Abstract

L'invention concerne un absorbant acoustique (100) s'étendant entre deux surfaces opposées (100a,100b) et combinant un ou plusieurs résonateurs acoustiques quart-d'onde (102,102') avec un élément microporeux (101). Chaque résonateur acoustique (102, 102') a une longueur L, sensiblement supérieure à une épaisseur t de l'absorbant acoustique (100), entre une première extrémité (102a) ouverte sur une première surface (100a) des deux surfaces opposées (100a,100b) de l'absorbant acoustique (100), et une deuxième extrémité (102b) fermée. L'élément microporeux (101) est constitué d'une pluralité de cellules unitaires périodiquement répétées et adjacent aux résonateurs acoustiques (102), (102'). L'invention concerne aussi un procédé de production de cet absorbant acoustique, comprenant au moins une étape de fabrication additive.

IPC Classes  ?

• G10K 11/172 - Methods or devices for protecting against, or for damping, noise or other acoustic waves in general using resonance effects
• B33Y 10/00 - Processes of additive manufacturing
• F02C 7/24 - Heat or noise insulation

Abstract

Disclosed is a structure (50) for linking and supporting a turbine engine (10) on an aircraft pylon (32), the structure (50) having a first longitudinal axis (C) intended to extend parallel to a second longitudinal axis (B) of the pylon (32), the structure (50) comprising: -a first axial portion (50a) for fastening to the pylon (32), said first portion (50a) having an upper end that defines a substantially horizontal plane (P3) of interface with the pylon (32), and a lower end bearing suspension rods (62, 64) for suspending the turbine engine (10), which extend in a rear vertical plane (P2); and -a second axial portion (50b) that is intended to extend forward of the pylon (32) and comprises at least one suspension member (66) of the turbine engine (10) in a front vertical plane (P1).

IPC Classes  ?

• B64D 27/40 - Arrangements for mounting power plants in aircraft
• B64D 27/12 - Aircraft characterised by the type or position of power plant of gas-turbine type within, or attached to, wing
• B64D 27/18 - Aircraft characterised by the type or position of power plant of jet type within, or attached to, wing

Abstract

L'invention concerne le domaine des méta-matériaux acoustiques et en particulier un procédé de fabrication d'un méta-matériau acoustique (100) comportant une pluralité de colonnes (101) s'étendant à partir d'une base commune, ainsi que le méta-matériau acoustique (100) obtenu par ce procédé. Ce procédé comprend une étape fabrication additive d'un moule (210) par dépôt d'une pluralité de strates empilés (203) comprenant chacun une pluralité de cellules (204) périodiquement répétées, séparées par des parois (205), les cellules (204) de la pluralité de strates empilés (203) étant alignées de manière à former des canaux (206), une étape de remplissage des canaux (206) avec un matériau fluide (220), une étape de solidification du matériau fluide (220), et une étape d'élimination du moule (210).

IPC Classes  ?

• B29C 64/118 - Processes of additive manufacturing using only liquids or viscous materials, e.g. depositing a continuous bead of viscous material using filamentary material being melted, e.g. fused deposition modelling [FDM]
• B33Y 10/00 - Processes of additive manufacturing
• B33Y 80/00 - Products made by additive manufacturing
• B29C 39/10 - Shaping by casting, i.e. introducing the moulding material into a mould or between confining surfaces without significant moulding pressure; Apparatus therefor for making articles of definite length, i.e. discrete articles incorporating preformed parts or layers, e.g. casting around inserts or for coating articles

Abstract

L'invention concerne un méta-matériau acoustiques (100), ainsi qu'un procédé pour sa fabrication. Le méta-matériau acoustique comprend une pluralité de canaux (101) ayant chacun une même section transversale avec un rayon hydraulique entre 5 et 300 pm, disposés avec un espacement périodique (t) entre canaux adjacents entre 2 et 600 pm. On peut ainsi obtenir un réseau hautement dense de micro-canaux acoustiques pouvant offrir une absorption et/ou impédance acoustiques optimales sur une large bande de fréquences. Son procédé de fabrication est un procédé de fabrication additive comprenant plusieurs étapes consécutives de dépôt de matériau pour former, dans chaque étape, une couche (203) comportant une pluralité de cellules (204) périodiquement répétées, séparées par des parois (104). Les couches (203) déposées dans les étapes consécutives de dépôt de matériau sont empilées avec leurs cellules (204) respectives alignées de manière à former les canaux (101).

