A control ring for a stage of variable-pitch vanes for a turbine engine includes at least one member for bearing on a casing and means for fixing the member. The member includes a bushing with an axial bore for the passage of the member or an element for supporting the member. A through-slot opens into the bore and allows for substantial radial deformation of the bushing. The member further includes an outer thread for screwing the bushing into a complementary thread of a hole in the body, thereby deforming the bushing. The member or support can be mounted and moved inside the bore, to a second position, in which the bushing is radially constrained and is tightly mounted on the member or support, which is thus immobilized in relation to the bushing.
Tooling for balancing a turbine engine module (10) in a balancing machine, the turbine engine module having at least one stator housing (14) and a rotor (16) having a shaft (18) with a longitudinal axis A and at least one blade stage (20) surrounded by the stator housing (14). The tooling has at least a balancing frame (14), having rotor (16) guide bearings, first and second annular plates (30, 32) designed to be attached to the stator housing (14), third and fourth attachment lugs (34, 36) provided on the balancing frame (24), to attach the first and second annular plates (30, 32) to the frame. A trolley is for transporting the frame (24), and a support (84, 94) for supporting the frame, provided on the balancing frame (24) and cooperating equally well with the balancing machine and with the trolley.
G01M 1/28 - Determining imbalance by oscillating or rotating the body to be tested with special adaptations for determining imbalance of the body in situ, e.g. of vehicle wheels
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
D03D 25/00 - Woven fabrics not otherwise provided for
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
4.
Method for knocking out a foundry core and method for manufacturing by casting comprising such a method
A method for knocking out a foundry core confined in an internal cavity in a part at the end of a casting operation, in particular a lost-wax casting operation, includes at least a primary chemical knocking-out step. During the primary chemical knowing-out step, the part is subjected to a chemical solution to dissolve the core, in a sealed enclosure. The method further includes a secondary step of knocking out by ultrasounds in water or an aqueous solution contained in an ultrasound tank, during which the part is subjected to ultrasounds to loosen core residues from walls of the cavity.
b), at least one acoustic panel (26) fastened using fastening elements (48, 54) to the inner surface of the fan casing, and at least one circumferential stiffener (40) of the fan casing (14). According to the invention, the fastening elements (48, 54) connect the fan casing (14) to the stiffener (40) incorporated with the acoustic panel (26).
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
A blade including at least one web and a vane having a leading edge and a trailing edge, wherein, for at least one aerofoil of the vane in the vicinity of the web, a maximum sweep angle associated with a position along a chord of the aerofoil extending from the leading edge to the trailing edge of the vane corresponding to a relative chord length of at least 50%.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
8.
Cooling device for a turbomachine supplied by a discharge circuit
F02C 7/14 - Cooling of plants of fluids in the plant
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
9.
Lubricating-oil collection cap for turbomachine equipment
An annular cap for collecting lubricating oil for turbomachine equipment is configured to extend around the equipment and to rotate about an axis. The cap includes through-orifices through which the oil can pass radially under the effect of spinning. The cap further includes means of deflecting the oil leaving the orifices in a direction substantially transverse to the axis and substantially tangential to the cap.
b) including a receptacle (20) suitable for retaining a blade root, and at least one arm (17) of organic matrix composite material extending laterally relative to said longitudinal axis (Z) and supporting a flyweight (16).
A warp yarn take-up system includes a clamping device for holding a plurality of layers of warp yarns, the clamping device being movable at least in a direction corresponding to the advance direction of the warp yarns. The clamping device includes a bottom clamp, a top clamp, and at least one intermediate clamping element present between the bottom clamp and the top clamp. The bottom clamp, the top clamp, and the at least one intermediate clamping element are held together by clamping.
D03D 13/00 - Woven fabrics characterised by the special disposition of the warp or weft threads, e.g. with curved weft threads, with discontinuous warp threads, with diagonal warp or weft
12.
Method for producing a turbine engine part, and resulting mould and intermediate blank
d), in which the third and fourth sides extend between the first and second sides, flaring apart from the first side towards the second side, first at a first angle and subsequently at a second larger angle, and said at least one part is machined in the blank.
B22D 25/02 - Special casting characterised by the nature of the product by its peculiarity of shapeSpecial casting characterised by the nature of the product of works of art
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
B22D 13/06 - Centrifugal castingCasting by using centrifugal force of solid or hollow bodies in moulds rotating around an axis arranged outside of the mould
B22D 27/15 - Treating the metal in the mould while it is molten or ductile by using vacuum
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
13.
Method for detecting a failure in a motive flow valve of an aircraft engine fuel circuit
A method detects a failure in a fuel return valve of an aircraft engine fuel circuit. A fuel system is connected to a fuel tank of the circuit and includes a high-pressure pump delivering a flow rate to an actuating cylinder, a cutoff valve capable of feeding the actuating cylinder disposed in a feed pipe of the engine; a fuel return pipe; a fuel return valve arranged to switch between an open position and a closed position. The method includes starting the engine at an engine speed; increasing the engine speed until a flow rate reaches a predefined value sufficient for opening the cutoff valve; measuring the position of the actuating cylinder and an engine speed corresponding to the opening of said cutoff valve.
a control center (600) suitable for automatically controlling one or more of the mechanisms and devices and/or the conveyor of the replacement device (1).
B23P 19/04 - Machines for simply fitting together or separating metal parts or objects, or metal and non-metal parts, whether or not involving some deformationTools or devices therefor so far as not provided for in other classes for assembling or disassembling parts
B23C 9/00 - Details or accessories so far as specially adapted to milling machines or cutters
B23P 21/00 - Machines for assembling a multiplicity of different parts to compose units, with or without preceding or subsequent working of such parts, e.g. with programme control
B23C 5/20 - Milling-cutters characterised by physical features other than shape with removable cutter-bits or teeth
B23P 19/00 - Machines for simply fitting together or separating metal parts or objects, or metal and non-metal parts, whether or not involving some deformationTools or devices therefor so far as not provided for in other classes
B23P 19/06 - Screw or nut setting or loosening machines
15.
Mobile turbine blade with an improved design for an aircraft turbomachine
A turbine blade for an aircraft turbomachine including a root, an airfoil and a platform inserted between the airfoil and the root and delimiting a gas circulation flowpath, the platform having two axial ends each forming an angel wing of which at least one has an internal cavity that will be supplied with air from the root of the blade. At least one of the two angel wings is drilled with at least one bleed hole for passage of a bleed flow from the internal cavity that will limit/prevent gas recirculation outside the flowpath.
F01D 11/02 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
F01D 5/18 - Hollow bladesHeating, heat-insulating, or cooling means on blades
F01D 11/04 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
The invention concerns a rotary assembly for a turbomachine, comprising a disk of which the outer periphery is formed from an alternation of cavities and teeth (12), and blades extending radially from the disk and of which the roots (16) are engaged axially and held radially in the cavities of the disk. According to the invention, the teeth of the disk and the blade roots comprise, at the upstream and/or downstream axial ends of same, axial shoulders (74, 76) disposed circumferentially end-to-end in alternation and together forming a cylindrical surface (78) facing radially towards the inside of the disk.
axial sealing means upstream and/or downstream of an annular zone extending radially from the platforms (30) as far as the disc (16).
According to the invention, the sealing means comprise radially an internal annular part (64) and an external annular part (66) structurally separate from each other, and the facing radial ends of which are adapted for a relative radial movements by sliding sealingly, only an absorption of the centrifugal forces of the external part (66) being provided rotationally by the platforms (30).
An afterbody for a turbojet engine having a central axis, provided with a nozzle comprising two doors facing each other between two lateral beams. The doors pivot around axes defining a pivot direction, between a retracted position, in which a middle portion of the downstream edge of the doors forms the edge of the outlet section of the nozzle combined with the downstream edges of the two lateral beams, and a deployed position, in which the middle portions of the downstream edges of the pivoting doors come together so as to block the channel between the two lateral beams in order to reverse the thrust of the turbojet engine gases. The edge of the outlet section of the nozzle further having a crown of noise-reducing chevrons alternating with indentations and the afterbody.
A method for assembling two blades of a turbomachine nozzle, includes positioning a first surface of a first blade and a second surface of a second blade facing one another, the first and second surfaces being spaced apart from one another by an assembly clearance, and vapor phase aluminizing the first and second surfaces so as to fill the assembly clearance.
B23P 15/00 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass
B23P 15/04 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from several pieces
B23K 1/00 - Soldering, e.g. brazing, or unsoldering
B23K 1/20 - Preliminary treatment of work or areas to be soldered, e.g. in respect of a galvanic coating
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
B23K 101/00 - Articles made by soldering, welding or cutting
The present disclosure relates to a propulsion unit for an aircraft including a nacelle which surrounds a turbojet engine. The nacelle has an inner structure surrounding a downstream compartment of the turbojet engine, and the inner structure includes two annular half-portions. The propulsion unit also includes a rail/guide unit and to move the annular half-portions between a working position and a maintenance position. In particular, the rail/guide unit radially moves away the annular half-portions relative to a longitudinal axis of the nacelle, during a translation movement of the annular half-portions. The nacelle is provided with a connecting rod which is connected to the annular half-portions and to the turbojet engine and so that the connecting rod contributes to rotate the annular half-portions about the rail.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
B64D 29/00 - Power-plant nacelles, fairings or cowlings
A closer for an attachment orifice of an OGV in a turbine engine, the vane having a blade fixed to a root provided with attachment orifices, the closer including a cup adapted to be fixed in the attachment orifice of an OGV; a cylindrical shaped plug engaging in the cup, the plug including an upper face inclined relative to the ground plane of the plug; a guide adapted to position the plug on the cup by rotation into a working position, in which the upper face of the plug coincides with the surface level of the vane root; a blocking system to block the plug in the working position.
