tangential to a larger diameter section of the surrounding wall. The auxiliary outlet passes through the main enclosure for connection to the auxiliary circuit.
Guide vane (10) made of composite materials intended to be fixed to at least one inner fixed structure (20) of a turbine engine, said guide vane (10) comprising an aerodynamic blade (12) extending radially outwards along a radial axis (X-X), an outer radial end (14) and an inner radial end (16).
Guide vane (10) made of composite materials intended to be fixed to at least one inner fixed structure (20) of a turbine engine, said guide vane (10) comprising an aerodynamic blade (12) extending radially outwards along a radial axis (X-X), an outer radial end (14) and an inner radial end (16).
The vane comprises an inner platform (18) designed to be fixed to the inner fixed structure (20) and connected to the inner radial end (16) by an inner connecting elbow (16a) and extending from said inner radial end (16) along a single first tangential direction, secant to the radial axis (X-X) of the aerodynamic blade (12).
A gear changing system for the driving of an apparatus by an accessory gearbox including a train including a ring gear, a planetary gear integral with an output shaft and a planet carrier integral with an input shaft, a clutch device for the input shaft and output shaft, and a locking device. The system including a first configuration in which the clutch device is in the engaged position and the locking device is in the disconnected position, to drive the output shaft at the speed of the input shaft, and a second configuration in which the clutch device is in the disengaged position and the locking device is in the connected position, to drive the output shaft at a speed higher than the speed of the input shaft.
F16H 3/54 - Transmissions ayant uniquement deux engrenages centraux reliés par des engrenages à mouvement orbital avec engrenages orbitaux uniques ou paires d'engrenages orbitaux rigidement liés comprenant des engrenages orbitaux droits un des engrenages centraux étant denté intérieurement et l'autre denté extérieurement
54.
Pitch change mechanism with an actuator surrounding a fluid transfer bearing
The invention relates to a pitch change mechanism (70) comprising a frame (72), an actuating cylinder (74) having a moving part (102) capable of translating relative to the frame (72), and a linkage system (78) connecting the moving part (102) to a variable-pitch vane (56) so as to convert the translation of the moving part (102) into a rotation of the variable-pitch vane (56). It further comprises a fluid transfer bearing (257) for supplying the actuating cylinder (74) with actuating fluid from a source (76) relative to which the actuating cylinder (74) is capable of rotating. The actuating cylinder (74) extends at least partly around the fluid transfer bearing (257).
A bladed assembly (10) includes two adjacent blades (12). Each blade (12) a platform (16) located at a free radial end. The platform (16) of the first blade (12) includes a lateral face (18) which is located circumferentially opposite an associated lateral face (18) of the platform (16) of the second blade (12). The platform (16) of the second blade (12) comes into contact with the platform (16) of the second blade (12) only via an excrescence (20) which projects circumferentially with respect to the lateral face (18) of the platform (16) of the first blade (12) in the direction of the platform (16) of the second blade (12), and the excrescence (20) is received in a groove (22) formed in the platform (16) of the second blade.
The disclosed fan blade is composed at least in part of a structure (50) made of composite material comprising a fiber reinforcement and a matrix. The fiber reinforcement includes a plurality of plies that are each formed of warp yarns and weft yarns. The warp yarns and weft yarns include incomplete yarns that each have a terminal end on the pressure side or the suction side (42). The pressure side and the suction side (42) each have at least one ply drop line (82, 83) connecting the terminal ends of the incomplete yarns of the same ply. At least one ply drop line (82) includes at least one transverse portion (90) extending generally along a transverse direction (T), including an apex portion (92). The at least one apex portion (92) extends over at most 10% of the distance between the leading edge (44) and the trailing edge (46).
The disclosed fan blade is composed at least in part of a structure (50) made of composite material comprising a fiber reinforcement and a matrix. The fiber reinforcement includes a plurality of plies that are each formed of warp yarns and weft yarns. The warp yarns and weft yarns include incomplete yarns that each have a terminal end on the pressure side (40) or the suction side. The pressure side (40) and the suction side each have at least one ply drop line (80, 81) connecting the terminal ends of the incomplete yarns of the same ply. At least one ply drop line (81) has an end portion (96) that extends from the blade tip up to 90% of the blade height. Said end portion (96) extends generally along a direction of extent (E) forming an angle between 5° and 85° with the local tangent of the warp yarns.