IPC Classes  ?

• G10K 11/175 - Methods or devices for protecting against, or for damping, noise or other acoustic waves in general using interference effects; Masking sound
• B64C 1/40 - Sound or heat insulation
• B64D 29/00 - Power-plant nacelles, fairings, or cowlings
• G10K 11/162 - Selection of materials

Abstract

Said method for determining aircraft landing runway conditions comprises the steps of: acquiring a set of data groups of different types (D1, D2) for evaluating and monitoring runway degradation conditions; deriving weighting coefficients (Ki) from each data group; filtering the data; determining, for each data group, a partial runway condition; modifying the weighting coefficients of each data group; and combining the partial runway conditions to derive a runway condition coefficient (RWYCC) associated with a confidence index (CI) derived from the modified weighting coefficients.

IPC Classes  ?

• G08G 5/02 - Automatic landing aids, i.e. systems in which flight data of incoming planes are processed to provide landing data

Abstract

Procédé (100) de fabrication d'un matériau poreux multicouches par fabrication additive (104), ledit matériau comprenant au moins deux couches superposées homogènes formées par des filaments, caractérisée en ce que le procédé de fabrication comprend un procédé de détermination de caractéristiques géométriques de chaque couche du matériau qui comprend les étapes suivantes : - détermination (101) des paramètres JCAL d'un modèle de fluide équivalent de différentes couches homogènes formées de filaments en fonction des paramètres microstructuraux des couches ; - détermination (102) du comportement acoustique d'un matériau poreux multicouches comprenant au moins deux couches homogènes superposées formées de filaments à partir des épaisseurs desdites couches et des paramètres JCAL déterminés précédemment ; et - détermination (103) de l'épaisseur et des paramètres microstructuraux des couches du matériau poreux multicouches de manière à ce qu'au moins une grandeur parmi une impédance acoustique, un coefficient d'absorption ou de réflexion du matériau poreux soit comprise dans un intervalle ciblé selon une fréquence acoustique.

IPC Classes  ?

• B29C 64/153 - Processes of additive manufacturing using only solid materials using layers of powder being selectively joined, e.g. by selective laser sintering or melting
• B33Y 50/00 - Data acquisition or data processing for additive manufacturing
• B33Y 80/00 - Products made by additive manufacturing
• B32B 7/022 - Mechanical properties
• B32B 5/08 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by structural features of a layer comprising fibres or filaments the fibres or filaments of a layer being specially arranged or being of different substances
• B32B 37/00 - Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding
• F02C 7/24 - Heat or noise insulation

Abstract

Revêtement de traitement acoustique comprenant un métamatériau (201), caractérisé en ce que le métamatériau comprend une structure solide (210) comprenant au moins un espace libre (211, 212, 213) formant une cavité repliée et au moins un matériau poreux placé dans ledit espace libre de la structure solide.

IPC Classes  ?

• G10K 11/175 - Methods or devices for protecting against, or for damping, noise or other acoustic waves in general using interference effects; Masking sound
• G10K 11/162 - Selection of materials

Abstract

The invention relates to a tool (22) for removing a fan disc from a module of a turbine engine comprising a rotor and a stator, the rotor comprising the fan disc, a drum and a sealing part secured to one another via bolts each comprising a screw and a nut, each screw passing through a retaining member having at least two lugs arranged around a head of the screw, the stator being delimited by a straightener having a flange, the tool (22) comprising: - a frame (23); - a plate (25) comprising pins and two holes, each of the pins being configured to support the head of a screw, each hole being configured to receive an indexing finger; - at least three supports (30) comprising bearing surfaces (31) configured to support the flange of the straightener; - a visual marker (32) configured to angularly orient the module with respect to the tool (22).

IPC Classes  ?