The invention relates to an aircraft turbine engine (1) comprising a fan (15) as well as a reduction gear (20) including a plurality of rotary elements (52, 58) and driving the fan, the turbine engine also including a gearbox (32) as well as a housing for drawing mechanical power (36) driving the gearbox, the turbine engine comprising a first gear (38) as well as a second gear (40) which is part of the housing for drawing mechanical power (36) and mates with the first gear. According to the invention, the first gear (38) is rotatably secured to a hub (60) of the fan (15).
The invention relates to a system (10) for supplying a turbomachine with fluid, the supply system (10) comprising a low-pressure pumping unit (101) intended to increase the pressure of the fluid flowing toward a downstream circuit (50, 60). According to the invention, the downstream circuit (50, 60) divides at an inlet node (E), situated between the low-pressure pumping unit (101) and the high-pressure volumetric pump (102), into a circuit (60) supplying an injection system (62) and a variable geometries supply circuit (50). The circuit (60) supplying the injection system comprises a high-pressure volumetric pump (102). The variable geometries supply circuit (50) is configured to convey the fluid toward variable geometry devices (54) from the inlet node (E) to an outlet node (S) connecting the variable geometries supply circuit (50) to the upstream circuit (100) between two pumps (101a, 111a) of the low-pressure pumping unit.
The invention relates to an assembly (100) for an aircraft turbine engine, comprising: a rolling bearing bracket (70) defining an inner space (78) on either side thereof; a rotary assembly (58, 60, 38) comprising a first gear (38); a housing for drawing mechanical power (36) comprising a second gear (40) mating with the first gear; a shaft for drawing mechanical power (42), inserted into the housing (36) and rotated by the second gear (40), the shaft (42) passing through a first opening (96) of the bracket (70). According to the invention, the assembly includes means (91) for mounting the housing (36) on the bearing bracket (70), said means passing through a second opening (95) via the bracket (70), said second opening (95) being configured so as to allow the housing (36) to be inserted into the space (78).
The internal cooling of the moving blades of the turbines in aircraft turbomachines is limited in effectiveness because of inhomogeneities of this cooling on each of the pressure-face and suction-face walls. To address this problem, there is proposed a blade comprising a circuit (50) for cooling the airfoil part (34) thereof, in which circuit the cavities interconnected in series are such that the stream of air flows radially toward the outside along the pressure-face wall (40) in pressure-face cavities (52, 56), and radially toward the inside along the suction-face wall (42) in a suction-face cavity (54) that is separated from the pressure-face cavities by an internal wall (58) of the airfoil part. In this way, the force of the Coriolis effect deflects the stream of air toward each of the pressure-face and suction-face walls thereby limiting the inhomogeneity.
The invention relates to a method for synchronising the engines of an airplane, according to an activation logic having a deactivated state, an armed state and an activated state, in which: the switching of the activation logic from the armed state to the activated state is carried out via a first and then a second successive intermediate state; every instance of the activation logic switching from the intermediate state to the activated state involves the following: taking into consideration, on each engine, the activation state of the synchronisation, and exchanging said data between the engines; the switching of the activation logic of one of the engines to the activated state requires that the safety and activation conditions of the other engines are all met; if one of the engines enters the deactivated state, the others do so as well; and, for each engine, the switching of the activation logic from the armed state to the first and then to the second intermediate state takes place automatically when a first portion and then a second portion of the safety and/or activation conditions are met.
A fuel gear pump (4') works to supply a pre-defined flow rate, but with low or zero pressure rise. In order to ensure the hydrodynamic lifting of bearings (19) supporting the pinions (11), or some of them, a sealing lining (46) is added between them, in order to delimit a closed cavity (47), in order to ensure lifting by a fluid having more suitable viscosity properties, instead of the lifting being ensured by the fuel itself which is pumped. Possible application to fuel pumps of aircraft engines, in which the pump (4') is a high-pressure pump combined with a low-pressure pump.
F02C 7/236 - Fuel delivery systems comprising two or more pumps
F04C 2/08 - Rotary-piston machines or pumps of intermeshing-engagement type, i.e. with engagement of co-operating members similar to that of toothed gearing
F04C 11/00 - Combinations of two or more machines or pumps, each being of rotary-piston or oscillating-piston typePumping installations
F04C 15/00 - Component parts, details or accessories of machines, pumps or pumping installations, not provided for in groups
28.
Turbine engine blade made of composite material with a bulb-shaped root
A turbine engine blade made of composite material including fiber reinforcement obtained by three dimensionally weaving yarns and densified with a matrix, the blade including an airfoil and a blade root forming a single part. The blade root includes two opposite lateral flanks that are substantially plane and that are clamped between two independent pads made of composite material, which pads are fastened against the lateral flanks of the blade root to form a blade root that is bulb-shaped.
The invention relates to fiber reinforcement for making an elongate mechanical part (10) out of composite material having a lug (14) at at least one end for receiving a pin for making a pivot connection with another part, the fiber reinforcement being made from a central fiber structure (106; 106′) for forming a core that is obtained by three-dimensional weaving, from a peripheral fiber structure (16; 16′) that is to form a belt surrounding the central structure so as to form at least one empty cylindrical space in the lug of the part, and from at least one annular fiber structure (112; 112′; 112″) that is to form a ring that is formed inside the empty space provided between the central structure and the peripheral structure.
B64C 25/14 - Undercarriages non-fixed, e.g. jettisonable retractable, foldable, or the like fore-and-aft
B64C 25/10 - Undercarriages non-fixed, e.g. jettisonable retractable, foldable, or the like
B32B 5/02 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by structural features of a layer comprising fibres or filaments
B32B 5/22 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by the presence of two or more layers which comprise fibres, filaments, granules, or powder, or are foamed or specifically porous
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
F16C 7/02 - Constructions of connecting-rods with constant length
B29L 31/30 - Vehicles, e.g. ships or aircraft, or body parts thereof
30.
BLADE FOR A TURBINE ENGINE PROPELLER, IN PARTICULAR A PROPFAN ENGINE, PROPELLER, AND TURBINE ENGINE COMPRISING SUCH A BLADE
The invention relates to a blade (11A) for a turbine engine propeller, in particular a propfan engine, comprising a protruding part (16) on the leading edge (17) thereof, characterised in that said blade comprises means for controlling the position of the protruding part along the leading edge thereof.
Mandrels (1, 2) for the rolling of a ring (7) comprise impressions (5, 6) to accommodate the ring, which impressions have fillet radii where the various faces meet, so as to avoid the creation of sharp corners on the rolled ring (7) and the accidental formation of cracks.
A method of fabricating a composite material casing for a gas turbine engine, the method including making an outer shroud including a platform and a flange, making an inner shroud of smaller diameter than the outer shroud and including a platform and a flange, making a plurality of casing arms, each including a blade that is terminated at each radial end by a respective platform, arranging a plurality of openings in the respective platforms of the shrouds, each opening serving to receive a platform of a casing arm, and assembling the casing arms with the outer shroud and with the inner shroud by inserting the platforms of the casing arms in the openings of the shrouds. A composite material casing is obtained by such a method.
The invention concerns a turbomachine bearing housing (E) comprising a fixed wall (9), a rotating shaft (5), first and second seals (10, 20) between the wall and the shaft, and a chamber (Cam) between the fixed wall (9) and a stator element (19) supplied with air via an opening (19a) close to said shaft (5). The housing is characterised by the fact that an air guide means (30) is arranged along the surface of the wall (9) of the housing, outside same, such that at least a portion of the air exiting the guide means passes between the first seal (10) and the shaft, said guide means being supplied with air by an air intake separated radially from the shaft, the air from the air intake being at a higher pressure than at the shaft.
F01D 11/04 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
The invention concerns a turbomachine (1) characterised in that it comprises: - an exhaust housing (7), comprising a plurality of arms (10), the space separating the arms defining openings (13) in which there circulates a primary air flow (29) of the turbomachine (1), - at least one conduit (2), a) configured to collect a compressed air flow at one of the ends (3) of same, b) the other end of the conduit (2) being connected to at least one opening (13) of the exhaust housing (7), so as to insert the collected air flow into said primary air flow (29), said collected air flow having, when inserted into the opening (13), a Mach number less than or equal to 0.5.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
35.
MOBILE MEMBER OF A TURBOMACHINE WHICH COMPRISES MEANS FOR CHANGING THE RESONANCE FREQUENCY OF SAME
The invention proposes a rotor (10) of an aircraft turbomachine having a main axis A, which comprises means (14) for modifying the critical speed of the rotor (10) depending on whether the rotational speed of the rotor (10) is lower or higher than a predefined rotational speed, comprising: a component (16) that is capable of occupying a first state or a second state depending on whether the rotational speed of the rotor (10) is lower or higher than the predefined rotational speed, each state of the component (16) corresponding to a critical speed of the rotor (10); and means (18) for driving the component (16) to one or the other of the two states thereof, depending on the rotational speed of the rotor (10), characterised in that the means (14) for modifying the critical speed of the rotor (10) further comprise a component (38) that engages with the drive means (18) and is capable of being deformed elastically between one or the other of two stable forms, each of which corresponds to a state of said component (16).