A stator vane assembly defining a chordwise direction and a crosswise direction perpendicular to the chordwise direction is provided. The stator vane assembly includes: an attachment plate; and an attachment assembly coupled to the attachment plate, the attachment assembly including: a first flange member attached to the attachment plate and extending along the chordwise direction; a second flange member attached to the attachment plate and extending along the chordwise direction, the first flange member spaced from the second flange member in the crosswise direction; and a crosswise support member extending between the first and second flange members and positioned between the attachment plate and the first flange member and between the attachment plate and the second flange member.
A jumper wire for connecting an electrical cable to a test device with a view to diagnosing the state of health of the electrical cable. The jumper wire includes an output port configured to interact with an input port of the test device, an input connector configured to interact with a connector of the electrical cable, and a data-storing element that contains predetermined test data on the electrical cable and that is electrically connected to the output port by at least one data-transmitting wire. The data-storing element being is configured to deliver the test data to the test device.
G01R 31/58 - Test de lignes, de câbles ou de conducteurs
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
A bladed assembly for a turbomachine includes: a plurality of blades distributed about an axis and each including an airfoil and a platform formed at a free end of the blade. The platform of each blade includes a friction member and an opening that is oblong in a circumferential direction with respect to the axis. The friction member is engaged through the opening formed in the platform of a circumferentially adjacent blade within the bladed assembly with a clearance allowing the friction member to move at least in the circumferential direction within the opening. Mutual separation of the platforms in the circumferential direction during operation due to vibrations of the blades causes the friction member to rub against an edge of the opening and thereby dissipate energy.
A method for controlling a turbomachine comprising including a fan positioned upstream of a gas generator. The turbomachine including an electric motor forming a torque injection device for a high-pressure rotary shaft, in which method a fuel flow setpoint for a combustion chamber and a torque setpoint (Tcons) for the electric motor (ME) are determined. The control method including the steps of determining (El) a hybridisation rate (TH) corresponding to the ratio of the power (Pe) consumed by the electric motor (ME) to the power generated by the high-pressure rotary shaft (P2), determining (E2) a torque threshold (Tseuil) from the hybridisation rate (TH), limiting (E3) the torque setpoint (Tcons) to the torque threshold (Tseuil) if the torque setpoint (Tcons) is higher than the torque threshold (Tseuil).
F02C 9/48 - Commande de l'alimentation en combustible combinée avec une autre commande de l'ensemble fonctionnel
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
62.
Bladed turbomachine assembly including means for limiting vibrations between platforms
A bladed turbomachine assembly including at least two adjacent blades, each blade including a radial aerofoil and a platform located at a free radial end of the aerofoil, the platform of each blade of the bladed assembly being located circumferentially facing a platform of the other circumferentially adjacent blade of the bladed assembly, the bladed assembly including an insert arranged circumferentially between the platforms and which cooperates with the platforms, each platform including a circumferential bearing face against which a circumferential end of the insert is in contact, and the direction of the contact between each circumferential bearing face and the insert being oriented circumferentially with respect to the main axis (A).
The present invention relates to a flow-straightening stator (5) of an unducted thrust-generating assembly (3) comprising a plurality of blades (7) each having: —a tip (11) and a radially inner boundary (10) corresponding to an intersection between the blade (7) and a casing (8) of the turbine engine (1); —a skeleton (16); and—a deviation (S) of the profile, corresponding to an absolute value of a difference between a tangent to the skeleton (16) at the leading edge (14) and a tangent to the skeleton (16) at the trailing edge (15) of the blade (7), of between 20° and 450 at the radially inner boundary (10) of the blade (7) and between 10° and 40° at the tip (11) of the blade (7).
A heat exchanger for an electric machine. The electric machine includes a stator having a stator core and stator slots extending along a longitudinal axis of the stator, and a rotor assembly having a rotor configured to rotate about the longitudinal axis. The electric machine includes a coolant system having a stator cooling path providing a first coolant fluid in direct contact with the stator to flow around the stator core and through the stator slots. The coolant system also includes a coolant circuit providing a second coolant fluid to at least one rotatable component of the electric machine, such as at least one bearing. The coolant circuit and the stator cooling path are fluidically isolated from one another and arranged such that heat is exchanged between the first coolant fluid and the second coolant fluid at the stator.