• F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
• 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 deformation; Tools or devices therefor so far as not provided for in other classes for assembling or disassembling parts

Abstract

A discharge grid (22) designed to be mounted inside or at the outlet of a line (21) of a discharge valve (20) of a turbomachine (1) of an aircraft, the discharge grid (22) comprising an upstream face (24) designed to receive a gas flow (F), a downstream face (25) parallel to the upstream face (24) and designed to deliver the gas flow (F) received at the upstream face (24), and orifices (230) passing through the perforated plate (23) from the upstream face (24) to the downstream face (25) and designed to convey the gas flow (F) through the perforated plate (23). The discharge grid (22) comprises, for each orifice (230) of the perforated plate (23), a tubular channel (26) that is coaxial with its associated orifice and projects from the downstream face (25) of the perforated plate (23).

IPC Classes  ?

• 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
• F01D 17/10 - Final actuators
• F02K 3/075 - 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 controlling flow ratio between flows

Abstract

An acoustic treatment grid (22) designed to be mounted inside or at the outlet of a line (21) of a discharge valve (20) of a turbomachine (1) of an aircraft, this being designed to convey a gaseous flow (F), the grid (22) comprising a perforated plate (23) and circular orifices (230) passing through the perforated plate (23) in a first direction (DA), the orifices (230) having a diameter (D) and a geometric center (C). Each orifice (230) is separated from an adjacent orifice by a space whose length (e) is equal to the product of the diameter (D) of said orifice (230) with a spacing coefficient whose value is between 1.1 and 6.

IPC Classes  ?

• F02C 7/24 - Heat or noise insulation
• F01D 17/10 - Final actuators
• 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
• F02K 3/075 - 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 controlling flow ratio between flows
• F04D 29/52 - Casings; Connections for working fluid for axial pumps
• F04D 29/66 - Combating cavitation, whirls, noise, vibration, or the like; Balancing

Abstract

One aspect of the invention relates to a method for controlling a turbojet thrust reverser during an aborted aircraft takeoff, the thrust reverser comprising: - doors that can move between a closed position, an over-closed position and a deployed position; - door actuators configured to move the doors between the closed, over-closed and deployed positions; - a device for locking the doors in the closed position, said device being movable between a locking position and an unlocking position; and - a lock actuator configured to move the locking device between the locking and unlocking positions. The method comprises the following steps: - reducing the engine speed of the turbojet by following a setpoint value below a first engine speed threshold value at which the aerodynamic forces exerted on the doors are equal to the forces developed by the door actuators; - controlling the door actuators in such a way as to bring the doors into the over-closed position; - controlling the lock actuator in such a way as to bring the locking device into the unlocking position; and - when the locking device is in the unlocking position, controlling the door actuators in such a way as to bring the doors into the deployed position; - in which method the setpoint value of the engine speed is above the idle speed.

IPC Classes  ?

• F02K 1/76 - Control or regulation of thrust reversers
• F01D 21/14 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to other specific conditions

Abstract

Abrégé Procédé de fabrication d'une pièce en matériau composite L'invention concerne un procédé de fabrication d'une pièce en matériau composite à matrice organique ayant une porosité réduite. Figure pour l'abrégé : Fig. 5. CA 3070727 2020-01-30

IPC Classes  ?

• C08J 5/24 - Impregnating materials with prepolymers which can be polymerised in situ, e.g. manufacture of prepregs
• C08K 3/013 - Fillers, pigments or reinforcing additives
• C08K 7/04 - Fibres or whiskers inorganic

Abstract

The invention relates to an assembly (1) for a turbine engine extending along an axis, comprising a stator, a low-pressure compressor shaft (4), a low-pressure compressor comprising a rotor comprising a drum (2) rotatably coupled to the shaft (4) of the low-pressure compressor, a fan (21) comprising a disk (23) rotatably coupled to the shaft (4) of the low-pressure compressor, characterised in that the drum (2) comprises a radially internal part (3) rotatably coupled to the shaft (4) of the low-pressure compressor and a radially external part (5) fixed to the radially internal part (3) by means of detachable fixing means (11), first axial retention means (30) capable of axially and detachably retaining the disk (23) of the fan (21) relative to the shaft (4) of the low-pressure compressor, second axial retention means (15) capable of axially and detachably retaining the radially internal part (3) of the drum (3) relative to the shaft (4) of the low-pressure compressor.