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
(1) Compresseurs de véhicules moteurs; partie de moteurs autres que pour véhicules terrestres nommément échangeurs thermiques; pompes et pompes à fluide nommément pompes à gasoline, pompes autorégulatrices à combustible, pompes centrifuges, pompes comme pièces de machines et de moteurs; turbopompes et vannes pour fluides, propulseurs à hydrazine, propulseurs bi-ergols, propulseurs à puissance augmentée par catalyse, propulseurs iioniques, propulseurs électriques, propulseurs à plasma et propulseurs à arc nommément propulseurs de fusées et d'avions; accouplements et organes de transmission et de propulsion, à l'exception de ceux pour véhicules terrestres, moteurs en tous genres pour installations fixes et mobiles, à l'exception des moteurs pour véhicules terrestres, nommément turbomachines à gaz, moteurs thermiques, électrothermiques, à réaction, à énergie nucléaire, à propulsion combinée, à propulsion par fusées; sources de plasma, à savoir machines pour le dépôt de revêtement et pour le soudage telles que machines de métallisation, robots de soudage; ensembles d'éjection pour moteurs d'aéronefs nommément tuyères, inverseurs de poussée; nacelles et inverseurs de poussée de moteurs d'avions; partie constitutives de ces machines nommément parties constitutives de moteurs d'avions et de fusées, propulseurs d'avions et de fusées, pompes nommément pompes à gasoline, pompes autorégulatrices à combustible, pompes centrifuges, pompes comme pièces de machines et de moeurs, ensembles d'éjection nommément tuyères et inverseurs de poussée, nacelles et inverseurs de poussée d'avions et de fusées. (1) Aide à la gestion et à l'exploitation d'entreprises industrielles et commerciales détenant une flotte d'aéronefs; services de conseils en matière de gestion commerciale d'un parc de moteurs, systèmes, équipements et pièces d'aéronefs; services de conseils en matière de gestion administrative et commerciale de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; services de conseils en matière de définition et choix des outillages dans le domaine de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; gestion administrative et commerciale des matériels de rechange pour utilisateurs de moteurs, systèmes, équipements et pièces d'aéronefs; compilation et études statistiques de données relatives à la gestion de la maintenance d'un parc de moteurs d'aéronefs; recueil de données dans un fichier central; gestion et compilation de bases de données; analyse, rassemblement, systématisation, gestion, traitement et stockage de données; exploitation de données commerciales nommément sélection, triage et mise en valeur des données textes, sons et images fixes et animées dans le domaine de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; exploitation de banques de données commerciales dans le domaine de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; services de fourniture nommément établissement de données statistiques nommément compilation de statistiques.
(2) Réparation, entretien, maintenance, reconditionnement de moteurs, de nacelles, d'inverseurs, d'ensemble d'éjection d'aéronefs; réparation d'appareils, d'instruments, de systèmes, de matériel et d'équipements électriques et électroniques utilisés dans le domaine aéronautique; assistance par tout moyen informatique et de télécommunication nommément assistance technique par le biais d'une ligne téléphonique et par le biais d'un site Internet accessible 7 jours sur 7 et 24 heures sur 24 en matière de remise en état et d'échanges standard de moteurs et de modules de moteurs utilisés dans le domaine aéronautique.
(3) Conseils techniques dans le domaine aéronautique (travaux d'ingénieurs); service d'ingénierie opérationnelle pour compagnies aériennes; service de conseils techniques en matière de méthodologies à utiliser dans le cadre de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance pour les moteurs, systèmes, équipements et pièces d'aéronefs; essais de machines; essais de matériaux; service d'inspection et de surveillance des moteurs, systèmes, équipements et pièces d'aéronefs; acquisition et déchargement de données de vol d'aéronefs; services d'analyse, d'expertise et de traitement de l'acquisition de données enregistrées lors du fonctionnement des moteurs, systèmes, équipements et pièces d'aéronefs; élaboration et conception de logiciels et programmation informatique.
37.
FIRE PROTECTION OF A PART MADE OF A THREE-DIMENSIONAL WOVEN COMPOSITE MATERIAL
The invention relates to a method for fire protection (S) of a part (1) of a gas-turbine engine made of a composite material comprising a main fibrous reinforcement compregnated by a main matrix, the protection method (S) comprising the following steps: preforming (S1) a panel of prepreg (20) such as to grant same a shape corresponding to the shape of a surface (3) of the part (1) to be protected against fire, said panel of prepreg (20) comprising a secondary fibrous reinforcement compregnated by a secondary matrix; applying (S2) the panel of prepreg (20) thus preformed to the part (1); and securing (S3) the panel of prepreg (20) to the surface (3) by thermal treatment of the part (1) provided with said panel of prepreg (20) in order to obtain a fire-protection layer (2).
Method (S) for protecting against fire a fan casing (1) comprising a roughly cylindrical barrel (10) having a main direction extending along a longitudinal axis (X) and an upstream flange (20) extending radially with respect to the longitudinal axis (X) from an upstream end of the barrel (10), the fan casing (1) being made of a composite comprising a fibrous reinforcement densified by a matrix, said matrix being polymerized, the protection method(S) comprising the following steps: - laying (SI) widths containing glass fibre pre-impregnated with a resin capable of affording the fan casing with thermal protection against fire on an upstream radial face (22) of the upstream flange (20), and - polymerizing (S2) the resin in order to obtain a protective layer (2).
The invention pertains to a method for monitoring an aircraft engine retractable doors thrust reverser, the thrust reverser being a reverser having hydraulic actuators and being provided with contactors (3a, 4a, 5a, 3B, 4b, 5b, Sa, Sb) arranged so as to each return an item of information about the position of the doors, the engine comprising a computer (3) configured to carry out measurements (E1) of a parameter representative of the position of the contactors on the basis of the information returned by the contactors, characterized in that it comprises a calculation (E2) of one or more statistical indicators of the parameter measured and an analysis (E3) of the temporal evolution of the statistical indicator or indicators calculated. The invention extends also to a computer program for the implementation of this method.
The present invention relates to an epicyclic reduction device (70) for rotating a first set of blades of a turbomachine, comprising a sun gear (74) centred on a longitudinal axis (12) of the turbomachine and connected to a rotor (76) of the turbomachine so as to be rotated; at least one planet gear (78) meshing with the sun gear; a planet gear carrier (80) rotationally bearing the planet gear and connected to a first set of blades (82) to rotate same; and an annulus gear (72) meshing with the planet gear; the sun gear being connected to the rotor by a first ball-type constant velocity joint (84).
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/072 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with counter-rotating rotors
41.
AIRCRAFT PROPULSION ASSEMBLY WITH FIRE EXTINGUISHING SYSTEM
Aircraft propulsion assembly (10) comprising an engine (16), a nacelle (18) surrounding the engine, and a system for extinguishing a fire that may occur in the engine and/or in the nacelle, this extinguishing system comprising means (34) for supplying an extinguishant to at least one extinguishant distribution pipe (36) which opens into a cavity (32) of the engine and/or a cavity (26) of the nacelle, characterized in that it further comprises means (48) for supplying said at least one pipe with air so as to ventilate the or each cavity.
The main subject of the invention is a turbomachine air starter (10) characterized in that it comprises a first compartment (1) in which the so-called "non-driven" elements of the air starter (10) are situated, these corresponding to the elements rotationally driven only in the phase of starting the turbomachine, and a second compartment (2) in which the so-called "driven" elements of the air starter (10) are situated, these corresponding to the elements rotationally driven throughout the time that the turbomachine is in operation, including during the starting phase, the first compartment (1) comprising a lubricating oil receptacle (3) performing splash lubrication of the "non-driven" elements and the second compartment (2) comprising an AWC lubricating oil type cavity (4) for lubrication of the "driven" components which cavity is fed by the oil return from the "driven" components which are themselves fed with lubricating oil under pressure (Hp) supplied by the turbomachine, the receptacle (3) and the cavity (4) being internal to the air starter (10).
The invention relates to an aircraft turbomachine (10) comprising a nacelle and an engine (12) comprising at least one outflowing jet of air, characterised in that a heat exchanger (20) of the precooler type for supplying air to the aircraft is mounted in the nacelle, said exchanger comprising a primary circuit, the inlet of which is connected to means for taking compressed air from the engine and the outlet of which is connected to means for supplying air to the aircraft, and a secondary circuit supplied with air taken from said air flow.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
Schneider, Julien
Hild, Francois
Leclerc, Hugo
Roux, Stephane
Abstract
The invention relates to a method for characterising a part (10), including a step of obtaining an X-ray tomography image of the part, followed by a step (200) of correlating said image with a reference (20), characterised in that the correlation step (200) includes searching, in a predefined set (30) of transformations of X-ray tomography images, for a transformation (40) that minimises the difference (50) between the image and the reference in order to characterise (300, 350) the inside of said part (10).
G01N 23/046 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material using tomography, e.g. computed tomography [CT]
45.
HOUSING MADE FROM AN ORGANIC-MATRIX COMPOSITE MATERIAL PROMOTING THE DISCHARGE OF SMOKE
A gas turbine housing (100) made from an organic-matrix composite material comprising a reinforcement densified by an organic matrix delimits an inner volume. The housing comprises, on the inner face (101) of same, a structural portion (120) having a first face (120a) facing the inner face of the housing and a second opposing face (120b) defining a flow channel portion (102). Recesses (130) opening into the inner volume of the housing are present between the inner face (101) of the housing and the first face (120a) of the structural portion (120) facing said inner face of the housing. The recesses (130) allow the gases produced by the degradation of the resin of the housing in case of a fire to be discharged from the flow channel side.