H02K 9/197 - Dispositions de refroidissement ou de ventilation pour machines avec enveloppe fermée et circuit fermé de refroidissement utilisant un agent de refroidissement liquide, p. ex. de l'huile dans lesquels l'espace du rotor ou du stator est étanche au fluide, p. ex. pour pourvoir le rotor et le stator d'agents de refroidissement différents
B64D 27/30 - Aéronefs caractérisés par des groupes moteurs électriques
DEVICE FOR AXIALLY RETAINING MOVING BLADES OF AN LP TURBINE IN THE SLOTS OF A ROTOR DISC OF THE LP TURBINE AND METHOD FOR ASSEMBLING SAID MOVING BLADES
A device for axially retaining a moving blade of a low-pressure turbine in a recess of a rotor disc of axis of the turbine. This device includes a stop part intended to extend between an upstream face of a blade root of the moving blade and a downstream face of the root of the moving blade, the stop part including a first wall which forms an end stop for the upstream face of the blade root, a second wall which forms an end stop for the downstream face of the blade root, and a connection surface intended to extend between the first and the second wall along a radially inner end face of the blade root.
The invention relates to a blade comprising, a blade body having an aerodynamic profile comprising a first fibrous reinforcement obtained by three-dimensional weaving and comprising a first matrix in which the first fibrous reinforcement is embedded, the blade body part comprising a cavity formed by a disconnection of the first fibrous reinforcement, a blade root intended to be connected to a mechanism for the variable adjustment of the blade, and a spar comprising a core made of composite material and two metal casings attached to the core on either side of the core. The core comprises a first part that extends inside the cavity of the blade body and a second part that forms the blade root. The two metal casings which are attached to the core made of composite material extend over the second part and continue to extend over the first part inside the cavity of the blade body.
An aeronautical propulsion unit includes a central engine, an upstream series and a downstream series of blades, the blades of at least one of these series being adapted to be driven in rotation about the central longitudinal axis by a turbine, and a nacelle which encloses the central engine. At least some of the blades of the two series are variable-pitch blades. At least one of the blade of the first series of blades has a trailing edge having serrations, and/or at least one of the blades of the second series of blades has a leading edge having serrations.
A bearing for a turbine engine includes an oil film compression damper, including an inner sleeve intended to receive an outer ring of a roller bearing, an outer sleeve surrounding the inner sleeve and delimiting with this inner sleeve, an axisymmetric cavity supplied hydraulically in order to form a film of damping oil, this cavity being closed at its ends by two segments. The cavity includes a first portion and a second portion separated from one another by a movable segment able to occupy a closed position in which it forms a controlled sealing barrier between the two portions, and an open position in which the two portions communicate with one another, and in that it is the first portion that is supplied hydraulically to form the film of oil.
F01D 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
F16C 19/52 - Paliers à contact de roulement pour mouvement de rotation exclusivement avec dispositifs affectés par des conditions anormales ou indésirables
F16C 27/04 - Roulements à billes ou à rouleaux, p. ex. à organes de roulement élastiques
69.
AIRCRAFT TURBOMACHINE HAVING VARIABLE-PITCH PROPELLER VANES
An assembly including a propeller blade and a system for angularly setting the pitch of the blade for an aircraft turbomachine, the blade having a root extending from an upper end connected to an airfoil of the blade to a free lower end, the root having a bulging segment, which bulging segment is referred to as the “bulb”, the system for angularly setting the pitch of the blade including a cup which is radially defined by an annular wall and which includes a lower bottom closed by a bottom wall and an upper opening through which the bulb is intended to be axially inserted into the cup, the bottom wall including a protrusion which engages with a cavity having a complementary shape in the free end of the root.
A turbomachine includes a heat exchanger, a first circuit of a first fluid, the first circuit including two ducts in the heat exchanger, a second circuit of a second fluid, and a third circuit of a third fluid. The second circuit includes a first channel in the heat exchanger and the third circuit includes a second channel in the heat exchanger, each of the first channel and the second channel are contiguous to one of the two ducts so as to be in thermal contact with one of the two ducts, the first channel and the second channel are contiguous to each other so as to be in thermal contact with each other, and the heat exchanger includes a section in which the two ducts, the first channel and the second channel are locally parallel to each other.