IPC Classes  ?

• F01D 5/02 - Blade-carrying members, e.g. rotors
• F01D 5/06 - Blade-carrying members, e.g. rotors - Details thereof, e.g. shafts, shaft connections
• 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 29/054 - Arrangements for joining or assembling shafts
• F04D 29/26 - Rotors specially adapted for elastic fluids

Abstract

A system (36) for controlling an aircraft thrust reversal means comprises a reverse idle control means (38), a first detection means (31) configured to detect, when the reverse idle control is active, a condition for activation of the thrust reversal means, and an actuation means (52) configured to activate the thrust reversal means when the first detection means (31) detects a condition for activation of the thrust reversal means. It further comprises a second detection means (42, 44, 46, 48, 49) configured to detect a condition for activation of the reverse idle control, the control means (38) being configured to activate the reverse idle control when the second detection means (42, 44, 46, 48, 49) detects a condition for activating the reverse idle control.

IPC Classes  ?

• F02C 9/00 - Controlling gas-turbine plants; Controlling fuel supply in air-breathing jet-propulsion plants
• F02K 1/76 - Control or regulation of thrust reversers

Abstract

Disclosed is a compressor (1) of a dual-flow aircraft turboshaft engine extending longitudinally along an axis X, the compressor (1) comprising a low-pressure compressor, a high-pressure compressor (3) and an intermediate casing (2) axially connecting the low-pressure compressor and the high-pressure compressor (3), the high-pressure compressor (3) comprising an inlet guide stator (33) comprising a retaining ring (34), an outer high-pressure compressor casing (35) and a plurality of vanes (38), the compressor (1) comprising an immobilising device (5) comprising at least one first member secured to the intermediate casing (2) and at least one second member secured to the retaining ring (34) and configured to cooperate with the first member in order to prevent tangential movement while allowing axial movement and radial movement of the retaining ring (34) relative to the intermediate casing (2) according to the axis X.

IPC Classes  ?

• F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings

Abstract

The invention relates to a turbomachine moving blade (2) comprising at least one cooling circuit comprising at least one cavity (16; 16a, 16b) extending radially between the foot and the vertex, at least one air intake opening at a radial end of the cavity, a plurality of first discharge slots (18) arranged to open out along the trailing edge between the foot and the vertex, and a plurality of second discharge slots (20) which are separate from the first discharge slots and provided along the trailing edge (14) between the foot and the vertex, the second discharge slots (20) being axially offset upstream from the first discharge slots (18) and each of the first discharge slots being radially offset from each of the second discharge slots, without any overlap between the first and second discharge slots.

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades

Abstract

The invention relates to a turbomachine module (1), comprising: - a rotating housing (7-8) supporting a propeller provided with a plurality of blades (5), - a system for varying the pitch of the propeller blades (5), the system comprising a control means, and a mechanism for varying the pitch of the propeller blades, characterised in that the system is supported by the rotary housing (7-8), in that the control means comprise an annular row of rotary actuators (16), and in that the mechanism for varying the pitch of the blades comprises a synchronisation ring (11) that is driven to rotate by rotary output shafts (17) of the actuators (16), the synchronisation ring (11) being guided in rotation relative to the rotary housing (7-8) by guide means and meshed by a first toothing (13) with pinions (14) of the blades (5).

IPC Classes  ?

• B64C 11/06 - Blade mountings for variable-pitch blades
• B64C 11/38 - Blade pitch-changing mechanisms fluid, e.g. hydraulic
• B64C 11/44 - Blade pitch-changing mechanisms electric
• F01D 7/00 - Rotors with blades adjustable in operation; Control thereof
• F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps

Abstract

A method for acquiring data for detecting damage to a bearing, comprising the following steps: (a) Obtaining (S10) a vibration signal from the bearing over a period of time; (b) determining (S14) spectrograms of said vibration signal at various instants of the period of time; (c) detecting (S22) the peaks on each spectrogram; (d) from among the detected peaks, retaining (S24) the peaks that, in spectrograms corresponding to successive instants and transformed through synchronous resampling at a rotational speed of the bearing, would be present at the same frequency, with a predetermined tolerance; (e) identifying (S26) spectral lines corresponding to the retained peaks; (f) storing (S44) information representative of said spectral lines in a record able to be used by a unit for detecting damage to the bearing.