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
The invention relates to a method for manufacturing a part coated with a protective coating, the method comprising the following step: forming a protective coating on the outer surface of a part by micro-arc oxidation treatment, the part comprising a niobium matrix which contains metal-silicide insertions, the current passing through the part being monitored during the micro-arc oxidation treatment in order to subject the part to a series of current cycles, the ratio (amount of positive charge applied to the part) / (amount of negative charge applied to the part) being, for each current cycle, 0.80 to 1.6.
The invention proposes a guide arm for guiding at least one element having an elongated shape (20), corresponding to a set of cables and/or pipes. The arm comprises an inner cavity (62) opening on the outside of the arm at each of the ends thereof, and in which the elements having an elongated shape can extend. According to the invention, this structure more particularly comprises - a frame (8) comprising a beam (18) linked to means (30, 36, 44) for holding the elements having an elongated shape on the outside and along the beam, and - a cover (54) of which the walls (56, 58) cover the holding means of the frame, and are engaged with the beam, in such a way as to form the inner cavity in which the elements having an elongated shape extend, shock-absorbing means (68) being arranged between the means (30, 36, 44) for holding the elements having an elongated shape (20), and the longitudinal walls (56, 58), in such a way as to reduce and damp the movements of the means for holding in the cavity (62).
F01D 9/06 - Fluid supply conduits to nozzles or the like
F16L 3/10 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets substantially surrounding the pipe, cable or protective tubing divided, i.e. with two members engaging the pipe, cable or protective tubing
F16L 3/233 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets specially adapted for supporting a number of parallel pipes at intervals for a bundle of pipes or a plurality of pipes placed side by side in contact with each other by means of a flexible band
48.
SUSPENSION OF A TUBULAR ELEMENT IN AN AIRCRAFT COMPARTMENT
The invention relates to an assembly including a structure (2) and a tubular element (6), isostatically mounted in the structure (2), said tubular element (6) comprising a first end (7) connected by at least four connecting rods (10a - 10b - 10c - 10d) to said structure, establishing four first degrees of freedom, and a second end (8) connected to said structure (2) by a linking means (20) establishing two second degrees of freedom, said connecting rods comprising a means for adjusting the length thereof. The invention relates, in particular, to an exhaust fitted to an auxiliary power unit in the compartment thereof and to the installation method for aligning same.
The invention relates to a device (100) for cleaning, and in particular for degritting or desanding, a turbomachine module (110), characterised by comprising: (i) means (102, 104) for isolating bearings of the module, by containment in a closed enclosure (106); (ii) means (112) for overpressurising said enclosure; (iii) means (114) for stripping material deposited in the walls of annular recesses of the module, for example by spraying compressed air onto said walls; and (iv) means (116) for sucking up the material thus stripped.
The invention relates to a turbine engine compressor, in particular of an aeroplane turboprop or turbofan, including a stator comprising an annular casing and at least one annular row of variable-pitch vanes, each vane comprising a radially external end including a pivot mounted in an opening of the casing and connected by a linking member to a control ring (38) capable of pivoting axially relative to the casing, the linking member comprising a first end attached to the pivot of the vane and a second end comprising a pin inserted in a hole (52, 58) of the control ring (38), characterised in that at least one (58) of the holes (52, 58) of the control ring (38), which is used for inserting the pins of the linking members, has an oblong shape and extends in the circumferential direction such as to enable the pin to move into said oblong hole (58), during the rotation of the control ring (38).
The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2), between the pressure face of the first blade (3I) and the suction face of the second blade (3E) has a non-axisymmetric surface (S) defining a plurality of fins (4) of substantially triangular section extending downstream of a leading edge (BA) of each of the blades (3I, 3E), each fin (4) being associated with a leading position and a trailing position on the surface (S), between which positions the fin (4) extends, such that: the leading position is situated at between 5% and 35% length relative to a chord of the blade (3I, 3E) extending from a leading edge (BA) to a trailing edge (BF) of the blade (3I, 3E); - the further a fin (4) is from the suction face of the second blade (3E), the further the leading position of said fin (4) is axially from the leading edge (BA) of the blades (3I, 3E).
Blade intended for a turbomachine impeller comprising N blades. At one end, the blade has a platform formed integrally with an airfoil of the blade. Over part of the axial extent of the blade, a section on a plane perpendicular to the axis (X) of the impeller of the flow path of the platform consists mainly of two straight-line segments arranged respectively on the two sides of the airfoil. These segments form, on each side of the airfoil, an angle of 90°-180°/N with respect to the radial direction. Figure 3.
The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2) has a non-axisymmetric surface (S) bounded by a first and a second end plane (PS, PR) and defined by at least two class C construction curves each one representing the value of a radius of said surface (S) as a function of a position between the pressure face of the first blade (3I) and the suction face of the second blade (3E) in a plane substantially parallel to the end planes (PS, PR), these including at least one upstream curve and one downstream curve; each construction curve being defined by at least one pressure face control end point and one suction face control end point such that: - the tangent to the downstream curve at the suction face control end point 20 is inclined by at most 5°; - any other tangent to a construction curve at a control end point is inclined by at least 5°.
Blade for a turbo machine impeller comprising a root, an air foil and a tip. The root and the tip comprise platforms having surfaces (15) on the side of the air foil, the surfaces respectively being referred to as the root and tip flow path. Each of these flow paths is made up of a pressure-face part and of a suction-face part which are situated respectively on the side of the pressure face and of the suction face and are separated by a crest curve (45, 65). Blade manufacture is made easier notably by virtue of the fact that any point on a first surface out of the pressure face and the suction face and any point on the root and tip flow path parts situated on the side of the first surface has a normal that makes an acute or right angle with respect to a direction referred to as the first direction of manufacture. Method for modelling the blade.
The invention relates to an assembly including a structure (2) and a tubular element (6), isostatically mounted in the structure (2), said tubular element (6) comprising a first end (7) connected by at least four connecting rods (10a - 10b - 10c - 10d) to said structure, establishing four first degrees of freedom, and a second end (8) connected to said structure (2) by a linking means (20) establishing two second degrees of freedom, said connecting rods comprising a means for adjusting the length thereof. The invention relates, in particular, to an exhaust fitted to an auxiliary power unit in the compartment thereof and to the installation method for aligning same.
Aircraft propulsion unit comprising a motor and a nacelle comprising a casing (16) of revolution delimiting a flow vein of an air flow, characterized in that this casing comprises at least two openings closed by removable and interchangeable panels (18), at least one of these panels carrying equipment (24) of the propulsion system.
The invention concerns a method and system for forecasting maintenance operations to be applied to an aircraft engine comprising a plurality of elements monitored by damage counters, each damage counter being limited by a corresponding damage ceiling, characterised in that it comprises: - processing means (7) suitable for simulating a consumption of said damage counters (C1-Cm) by iteratively pulling a series of simulation missions from a learning database (9) containing test missions, - processing means (7) suitable for determining, at each iteration, an accumulation of consumption of each of said damage counters until at least one counter counting damage related to a current simulation mission reaches the damage ceiling associated with same, - processing means (7) suitable for applying a maintenance strategy to said current simulation mission to determine maintenance indicators representative of the maintenance operations to be planned on the aircraft engine.
The invention concerns a cutting table (100) for cutting a fibrous preform obtained by three-dimensional weaving and comprising two portions that are linked together by at least one separating area and that have contours of different shapes, the cutting table comprising a plate (104) provided with a cavity (108) intended to receive, flat, one of the portions of the preform to be cut, sacrificial plates (110) intended to be interposed between the portions of the preform to be cut and to be secured to the plate, at least one cutting template (114) intended to be applied to the portion of the fibrous preform that is not positioned in the cavity, and means (118) for applying a compacting pressure to the cutting template. The invention also concerns a method for cutting a fibrous preform using such a cutting table.
The invention relates to a method for mounting an engine module (1) in a mount comprising a first structure (3) and a second structure (2) that is offset relative to the first structure (3), which method is intended to position at least a determined part (4) of the engine module (1) with respect to an element (6) of the second structure (2) using a statically determinate suspension connecting the engine module (1) by first connecting rods (10a, 10b, 10c, 10d) to said first structure (3) and by second connecting rods (10e, 10f) to said second structure (2), the length of said first and second connecting rods being defined beforehand, in which method the length of at least two (10c, 10f) of said first and second connecting rods is adjusted relative to the predefined length thereof in order to position said determined part (4) of the engine module (1) relative to said element (6) of the second structure (2) in said mount. The invention also relates to said mount with adjustable connecting rods in an installation notably comprising an auxiliary power unit for an aircraft.
The invention relates to a method for mounting an engine module (1) in a mount comprising a first structure (3) and a second structure (2) that is offset relative to the first structure (3), which method is intended to position at least a determined part (4) of the engine module (1) with respect to an element (6) of the second structure (2) using a statically determinate suspension connecting the engine module (1) by first connecting rods (10a, 10b, 10c, 10d) to said first structure (3) and by second connecting rods (10e, 10f) to said second structure (2), the length of said first and second connecting rods being defined beforehand, in which method the length of at least two (10c, 10f) of said first and second connecting rods is adjusted relative to the predefined length thereof in order to position said determined part (4) of the engine module (1) relative to said element (6) of the second structure (2) in said mount. The invention also relates to said mount with adjustable connecting rods in an installation notably comprising an auxiliary power unit for an aircraft.