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
F01D 15/10 - Adaptations pour la commande des générateurs électriques ou combinaisons avec ceux-ci
F28D 7/00 - Appareils échangeurs de chaleur comportant des ensembles de canalisations tubulaires fixes pour les deux sources de potentiel calorifique, ces sources étant en contact chacune avec un côté de la paroi d'une canalisation
F28D 7/10 - Appareils échangeurs de chaleur comportant des ensembles de canalisations tubulaires fixes pour les deux sources de potentiel calorifique, ces sources étant en contact chacune avec un côté de la paroi d'une canalisation les canalisations étant disposées l'une dans l'autre, p. ex. concentriquement
71.
ACCESSORY GEARBOX AND AIRCRAFT TURBINE ENGINE COMPRISING SUCH A BOX
Accessory gearbox for an aircraft turbine engine, including: a casing, a reduction gear, and a freewheel which are permanently integrated into the casing, the accessory gearbox being configured to interface with a starter of the turbomachine that is intended to start up the turbomachine, the reduction gear and the freewheel being dedicated to the starter of the turbomachine.
An assembly for a turbomachine of longitudinal axis includes a nozzle carried by an annular part which is fixed to an outer casing. The nozzle has an outer annular platform from which extends, radially outwards, a radial annular wall which carries an annular attachment tab extending longitudinally upstream from the radial annular wall. The assembly has a cooling circuit with an inner annular chamber delimited radially by the outer annular platform and the annular part and delimited downstream by the radial annular wall. An intermediate annular chamber is formed radially externally to the annular attachment tab and in fluid communication with the outer casing. Means of fluid communication between the inner annular chamber and the intermediate annular chamber is formed in the annular attachment tab of the nozzle.
A method for monitoring a rotating machine in order to detect a fault in a bearing, the method including acquiring, from the rotating machine, a vibration signal measured by a vibration sensor; determining a first-order spectrogram by first-order cyclostationary analysis of the vibration signal using a delta transform and spectral standardisation; determining a second-order spectrogram by second-order cyclostationary analysis of the vibration signal using averaged cyclic coherence, a delta transform and spectral standardisation; and detecting a vibration signature of the fault in the bearing on the basis of the first-order spectrogram and the second-order spectrogram.
G06G 7/70 - Calculateurs analogiques pour des procédés, des systèmes ou des dispositifs spécifiques, p. ex. simulateurs de véhicules, p. ex. pour la détermination du chargement admissible de navires
G01M 13/045 - Analyse acoustique ou des vibrations
G01M 15/12 - Test des moteurs à combustion interne par contrôle des vibrations
74.
Propeller blade or vane having a hollow composite root
A propeller blade or airfoil for a turboprop engine made from composite material including a matrix-densified fibrous reinforcement, the propeller blade or airfoil including, in the direction of its span, a root and an aerodynamic profile. The fibrous reinforcement includes a fibrous preform having three-dimensional weaving with a root preform portion and an aerodynamic profile preform portion. The fibrous preform includes a separation delimiting a recess that forms a cavity extending both into the root and into the aerodynamic profile. A spar is present in the cavity, the spar including an aerodynamic profile shaping portion positioned in a first portion of the cavity and a root shaping portion positioned in a second portion of the cavity. The root has a rotationally symmetric shape.
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
a second lateral wall (52.1, 52.2) including a projecting portion (57.1, 57.2) intended to engage with a mounting groove (46.1, 46.2) formed in a vertical face of the compressor drum (20).
A module having a longitudinal axis, a hydraulic actuator, and a pump for supplying the hydraulic actuator with fluid, the pump can include axial pistons intended to be movable to rotate about the longitudinal axis and configured to transfer the fluid to the hydraulic actuator, a connecting plate connected to the axial pistons and engaging with an annular main plate centered on the longitudinal axis intended to be rotated about the longitudinal axis and tilted relative to the longitudinal axis, the connecting plate being arranged between the main plate and the axial pistons, the axial pistons being movable to rotate in a direction parallel to the longitudinal axis, the tilting of the main plate resulting in the movement of the axial pistons in the direction.
A sound absorber extending between two opposite surfaces, and combining one or more quarter-wave acoustic resonators, with a microporous element. Each acoustic resonator, has a length L, which is substantially greater than a thickness t of the sound absorber, between a first open end on a first surface of the two opposite surfaces, of the sound absorber and a second closed end. The microporous element is made up of a plurality of periodically repeating unit cells adjacent to the acoustic resonators. A method for producing this sound absorber, which method includes at least one additive manufacturing step.