B64D 41/00 - Power installations for auxiliary purposes
B64F 5/00 - Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided forHandling, transporting, testing or inspecting aircraft components, not otherwise provided for
B64D 27/00 - Arrangement or mounting of power plants in aircraftAircraft characterised by the type or position of power plants
61.
METHOD FOR IMPREGNATION OF A FIBROUS PREFORM AND DEVICE FOR IMPLEMENTATION OF THE SAID METHOD
A method for impregnation of a fibrous preform (10) by an impregnation composition (20), the method comprising the following step: a) application of a liquid (30) onto a structure, the structure comprising: - a chamber (2) in which a fibrous preform (10) to be impregnated is present, the chamber (2) being defined between a rigid support (3) on which the fibrous preform (10) is placed and a wall (4), the wall (4) comprising a face (4a) located facing the fibrous preform (10), and - an impregnation composition (20), intended to impregnate the fibrous preform (10), the impregnation composition being present in the chamber (2), the liquid (30) being applied on the wall (4) of the opposite side of the chamber (2), the wall (4) being configured so that the face (4a) located facing the fibrous preform (10) retains its shape during application of the liquid (30), the applied liquid (30) enabling creation of a sufficient pressure to displace the wall (4) towards the rigid support (3) and impregnating the fibrous preform (10) with the impregnation composition (20).
B29C 43/10 - Isostatic pressing, i.e. using non-rigid pressure-exerting members against rigid parts or dies
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B30B 1/00 - Presses, using a press ram, characterised by the features of the drive therefor, pressure being transmitted directly, or through simple thrust or tension members only, to the press ram or platen
62.
DRAINED FLUID EVACUATION STUB FOR A PROPULSION ASSEMBLY
Drained fluid evacuation stub (16) for a propulsion assembly (10), comprising a drained fluid storage cavity and at least one orifice (32) for evacuation of the fluids contained in said cavity, characterized in that it comprises means (36, 38) for detecting a pressure difference with the exterior of the stub and a component for purging the cavity which is movable between a first closed position of the evacuation orifice and a second release position of the orifice, the component being configured to move from the first to the second position when the pressure difference is greater than or equal to a predetermined value.
The invention relates to a device (110) for retaining drained fluids for a propulsive assembly, comprising a cavity for storing the drained fluids and two walls (118, 120) mounted at the opening of said cavity, the cavity having a fluid storage volume V1 when the device is in a substantially vertical position, and each wall being configured such as to define a fluid storage volume (V2 and V3 respectively) in the cavity when the device is in a substantially horizontal position, each of the volumes V2 and V3 being at least equal to the volume V1. The invention also relates to a propulsive assembly comprising a device for retaining drained fluids.
Motive flow valve for aircraft engine comprising: a shaft comprising high-pressure and low-pressure chambers, a fuel inlet and outlet, a shoulder arranged between the inlet and outlet, defining an abutment surface, a drawer displaceable relatively to the shaft under a pressure difference between the high- and low-pressure chambers, between a closed and open position wherein the drawer obstructs and releases the fuel inlet, the drawer comprising a portion extending into an intermediate chamber, and a channel exiting on the one hand in a first portion of the intermediate chamber and communicating with the fuel outlet in the open and closed positions, and on the other hand in a second portion of the intermediate chamber that only communicates with the fuel inlet or outlet via the channel in the open and closed positions, and a sealing element interposed between the abutment surface and the drawer in the closed position.
F02C 7/232 - Fuel valvesDraining valves or systems
F16K 3/26 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing with sealing faces shaped as surfaces of solids of revolution with cylindrical valve members with fluid passages in the valve member
F16K 3/30 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing Details
F16K 17/04 - Safety valvesEqualising valves opening on surplus pressure on one sideSafety valvesEqualising valves closing on insufficient pressure on one side spring-loaded
65.
BALANCED TURBINE ENGINE PORTION AND TURBINE ENGINE
The invention relates to a balanced turbine engine portion. Said portion comprises at least one angular section (21) arranged such as to form a balancing ring (20) centred on a ring axis (C). Said angular section (21) comprises a plurality of attachment elements (30), a bearing surface (5) with a shape that matches the balancing ring (20), the angular section (21) abutting with said bearing surface (5). Said portion also comprises a plurality of balance weights (40), each attached to the corresponding attachment element (30) of the angular section (21), at least one of said balance weights also being useful as an attachment means for attaching the angular section (21) to the bearing surface (5). The invention also relates to a turbine engine comprising such a balanced portion.
The invention relates to a device (10) for centring and guiding the rotation of a turbine engine shaft, in which the outer ring (18) of a bearing is retained axially upstream and downstream by retaining means (52, 72) that engage with a bearing mounting (20) and with coupling means (29) including resiliently deformable means (32) connecting the outer ring to the bearing mounting, said retaining means being separate from a binding band (28) of the device. The invention also provides a method for assembling such a device in which the retaining means (52, 72) are pre-assembled with the coupling means (29) prior to the final assembly of the coupling means with the bearing mounting (20). The device and the method have the combined advantages of axially retaining the outer ring in two opposing directions and having a particularly straightforward assembly.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
67.
DEVICE FOR GUIDING SYNCHRONIZING RING VANES WITH VARIABLE PITCH ANGLE OF A TURBINE ENGINE AND METHOD FOR ASSEMBLING SUCH A DEVICE
The invention relates to a device for guiding synchronizing ring vanes with variable pitch angle of a turbine engine, including a plurality of angular inner ring sectors placed end-to-end to form an inner ring (26), each inner ring sector including shafts (24) passing radially from one side of the inner ring sector to the other, a plurality of cylindrical bushes (22) which are each mounted in a shaft of the inner ring from the inside and which are each intended for receiving a guiding pivot (12) of a synchronizing ring vane (4), a plurality of angular reconstitution ring sectors which are placed end-to-end to form a reconstitution ring (36) and which are mounted radially from the inside on the inner ring, and a plurality of locking elements passing axially through the inner and reconstitution rings such as to assemble said rings together. The invention also relates to a method for assembling such a device.
The invention relates to a fan, in particular for a small turbine engine such as a jet engine, having a hub ratio corresponding to the ratio of the diameter of the inner limit of the air intake section (26) at the radially internal ends of the leading edges of the fan blades (10), divided by the diameter of the circle through which the outer ends of the fan blades pass, which has a value of 0.25 to 0.27.
F01D 5/32 - Locking, e.g. by final locking-blades or keys
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps
The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream (26) at the radially inner ends of the leading edges of the fan blades (10), divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.
Fan, in particular for a turbine engine The invention relates to a fan, in particular for a small turbine engine such as a jet engine, having a hub ratio corresponding to the ratio of the diameter of the inner limit of the air intake section (26) at the radially internal ends of the leading edges of the fan blades (10), divided by the diameter of the circle through which the outer ends of the fan blades pass, which has a value of 0.20 to 0.265.
The invention relates to a system (2) for generating non-propulsive energy in an aircraft, including: an auxiliary power unit (20) including a gas turbine (21) and a fuel cell (22); a pathway (23) for the intake of outside air into the aircraft; an exhaust pipe (24) of the gas turbine, the system being characterized in that the air intake pathway (23) includes a pipe (230) for cooling the fuel pipe, in that the pipe is in fluid communication with the exhaust pipe (24) of the gas turbine such that the ejection of the gas from the gas turbine into the exhaust pipe causes a suction of the air outside the aircraft into the cooling pipe (230) by Venturi effect. The invention also relates to a method for generating non-propulsive energy.
The invention relates to a system (2) for generating non-propulsive energy in an aircraft, including: an auxiliary power unit (20) including a gas turbine (21) and a fuel cell (22); a pathway (23) for the intake of outside air into the aircraft; an exhaust pipe (24) of the gas turbine, the system being characterized in that the air intake pathway (23) includes a pipe (230) for cooling the fuel pipe, in that the pipe is in fluid communication with the exhaust pipe (24) of the gas turbine such that the ejection of the gas from the gas turbine into the exhaust pipe causes a suction of the air outside the aircraft into the cooling pipe (230) by Venturi effect. The invention also relates to a method for generating non-propulsive energy.