F02C 7/045 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction comportant des dispositifs destinés à supprimer le bruit
B33Y 80/00 - Produits obtenus par fabrication additive
F02K 1/82 - Parois des tubulures de jet, p. ex. chemises
G10K 11/172 - Procédés ou dispositifs de protection contre le bruit ou les autres ondes acoustiques ou pour amortir ceux-ci, en général utilisant des effets de résonance
78.
METHOD FOR PRODUCING A PART MADE OF A Y/Y' NICKEL-BASED ALLOY BY HOT FORGING
The invention relates to a method for producing a part made of a nickel-based alloy with a Y/Y′ microstructure in which at least one hot-forging step is performed, which method is characterised in that the temperature at which the part is heated in the forging step is maintained within a temperature range lower than the abnormal grain growth temperature range of the alloy, the hot-forging temperature being maintained at a temperature lower than the temperature of the part, the hot-forging temperature being such that the difference compared with the temperature at which the part is heated is less than 250° C. and preferably less than 150° C.
C22C 19/05 - Alliages à base de nickel ou de cobalt, seuls ou ensemble à base de nickel avec du chrome
B21K 1/32 - Fabrication d'éléments de machines de rouesFabrication d'éléments de machines de disques de disques, p. ex. de roues pleines
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
79.
Aeronautical propulsion system comprising an optimized fan section
A fan section of an aeronautical propulsion system includes a fan rotor including seventeen blades to twenty blades and having a solidity strictly less than 1.0. The solidity is equal to a ratio between a chord at the blade tip and an inter-blade pitch at the blade tip. The fan section has a hub-tip ratio greater than or equal to 0.22 and less than or equal to 0.32, a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5, and a peripheral speed at the blade tip greater than or equal to 260 m/s and less than or equal to 400 m/s. The pressure ratio and the peripheral speed are measured at cruising speed.
F01D 5/00 - AubesOrganes de support des aubesDispositifs de chauffage, de protection contre l'échauffement, de refroidissement, ou dispositifs contre les vibrations, portés par les aubes ou les organes de support
F01D 1/04 - "Machines" ou machines motrices à déplacement non positif, p. ex. turbines à vapeur avec des moyens stationnaires de guidage de fluide de travail et un rotor à ailettes ou de structure analogue traversées par le fluide de travail principalement dans le sens axial
F01D 25/00 - Parties constitutives, détails ou accessoires non couverts dans les autres groupes ou d'un intérêt non traité dans ces groupes
A multi-flow turbine engine for an aircraft, the turbine engine having a gas generator with geometrical characteristics allowing it to have at its periphery a cold compartment larger than its hot compartment.
F04D 29/52 - Carters d'enveloppeTubulures pour le fluide énergétique pour pompes axiales
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
81.
OUTLET NOZZLE PROVIDED WITH CHEVRONS FOR AERONAUTICAL THRUSTER
Exhaust nozzle for an aeronautical propulsion system disposed downstream of at least one stator blade of the propulsion system, a trailing edge of the exhaust nozzle including a plurality of chevrons distributed circumferentially around a central axis, said trailing edge including a first trailing edge pattern extending over an angular range Δθ including the at least one stator blade, the first pattern being a portion of the trailing edge including at least one chevron or being a portion of the trailing edge not having a chevron, and a second trailing edge pattern extending over an angular range Δθ′ not including the at least one stator blade, the second pattern being a portion of the trailing edge including at least one chevron or being a portion of the trailing edge not having a chevron, and being distinct from the first pattern.
A method for controlling a gas turbine engine including estimating a temperature gradient between first and second portions of a spool, comparing the estimated temperature gradient with a temperature gradient threshold, if the temperature gradient is less than the threshold, driving a drive shaft in rotation at a rotational speed varying over time according to a first profile of rotational speed variation, and if the temperature gradient is greater than the threshold, controlling the electric motor to drive the drive shaft in rotation such as to vary the rotational speed of the drive-shaft according to a second profile of rotational speed variation over time so that, when the rotational speed lies within a critical interval of rotational speeds, acceleration of rotation of the drive shaft is greater than acceleration of rotation of the drive-shaft according to the first profile of rotational speed variation, within the same critical interval.