09 - Scientific and electric apparatus and instruments
35 - Advertising and business services
36 - Financial, insurance and real estate services
37 - Construction and mining; installation and repair services
38 - Telecommunications services
42 - Scientific, technological and industrial services, research and design
Goods & Services
(1) Équipements de traitement des informations nommément ordinateurs, matériel informatique de traitement des données; logiciels et programmes informatiques fournissant des informations à propos des moteurs d'aéronefs, du fonctionnement, de la maintenance, de l'intervention, de la réparation en relation avec les moteurs d'aéronefs et leurs modules et pièces; logiciels de gestion du fonctionnement et des opérations de réparation, révision, entretien, maintenance et remise en état effectuées sur un moteur d'aéronefs et sur ses modules et pièces; supports d'enregistrement magnétiques, optiques et numériques nommément CDrom, DVD, cassettes audio et vidéo et bandes magnétiques vierges et préenregistrés contenant des données textes, sons et images fixes et animées, logiciels et informations pour la conception, fabrication, maintenance, monitoring et réparation de moteurs aéronautiques et de leurs parties constitutives; banques de données électroniques et informatiques contenant des informations concernant les moteurs d'aéronefs, leur fonctionnement, la réparation, la révision, l'entretien, la maintenance et la remise en état de moteurs, modules et pièces de moteurs pour aéronefs; banques de données contenant des informations concernant les moteurs d'aéronefs, leur fonctionnement, la réparation, la révision, l'entretien, la maintenance et la remise en état de moteurs, modules et pièces de moteurs pour aéronefs; appareils pour l'enregistrement, la transmission, la reproduction du son, d'images et de données nommément systèmes et équipements électroniques embarqués et non, d'acquisition et de traitement de paramètres et de données textes, sons et images fixes et animées; tous ces produits étant utilisés et destinés au domaine aéronautique. (1) Aide à la gestion et à l'exploitation d'entreprises industrielles et commerciales détenant une flotte d'aéronefs; services de conseils en matière de gestion commerciale d'un parc de moteurs, systèmes, équipements et pièces d'aéronefs; services de conseils en matière de gestion administrative et commerciale de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; services de conseils en matière de définition et choix des outillages dans le domaine de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; gestion administrative et commerciale des matériels de rechange pour utilisateurs de moteurs, systèmes, équipements et pièces d'aéronefs; compilation et études statistiques de données relatives à la gestion de la maintenance d'un parc de moteurs d'aéronefs; recueil de données dans un fichier central; gestion et compilation de bases de données; analyse, rassemblement, systématisation, gestion, traitement et stockage de données nommément systématisation de données dans un fichier central; exploitation pour le compte de tiers de données commerciales contenant de l'information technique, financière, des bulletins d'informations financières et de l'information sur les clients et les contrats nommément production, analyse, sélection, triage et mise en valeur des données dans les domaines aérospatial, de la défense et de la sécurité civile; exploitation de banques de données commerciales pour le compte de tiers contenant de l'information technique, financière, des bulletins d'information dans les domaines aérospatial, de la défense et de la sécurité civile; services de fourniture nommément établissement de données statistiques nommément compilation de statistiques; tous ces services étant utilisés et destinés au domaine aéronautique.
(2) Assurances; affaires financières nommément informations, consultations, analyse et gestion financières dans le domaine aéronautique; affaires monétaires nommément analyse et gestion monétaires dans le domaine aéronautique; consultations, estimations, analyses et expertises en affaires financières; études et montages de projets financiers ainsi que la fourniture de conseils et d'assistance dans ce domaine; ingénierie financière, montage de dossiers de financement et accompagnement de projets financiers; financement nommément prêts pour systèmes, équipements et pièces d'aéronefs; services financiers concernant la location de systèmes, équipements et pièces d'aéronefs; services d'assurances de systèmes, équipements et pièces d'aéronefs; services de garantie pour systèmes, équipements et pièces d'aéronefs; crédit bail nommément service de location vente de systèmes, équipements et pièces d'aéronefs; tous ces services étant destinés aux domaines aéronautique et spatial.
(3) Réparation et maintenance de véhicules aérospatiaux et de leurs pièces constitutives; construction et installation de bâtiments industriels et de bancs d'essais pour systèmes, équipements et pièces de véhicules aérospatiaux; réparation d'appareils, d'instruments, de systèmes, de matériels et d'équipements électriques et électroniques utilisés dans le domaine aérospatial; services de réparation, de révision, d'entretien et de maintenance pour des systèmes, équipements et pièces de véhicules aérospatiaux et de leurs pièces constitutives, y compris de moteurs, turbines, nacelles de moteurs, réacteurs, propulseurs, inverseurs de poussée, systèmes de freinage, systèmes d'atterrissage, câblages; services de réparation, de révision, d'entretien et de maintenance sous l'aile pour tous types de systèmes, équipements et pièces de véhicules aérospatiaux; mise au standard, remise en état et échange standard de systèmes, équipements et pièces de véhicules aérospatiaux; service d'assistance en ligne 24h/24h et 7 jours sur 7 en matière de réparation, révision, de remise en état, d'entretien, de maintenance et d'échange standard pour les systèmes, équipements et pièces de véhicules aérospatiaux; conseils et assistance techniques concernant la fabrication, l'entretien, la réparation, la révision de véhicules aéronautiques et de leurs parties constitutives, y compris de moteurs, turbines, nacelles de moteurs, de réacteurs, de propulseurs, d'inverseurs de poussée, de véhicules aéronautiques, de systèmes de freinage, de systèmes d'atterrissage, de câblages; réparation des outillages et matériels de servitude pour systèmes, équipements et pièces de véhicules aérospatiaux.
(4) Fourniture d'accès à des bases de données électroniques et informatiques dans le domaine aéronautique; service de télécommunication nommément transmission d'informations contenues dans une banque de données et sur un serveur via un accès Internet et Intranet nommément diffusion d'informations dans le domaine de la réparation, la révision, l'entretien, la maintenance et la remise en état de moteurs d'aéronefs via une base de données informatiques; location de temps d'accès à un centre serveur de bases de données; tous ces services étant utilisés et destinés au domaine aéronautique.
(5) Services de développement nommément constitution de banques de données; services d'élaboration, conception, installation, maintenance, mise à jour de logiciels d'ordinateurs dans le domaine aéronautique; location de logiciels d'ordinateurs dans le domaine aéronautique; services scientifiques et technologiques et services de recherche et conception y relatifs, dans le domaine aéronautique à savoir: expertises nommément travaux d'ingénieurs; service d'analyse, d'expertise et de traitement de l'acquisition de données enregistrées lors du fonctionnement de moteurs d'aéronefs, de leurs modules et pièces; tous ces services étant utilisés et/ou destinés au domaine aéronautique.
74.
FRONT ENCLOSURE WHICH IS SEALED DURING THE MODULAR DISMANTLING OF A TURBOJET WITH REDUCTION GEAR
The invention relates to a turbofan engine comprising a fan (S) driven, via a fan shaft (3) supported by at least two first bearings (11, 12), by a turbine shaft (4) supported by at least one second bearing (10) comprising a stationary ring (25) and a movable ring (26), said turbine shaft driving said fan shaft (3) through a device for reducing the speed of rotation (7), said device for reducing the speed of rotation and said first and second bearings being housed in a lubrication enclosure (E1) in which the shell comprises stationary portions and movable portions connected to one another by sealing means (29, 30, 31), said reducing device comprising an inducer (27) shaped so as to receive the torque transmitted by said turbine shaft via driving means (8, 9) connected to said movable ring, characterised in that the lubrication enclosure forms a coaxial ring with the turbine shaft and said driving means (8, 9) comprise a girth gear which is part of the movable sealing walls of the shell of the lubrication enclosure (E1).
F02C 3/067 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages having counter-rotating rotors
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
75.
MODULAR ENGINE, SUCH AS A JET ENGINE, WITH A SPEED REDUCTION GEAR
The present invention relates to an engine (1) with a modular structure comprising a plurality of coaxial modules (A, B, C) with, at one end, a first module (A) comprising a power transmission shaft (3) and a speed reduction gear (7), said power transmission shaft being driven via the speed reduction gear (7) by a turbine shaft (2) secured to one (C) of said coaxial modules that is separate from the first module, the speed reduction gear comprising a drive means (8 and 9) fixed to the turbine shaft (2) and to a journal (13) of a shaft of a low-pressure compressor rotor (1 a), characterized in that it comprises a first nut (16) for fastening the drive means to the journal and a second nut (14) for fastening the drive means to the turbine shaft.
A YSZ-type ceramic layer is deposited on a tie sublayer by thermal spraying using a plasma arc torch, said tie sublayer being itself deposited on the part to be protected. A sintering post treatment is carried out by means of a sweep of the ceramic layer by the beam of the plasma arc torch, the temperature at the point of impact of the beam at the surface of the ceramic layer (C) being, during this sweep, between 1300°C and 1700°C.
C23C 4/08 - Metallic material containing only metal elements
C23C 4/10 - Oxides, borides, carbides, nitrides or silicidesMixtures thereof
C23C 4/12 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
The invention relates to a turbomachine (30) comprising: a compressor stage and a turbine stage, each stage comprising at least one disk (42); and a tubular shaft (31) sleeve (33) extending along the axis (32) of the turbomachine, wherein the sleeve (33) comprises at least one tab (40) extending from an outer radial surface (41) of the sleeve and facing the disk (42), the tab (40) being designed to come into contact with the disk (42) when the sleeve (33) is in rotation about the axis (32) of the turbomachine.
The present invention relates to a bypass engine bearing holder (1) that holds an upstream bearing (6) and defines, with said upstream bearing, an oil chamber (100) and an air chamber (200), comprising a frusto-conical portion (11) defining an upstream bearing chamber (160) and a downstream inner chamber (150), and comprises an outer collar (13) connected, by a weld (135), to a flange (15) that extends outward from the frusto-conical portion (11). The outer collar (13) has a sealable gimlet (131) engaging with the upstream bearing (6) such as to seal the oil chamber (100). The bearing holder (1) comprises a plurality of oil recovery ducts (8) leading to the downstream inner chamber (150) and to the upstream bearing chamber (160). The oil recovery ducts (8) lead to the upstream bearing chamber (160), downstream from the weld (135) of the outer collar (13) on the flange (15), the weld (135) of the outer collar (13) being axisymmetric.