F01D 19/02 - Démarrage des "machines" ou machines motricesDispositifs de régulation, de commande ou de sécurité en rapport avec les organes de démarrage dépendant de la température des éléments constitutifs, p. ex. du carter de la turbine
F01D 25/36 - Vireur ou dispositif de rotation par mouvements saccadés utilisant des moteurs électriques
A method for controlling a motor assembly. This motor assembly includes at least a first electric machine and a gas turbine engine. The gas turbine engine includes a low-pressure shaft and a high-pressure shaft. The electric machine is coupled to the low-pressure shaft, and the control method includes a step in which a take-off of mechanical work is ordered from the first electric machine to brake a rotation of the low-pressure shaft in response to an activation of a thrust reverser of the gas turbine engine and/or a disturbance of the air flow in a transverse plane at an air intake of the gas turbine engine. A control unit suitable for carrying out this method, a motor assembly incorporating this control unit, the electric machine and the gas turbine engine, and a computer program to carry out this method.
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 27/16 - Aéronefs caractérisés par le type ou la position des groupes moteurs du type à réaction
A gas turbine includes two unducted propellers, a main channel, a compression stage located in the main channel, the main channel supplying the compression stage with air, a cooling channel extending from a cooling inlet opening into the main channel, the cooling inlet being located upstream of the compression stage with reference to a direction of flow of the air through the gas turbine, a heat exchanger located in the cooling channel, an ejection channel opening into the main channel and into the cooling channel so that air passing through the ejection channel ventilates the heat exchanger, and a variable bleed valve that adjust an air flow rate in the ejection channel.
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
B64C 11/48 - Ensembles de plusieurs hélices coaxiales
B64D 33/08 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des systèmes de refroidissement des ensembles fonctionnels de propulsion
F02C 6/08 - Ensembles fonctionnels de turbines à gaz délivrant un fluide de travail chauffé ou pressurisé à d'autres appareils, p. ex. sans sortie de puissance mécanique délivrant des gaz comprimés le gaz étant prélevés sur le compresseur de la turbine à gaz
F02C 7/04 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction
F02C 7/042 - Entrées d'air pour ensembles fonctionnels de turbines à gaz ou de propulsion par réaction à géométrie variable
F02C 7/18 - Refroidissement des ensembles fonctionnels caractérisé par l'agent refroidisseur l'agent refroidisseur étant gazeux, p. ex. l'air
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
B64D 33/02 - Aménagement sur les aéronefs des éléments ou des auxiliaires des ensembles fonctionnels de propulsion, non prévu ailleurs des entrées d'air de combustion
85.
Suspension of a triple-flow aircraft turbine engine
A triple-flow turbine engine for an aircraft, this turbine engine having a longitudinal axis and including upstream suspension elements which are located in a first plane perpendicular to the axis and are connected or fixed to the gas generator of the turbine engine; downstream suspension elements which are located in a second plane perpendicular to the axis and are connected or fixed to the gas generator; and thrust-absorbing rods which include first ends which are connected or fixed to the gas generator and opposing second ends which are located in a third plane perpendicular to the axis, wherein the first, second and third planes are located at a cold compartment of the gas generator.
F02C 7/20 - Montage ou bâti de l'ensemble fonctionnelDisposition permettant la dilatation calorifique ou le déplacement
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
A suspension assembly for a turbine engine includes a beam that attaches to an aircraft pylon. The beam has first and second portions, each carrying a half-housing together defining a ball joint housing. A ball joint nut is hingedly engaged to rotate in the ball joint housing, and a cylindrical part is configured to be attached to a fixed portion of the turbine engine, hinged in the ball joint housing, and mounted so as to be pivotable about its axis in the ball joint nut. The suspension assembly further includes a support ring surrounding the two half-housings around their entire circumference. The support ring performs a safety function in the event of failure of at least one element of the suspension assembly.
A tool is configured for molding a duct for an aircraft turbine engine, wherein the duct is made of a fiber-based composite material and includes a curved tubular portion, one end of which is connected to a peripheral flange. The tool has a base and a generally elongate and curved body. The body is formed by an assembly of first parts which are fitted tightly together and which include a first central release key which extends from one end of the body to the other and which is configured to be removed first upon release of the body.
A method for maintaining a bladed wheel of a high-pressure turbine of a turbomachine having a disc with cavities, and blades extending radially from a radially internal root mounted in a cavity of the disc, the root being supported on the disc by means of surfaces of the root and of the disc forming bearing surfaces. At least one foil is removably mounted between the root of at least one blade and the disc, at the corresponding bearing surfaces. The method includes measuring a radial clearance between the tip of at least one blade, meaning the radially external end of the blade, and a shroud of the turbomachine situated facing the bladed wheel, and if the radial clearance is greater than a determined value, removing the foil.
The method for manufacturing a blade (32) of a turbomachine (100) comprises: —manufacturing a part (4) comprising an air flow path zone (10) and a layer (20) covering the zone, the layer having cavities (21) forming a periodic pattern, the manufacture taking place by injecting a mixture comprising a binder and a power; —removing a larger part of the binder from the part; —sintering the part, and —removing the layer (20) from the part to obtain the blade (32).
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 5/04 - Fabrication de pièces ou d'objets à partir de poudres métalliques caractérisée par la forme particulière du produit à réaliser d'aubes de turbines
B23C 3/18 - Usinage de surfaces à double courbure pour le façonnage d'hélices, d'aubes de turbines ou de roues à ailettes
B23P 15/02 - 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 une seule pièce
An aeronautical thruster of longitudinal axis has a hub, an annular row of unducted upstream rotor blades and an annular row of unducted downstream stator blades. Each downstream stator blade is of variable pitch, and at least one of the downstream stator blades is in a closed-pitch configuration relative to another of the downstream stator blades in that it has a pitch angle smaller than the pitch angle of the other downstream stator blade.
A method for locating a mark on a casing of a turbomachine of an aircraft, wherein the method includes a) a step of visually inspecting the arms, the outer surface of the internal shroud and the inner surface of the external shroud, and, in the event that at least one mark is detected at step a), b) a step of manually mounting a visual inspection device around the inner end or outer end of one or more of the arms, and c) a step of determining the location of the at least one mark in a zone of the casing.
G01B 5/14 - Dispositions pour la mesure caractérisées par l'utilisation de techniques mécaniques pour mesurer une distance ou une marge entre des objets ou des ouvertures espacés
F01D 25/24 - Carcasses d'enveloppeÉléments de la carcasse, p. ex. diaphragmes, fixations
92.
METHODS AND APPARATUS FOR MOUNTING A GAS TURBINE ENGINE
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 blading (25, 26) for a turbomachine (10), comprising: —a blade (31) having an aerodynamic profile; —a platform (32, 33) comprising a flow-path surface (321) intended to delimit a primary flow path (21A) of the turbomachine (10), which path is intended, when the turbomachine (10) is in operation, to receive a flow that splits, upstream of the blade (31), into a suction-face flow (EE) and a pressure-face flow (EI); and —an internal canal (34) which has an intake opening (35) and an ejection opening (36), these each opening onto the flow-path surface (321) of the platform (32, 33), the ejection opening (36) opening downstream of the intake opening (35) and the intake opening (35) opening toward the pressure-face flow (EI).
An aeronautical thruster of longitudinal axis includes a hub, an annular row of unducted upstream rotor blades, and an annular row of unducted downstream stator blades. The annular row of downstream stator blades includes at least one downstream stator blade of a first type, each of which is located about the longitudinal axis in a first angular sector about the longitudinal axis. Each downstream stator blade of the first type has a fixed pitch. The annular row of downstream blades further includes at least one downstream stator blade of a second type, each of which is located about the longitudinal axis outside said first angular sector, wherein each downstream stator blade of the second type has a variable pitch.
A method for controlling a motor assembly. The motor assembly includes a gas turbine engine and at least one electric machine which is mechanically coupled to a rotating shaft of the gas turbine engine so as to be rotated for generating electricity. In this control method, a mechanical power take-off setpoint by the electric machine is changed when an operating parameter of the gas turbine engine reaches a predetermined limit. A control unit adapted to perform this method, a motor assembly incorporating this control unit, the electric machine and the gas turbine engine, and a computer program for performing this method.
F02C 6/20 - Aménagements des ensembles fonctionnels de turbines à gaz pour l'entraînement des véhicules
F02C 9/00 - Commande des ensembles fonctionnels de turbines à gazCommande de l'alimentation en combustible dans les ensembles fonctionnels de propulsion par réaction alimentés en air ambiant
96.