The invention relates to a multi-point fuel injection device (1) for an aircraft engine (M), comprising an inlet line (10), at least two injection lines (11, 12), and a purge line (14), a fuel distributor member (2) connected to each line and comprising a moveable element (22) which comprises an injection passage (223), in which the moveable element (22) additionally comprises a purge passage (226), and is configured to adopt a first range of positions in which the injection passage (223) interconnects the inlet line (10) and the injection lines (11, 12), and a second range of positions in which the injection passage (223) interconnects the inlet line (10) and at least a first injection line (11) while the purge passage (226) interconnects the purge line (14) and at least a second injection line (12), the device being characterized in that it additionally comprises an actuator adapted to move the moveable element into a safety position when a failure of the distribution member is detected, the injection passage (223) interconnecting, in this safety position of the moveable element (22), the inlet line (10) and the first injection line (11) while the purge passage (226) does not interconnect the purge line (14) to any of the injection lines (12).
F02C 7/228 - Dividing fuel between various burners
F02C 7/232 - Fuel valvesDraining valves or systems
F02C 9/34 - Joint control of separate flows to main and auxiliary burners
F02M 41/04 - Fuel-injection apparatus with two or more injectors fed from a common pressure-source sequentially by means of a distributor the distributor being spaced from pumping elements the distributor reciprocating
F23D 11/00 - Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
F23K 5/06 - Liquid fuel from a central source to a plurality of burners
F23R 3/34 - Feeding into different combustion zones
80.
SEALING SYSTEM WITH TWO ROWS OF COMPLEMENTARY SEALING ELEMENTS
The invention concerns a sealing system, in a cavity (C) under a stator (10), of a turbomachine vein (VC, VT), the cavity (C) being located between a stator (10) vane (PS) root (SI) and an additional rotor member (11), the root (SI) comprising two surfaces (21, 24a) each provided with an abradable coating (22, 32), the rotor member (11) being provided with first and second sealing elements (23, 33), disposed respectively facing the first and second surfaces (21, 24a), the first surface (21) and the first sealing element (23) forming a first sealing pair (20) and together delimiting a first leakage section, the second surface (24a) and the second sealing element (33) forming a second sealing pair (30) and together delimiting a second leakage section, one of the two pairs (20, 30) moving to a minimum leakage section when the other (30, 20) moves to a maximum leakage section, and vice versa.
The invention relates to a device (1) for replacing machining inserts (900) on a tool (800) including a body (810) and a head (820) supporting at least one machining insert (900), each insert being maintained on the head (820) by a screw. The device (1) includes: a positioner (50) comprising a supporting element (59) capable of supporting the body (810) of the tool; a screwing station having a screwdriver (60) capable of screwing and unscrewing the screws, the positioner (50) being capable of moving the tool relative to the screwing station; a gripping device (70) capable of gripping and placing an insert (900); a conveyor (500) comprising a plurality of insert containers (510), along which a first station (100) having the gripping device (70), a second station (200) having a mechanism (20) for rotating the inserts on the axes thereof, a third station (300) having a mechanism (30) for unloading the inserts, and a fourth station (400) having a mechanism (40) for supplying inserts are distributed; a transport mechanism (80) capable of moving the gripping device (70) between the positioner (50) and the first station (100) of the conveyor (500); and a control centre (600) capable of automatically controlling one or more of the mechanisms and devices and/or the conveyor of the replacement device (1).
The invention relates to a turbine engine, comprising two structural annular casings (16, 22) connected to one another by means (40, 54) for absorbing stresses from the thrust of the engine, which include connecting rods (54), characterised in that said thrust-absorbing means also include at least one accessory gearbox (40) which is attached to a first one of said casings (16) and which is connected by said connecting rods to the other one of said casings (22).
The invention concerns a method for generating auxiliary power in an aircraft, comprising the step consisting of: starting up an auxiliary power unit (6) of the aircraft by supplying compressed air to the auxiliary power unit (6) from a supercharger (7), and transferring non-propulsive energy from the auxiliary power unit (6) to the aircraft. The invention also concerns a system (5) for generating auxiliary power in an aircraft and an aircraft implementing such a method.
The invention relates to a method for monitoring a locking system (1) comprising N locks (2a, 2b, 2c, 2d, 2e). Each lock (2a, 2b, 2c, 2d, 2e) is monitored by two locking sensors (3a, 4a, 3b, 4b, 3c, 4c, 3d, 4d, 3e, 4e). Each locking sensor is capable of indicating if the lock that said locking sensor is monitoring is in a locked or unlocked state. Each locking sensor can be in a valid or invalid state. The method comprises the following steps: - determining the state of the locking system (1) on the basis of the state of the locks detected by the locking sensors; and - determining a reliability level associated with the state of the locking system on the basis of the number of valid locking sensors monitoring the locks that are in the same state as the locking system.
A quenching agent delivery apparatus is provided for delivering a quenching agent to a component to be quenched. The delivery apparatus comprises an inlet through which the quenching agent is configured to be delivered into the apparatus, a first outlet configured to deliver quenching agent in a first direction to an inner surface of the component, and a second outlet configured to deliver quenching agent in a second direction to an inner surface of the component.
SOCIETE LORRAINE DE CONSTRUCTION AERONAUTIQUE (France)
Inventor
Poisson, Mathieu Ange
Orcel, Stephane
Glemarec, Guillaume
Pacary, Jean-Luc
Abstract
Turboprop air intake A turboprop (110), comprising a rotary propeller (112) upstream from an engine (114) and an air intake (116) that is not coaxial to the propeller, said air intake defining a conduit (119) for supplying air to the engine and further defining a bypass (124) to said conduit, the bypass comprising an outlet (126) oriented substantially axially towards the downstream of the engine, the turboprop further comprising a nacelle (130) surrounding the engine and the air intake, characterised in that the air intake is secured to a housing (123) of the engine and is not rigidly connected to the nacelle, so as to allow, during operation, relative movements between the air intake and the nacelle, said outlet (126) being connected by a flexible link (140) to an intake of an air circuit carried by the nacelle.
The present invention concerns a turbomachine part (1) comprising at least first and second blades (3, 31, 3E), and a platform (2) from which the blades (3, 31, 3E) extend, characterised in that the platform (2) has a non-axisymmetric surface (S) limited by first and second end planes (PS, PR), and defined by at least three construction curves (PC-A, PC-C, PC- F) of class C1 each representing the value of a radius of said surface (S) on the basis of a position between the lower surface of the first blade (31) and the upper surface of the second blade (3E) according to a plane substantially parallel to the end planes (PS, PR), including: - a first curve (PC-C) that increases in the vicinity of the second blade (3E); - a second curve (PC-F) disposed between the first curve (PC-C) and a trailing edge (BF) of the first and second blades (3, 31, 3E), and that decreases in the vicinity of the second blade (3E); - a third curve (PC-A) disposed between the first curve (PC-C) and a leading edge (BA) of the first and second blades (3, 31, 3E), and having a minimum at the second blade (31).
The invention relates to a method for the acoustic analysis of a machine (M), comprising the acquisition of at least one acoustic signal supplied by at least one microphone (7) positioned in the machine, characterised in that it comprises the following steps: separation of at least one acoustic signal into a plurality of sound sources, said signal being modelled as a mixture of components, each one corresponding to a sound source; for at least one separate sound source, determination of a characteristic acoustic signature; comparison of at least one characteristic acoustic signature with at least one reference acoustic signature recorded in a reference database (5).
G01M 15/12 - Testing internal-combustion engines by monitoring vibrations
G01N 29/14 - Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic wavesVisualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object using acoustic emission techniques
G01N 29/44 - Processing the detected response signal
G01N 29/46 - Processing the detected response signal by spectral analysis, e.g. Fourier analysis
89.
DEVICE FOR TRANSFERRING OIL BETWEEN TWO REPOSITORIES ROTATING RELATIVE TO EACH OTHER, AND PROPELLER TURBOMACHINE FOR AN AIRCRAFT WITH SUCH A DEVICE
The device (20) comprises two outer and inner concentric rings (22, 23), one of which is connected to an oil supply from one of the repositories, the other ring being connected to the other repository, the oil flowing between said rings, and bearings between the rings in order to change repositories between the two rings. According to the invention, the device (20) further comprises a flexible means (31) forming a shock absorber, provided between a first of said rings and an intermediate ring (41) that is separated from a second of said rings by said bearings (25), said flexible means (31) defining a deformable sealed chamber (32) in which oil travels between the two repositories.
The invention relates to a method and system for monitoring an aircraft engine (2), comprising: - acquisition and processing means (11) configured to collect a time signal for a residual temperature margin at the exhaust gas outlet from said aircraft engine (2), - acquisition and processing means (11) configured to smooth said time signal, thus forming a curve representative of said residual temperature margin, - acquisition and processing means (11) configured to identify descending pieces of said first curve, - acquisition and processing means (11) configured to build a second curve by concatenating said descending pieces, said second curve being continuous while being restricted to said descending pieces of said first curve, - acquisition and processing means (11) configured to build a prediction model from said second curve to determine at least one failure prognosis indicator.