FAN MODULE IMPELLER, AND JET TURBINE ENGINE EQUIPPED WITH SUCH AN IMPELLER
Turbojet engine fan module wheel, including a plurality of blades made of composite material, each blade having a root assembled with a base distinct from the bases of the other blades, each base having a groove extending axially and opening out on the side of the upstream face and on the side of the downstream face, each root cooperating by axial interlocking in a form-fitting manner, for example in the shape of a dovetail, with the groove of the base, whereby the root is retained on the base along the radial and circumferential directions, and each base cooperates with at least one part configured to axially block the root within the groove of the base, whereby the root is retained along the axial direction.
F02K 7/00 - Ensembles fonctionnels dans lesquels le fluide de travail est utilisé uniquement sous forme de jet, c.-à-d. ensembles ne comportant ni turbine ni autre moteur entraînant un compresseur ou une soufflante carénéeLeur commande
B64C 11/06 - Montage des pales dans le cas de pales à pas variable
F01D 7/00 - Rotors à aubes réglables en marcheLeur commande
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
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
97.
ALLOY POWDER, METHOD FOR MANUFACTURING A PART BASED ON SAID ALLOY AND RESULTING PART
The invention relates to a nickel-based alloy powder, that comprises in weight percentages, 14.00 to 15.25% of chromium, 14.25 to 15.75% of cobalt, 4.00 to 4.60% of aluminium, 0 to 0.50% of iron, 0 to 0.15% of manganese, 3.00 to 3.70% of titanium, 3.90 to 4.50% of molybdenum, 0 to 0.015% of sulphur, 0 to 0.06% of zirconium, 0.012 to 0.020% of boron, 0 to 0.20% of silicon, 0 to 0.10% of copper, 0 to 150 ppm of carbon, 0 to 0.5 ppm of bismuth, 0 to 5 ppm of lead, 0 to 1000 ppm of platinum, 0 to 1000 ppm of palladium, 0 to 50 ppm of hydrogen, 0 to 5 ppm of silver, 0 to 120 ppm of nitrogen, 0 to 1000 ppm of rhenium, 0 to 410 ppm of oxygen and 0 to 500 ppm of inevitable impurities, the rest being made up of nickel, and has a particle size D10 between 3 and 10 μm, a particle size D90 between 20 and 40 μm and a particle size D50 between 10 and 20 μm, the values of the particle sizes D10, D50 and D90 having been measured by laser diffraction according to standard ISO 13322-2. The invention also relates to a method for manufacturing a part using said powder and a resulting part.
B22F 1/10 - Poudres métalliques contenant des agents lubrifiants ou liantsPoudres métalliques contenant des matières organiques
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
A system including: a blade including, at the level of the blade root, two portions spaced apart from each other in the transverse direction X so as to provide therebetween a cavity extending downwards in the direction Y of the height of the blade, from a cavity bottom to a cavity opening on the outside, located at the level of the lower end of the blade root, an attachment element of the blade partially engaged inside the cavity, the engaged part having a shape, taken in a plane defined by the directions X and Y, which extends towards the bottom while flaring out in the direction X, the two spaced apart portions of the blade root which are in contact with the flared shape of the attachment element having corresponding flared shapes.
A method for carrying out tomographic analysis of a composite part comprising a matrix and fibers, the method including: acquiring at least one two-dimensional image of the part by a tomographic device, generating a plurality of shapes characteristic of the fibers and of convolution masks having a plurality of copies of the characteristic shape, computing, for each of the characteristic shapes, a product of convolution of the two-dimensional image with the corresponding convolution mask, and obtaining a convolution image, detecting a position corresponding to an overall maximum in all the obtained convolution images and attributing a fiber center to the position, marking as processed a region of the image placed around the fiber center, and iterating the steps of computing, detecting and marking on the unprocessed regions of the image.
A sealing system for a turbomachine bearing enclosure includes bearing with a stator carrying a rotor; a seal carried by the stator and surrounding the rotor; an external peripheral groove carried by the stator in the vicinity of the seal for collecting oil which exits the enclosure running along the rotor through the seal and which is centrifuged by the rotor towards this groove; and a drain at the bottom for discharging the oil from the groove. This groove is open towards the inside and is delimited by a base extended by two side walls, the side wall opposite the seal is extended by a skirt oriented obliquely towards the base to delimit a gutter opening towards the base.
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 25/16 - Aménagement des paliersSupport ou montage des paliers dans les stators
brevets