36 - Financial, insurance and real estate services
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
(1) Aide à la gestion et à l'exploitation d'entreprises industrielles et commerciales détenant une flotte d'aéronefs; services de conseils en matière de gestion commerciale d'un parc de moteurs, systèmes, équipements et pièces d'aéronefs; services de conseils en matière de gestion administrative et commerciale de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; services de conseils en matière de définition et choix des outillages dans le domaine de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; gestion administrative et commerciale des matériels de rechange pour utilisateurs de moteurs, systèmes, équipements et pièces d'aéronefs; compilation et études statistiques de données relatives à la gestion de la maintenance d'un parc de moteurs d'aéronefs; recueil de données dans un fichier central nommément conception et gestion de bases de données électroniques et informatiques; gestion et compilation de bases de données; analyse, rassemblement, systématisation, gestion, traitement et stockage de données textes, images et sons, toutes enregistrées lors du fonctionnement de systèmes, équipements et pièces d'aéronefs; exploitation pour le compte de tiers de données commerciales contenant de l'information technique et financière, des bulletins d'informations financières et de l'information sur les clients et les contrats nommément production, analyse, sélection, triage et mise en valeur des données dans le domaine aérospatial; exploitation de banques de données commerciales nommément interrogation, stockage et modification de banques de données pour le compte de tiers contenant de l'information technique et financière, des bulletins d'information dans le domaine aérospatial; services de fourniture nommément établissement de données statistiques nommément compilation de statistiques
(2) Assurances; affaires financières et monétaires nommément évaluations financières à des fins d'assurance, gestion d'actifs; consultations, estimations, analyses et expertises en affaires financières nommément consultation et audit en matière de gestion du risque pour les compagnies aériennes et les exploitants d'aéronefs; études de projets financiers ainsi que la fourniture d'assistance dans ce domaine et montage de dossiers de financement et accompagnement de projets financiers nommément financement de crédit-bail, financement de prêts, financement de projets pour systèmes, équipements et pièces d'aéronefs; ingénierie financière; prêts (financement) pour systèmes, équipements et pièces d'aéronefs; services financiers concernant la location de systèmes, équipements et pièces d'aéronefs nommément services financiers de garantie et de cautionnement; services d'assurances de systèmes, équipements et pièces d'aéronefs nommément services d'assurance dans le domaine aéronautique et spatial; services de garantie pour systèmes, équipements et pièces d'aéronefs nommément garantie prolongée; services de location vente (crédit bail) de systèmes, équipements et pièces d'aéronefs, tous les services précités étant utilisés et destinés au domaine aéronautique et spatial
(3) Réparation et maintenance d'aéronefs; services de réparation, de révision, d'entretien et de maintenance pour des systèmes, équipements et pièces d'aéronefs; mise au standard, remise en état et échange standard de systèmes, équipements et pièces d'aéronefs; services de conseils en matière de réparation, révision, entretien, mise au standard, maintenance de moteurs, systèmes, équipements et pièces d'aéronefs; service d'assistance en ligne 24h/24h et 7 jours sur 7 en matière de réparation, révision, de remise en état, d'entretien, de maintenance et d'échange standard pour les systèmes, équipements et pièces de véhicules aéronautiques et spatiaux
(4) Conseils techniques dans le domaine aéronautique (travaux d'ingénieurs); service d'ingénierie opérationnelle pour compagnies aériennes; service de conseils techniques en matière de méthodologies à utiliser dans le cadre de la réparation, des révisions, de l'entretien, de la mise au standard, de la maintenance pour les moteurs, systèmes, équipements et pièces d'aéronefs; essais de machines nommément de systèmes, équipements et pièces d'aéronefs; essais de matériaux; service d'inspection et de surveillance des moteurs, systèmes, équipements et pièces d'aéronefs; acquisition et déchargement de données de vol d'aéronefs; services d'analyse, d'expertise et de traitement de l'acquisition de données enregistrées lors du fonctionnement des moteurs, systèmes, équipements et pièces d'aéronefs; élaboration et conception de logiciels et programmation informatique
92.
TURBOMACHINE COMBUSTION CHAMBER PROVIDED WITH AIR DEFLECTION MEANS FOR REDUCING THE WAKE CREATED BY AN IGNITION PLUG
In order to improve the cooling of an annular wall (13) of a turbomachine combustion chamber provided with microperforations (53) and, in particular, the cooling of a region of the wall facing a wake (52) caused by an ignition plug, deflector means (60, 68) are proposed, these being designed to deflect the air (34') bathing the ignition plug towards a mid plane (P) of the wake (52) and in the direction of the annular wall (13) of the combustion chamber so as to increase the pressure of the air within the wake (52) near the annular wall (13).
Compacting assembly comprising a shaping mould (24) delimiting a housing open at the top able to receive a precut woven preform (10a), and a compacting tool (128) that is able to move vertically and forms, with the shaping mould (24), a compacting assembly for compacting said preform placed beforehand in the housing. The compacting tool (128) comprises at least one root portion (128A). Application to the manufacture of turbomachine composite blades.
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29D 99/00 - Subject matter not provided for in other groups of this subclass
A fibrous structure (200) comprises a preform portion (210) formed as a single piece by three-dimensional weaving between a first plurality of layers of threads and a second plurality of layers of threads, the preform portion corresponding to all or part of a fibrous reinforcement preform for a component made of composite. The fibrous structure (200) comprises, outside of the preform portion (210), one or more layers of two-dimensional woven fabric (220a, 220b), each layer of two-dimensional woven fabric grouping together the threads (2010a) of one same layer (201a) belonging at least to the first plurality of layers of threads and situated outside of the preform portion (210).
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
D03D 25/00 - Woven fabrics not otherwise provided for
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
The invention relates to a fuel injector (10) such as an injector for an annular combustion chamber of a turbomachine, comprising a downstream head (16) having a central outlet (22) and an annular peripheral outlet (24) surrounding the central outlet (22), and an injector arm (12) upstream of the head (16) comprising a coaxial central channel (18) and a coaxial annular channel (20), characterised in that the central channel (18) is in fluid communication with the peripheral outlet (24) and the annular channel (20) is in fluid communication with the central outlet (22).
The invention relates to a combustion chamber for a turbine engine, including an annular bottom wall (18) provided with injection systems (20) each centred on a respective axis (24) and each having an upstream end forming a socket (26') intended for receiving a head of a fuel injector, and an annular fairing (40') covering said bottom wall (18) and including injector-passage openings (42) arranged respectively opposite said injection systems (20), wherein said annular fairing (40) comprises air-intake openings separated from said injector-passage openings (42), and said socket (26') of each injection system passes through the corresponding injector-passage opening (42) and includes, at the upstream end thereof, a flange (62) having a free end (64) separated from said axis (24) of the injection system by a first distance (d1) which is greater than a second distance (d2) separating an edge of said injector-passage opening and said axis.
The invention relates to a compacting assembly comprising a forming mould (24) defining an upwardly open housing that can receive a previously cut woven preform (10a), and a vertically mobile compacting tool (128), and forming, with the forming mould (24), an assembly for compacting said preform previously placed in the housing. The compacting tool (128) comprises at least one foot portion (128A). The compacting tool comprises at least three separate compacting blocks (1281-1287). The invention is applicable to the production of composite fan blades for turbomachines.
B29C 43/36 - Moulds for making articles of definite length, i.e. discrete articles
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
The invention concerns a rotary assembly for a turbomachine, comprising a disk of which the outer periphery is formed from an alternation of cavities and teeth (12), and blades extending radially from the disk and of which the roots (16) are engaged axially and held radially in the cavities of the disk. According to the invention, the teeth of the disk and the blade roots comprise, at the upstream and/or downstream axial ends of same, axial shoulders (74, 76) disposed circumferentially end-to-end in alternation and together forming a cylindrical surface (78) facing radially towards the inside of the disk.
The present invention concerns a hydraulic ram comprising a fixed support (30), a cylinder (24) that is movable in translation relative to the support, a piston (22) secured inside the cylinder delimiting two chambers with the cylinder (24) and a device for supplying the chambers with hydraulic fluid upstream from the fixed support (30, 31). The ram is characterised by the fact that the supply device comprises telescopic channels (25, 26, 28), each telescopic channel comprising two tubular elements sliding one into the other, a first tubular element being rigidly connected to the fixed support (31) at one end and the second tubular element being rigidly connected to the cylinder at at least two points separated from each other along a generatrix of the cylinder. The invention is applicable to controlling the pitch of the blades of a turbine engine propeller.
B64C 11/38 - Blade pitch-changing mechanisms fluid, e.g. hydraulic
B64C 11/48 - Units of two or more coaxial propellers
F15B 15/14 - Fluid-actuated devices for displacing a member from one position to anotherGearing associated therewith characterised by the construction of the motor unit of the straight-cylinder type
100.
METHOD FOR MONITORING THE ENERGY DENSITY OF A LASER BEAM BY IMAGE ANALYSIS AND CORRESPONDING DEVICE
The invention relates to a method for monitoring (S) the energy density of a laser beam (3) using parameters of the laser beam (3) including the following steps: regularly applying the laser beam (3) to a reference substrate and measuring (S4), with each application, the resulting light intensity; identifying (S6, S7, S8) a change in the light intensity on the reference substrate between at least two measurements; and, when the change in the light intensity is higher than a predetermined threshold, determining the unstable parameter or parameters of the energy density of the laser beam (3) (S8).
B01D 53/00 - Separation of gases or vapoursRecovering vapours of volatile solvents from gasesChemical or biological purification of waste gases, e.g. engine exhaust gases, smoke, fumes, flue gases or aerosols
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
B22F 5/04 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product of turbine blades