IPC Classes  ?

• G01H 1/00 - Measuring vibrations in solids by using direct conduction to the detector
• G01M 13/045 - Acoustic or vibration analysis

Abstract

The invention relates to a method for determining a predictive model of a pressure ratio P1/P2 for a predefined type of dual-flow turbine engine (100) comprising primary (VP) and secondary (VS) flows which are separated by a core compartment (110), the pressure P1 corresponding to an air pressure within the core compartment, the pressure P2 corresponding to an air pressure in the secondary flow. Moreover, the method comprises: - a step (E10) of acquiring, on a test turbomachine operating like the predefined type, at respective measurement times, sets of measurements each comprising a measurement P1, a measurement P2, a measurement M of a Mach number, and a measurement PCN12R representative of a low pressure speed, - a step (E20) of obtaining the sets of measurements by means of a processing device, - a step of supervised learning (E30), by the processing device and from the sets of obtained measurements, of a predictive model of the P1/P2 ratio according to the M number and the speed PCN12R which are considered to be explanatory variables.

IPC Classes  ?

• A61K 31/4427 - Non-condensed pyridines; Hydrogenated derivatives thereof containing further heterocyclic ring systems
• A61K 31/454 - Non-condensed piperidines, e.g. piperocaine containing further heterocyclic ring systems containing a five-membered ring with nitrogen as a ring hetero atom, e.g. pimozide, domperidone
• A61K 39/12 - Viral antigens
• A61K 39/29 - Hepatitis virus
• A61K 39/39 - Medicinal preparations containing antigens or antibodies characterised by the immunostimulating additives, e.g. chemical adjuvants
• A61K 45/06 - Mixtures of active ingredients without chemical characterisation, e.g. antiphlogistics and cardiaca
• A61P 31/20 - Antivirals for DNA viruses
• C07D 407/12 - Heterocyclic compounds containing two or more hetero rings, at least one ring having oxygen atoms as the only ring hetero atoms, not provided for by group containing two hetero rings linked by a chain containing hetero atoms as chain links
• C07D 417/12 - Heterocyclic compounds containing two or more hetero rings, at least one ring having nitrogen and sulfur atoms as the only ring hetero atoms, not provided for by group containing two hetero rings linked by a chain containing hetero atoms as chain links

Abstract

The present invention relates to a bypass fan casing (4), characterised in that it comprises a single wall (14) configured to form at least part of an outer skin of a nacelle (17), said wall (14) comprising at least one reinforcement (13, 130) for stiffening the fan casing (4), at least one such reinforcement (13) being made up of a part of the wall (14) forming at least one annular zone (13A, 13B) projecting towards the inside of the nacelle and/or having an extra thickness, said reinforcement (13) being further configured for attaching the fan casing (4) to arms (7) that pass through a secondary flow stream of the turbojet engine.

IPC Classes  ?

• F01D 25/26 - Double casings; Measures against temperature strain in casings

Abstract

The invention concerns a turbine blade, comprising a root, a vane (12) comprising a leading edge and a trailing edge (17) and a pressure-side wall and a suction-side wall, and comprising cooling vents (26, 27) at the trailing edge (17), this vane also comprising: ? first (T1) and second (T2) serpentine circuits (T1); ? each serpentine circuit (T1, T2) comprising several ducts (CA1, CM1, CT1, CA2, CM2, CV2, CT2) extending in the span direction (EV), being connected to each other by angled portions; ? each serpentine circuit (Tl, T2) being supplied with air by its duct (CA1, CA2) that is closest to the leading edge (16); and in which the vents (26, 27) are supplied by the first and by the second serpentine circuit (Tl, T2).

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades

Abstract

The invention concerns a method for regulating the acceleration of a turbomachine, the turbomachine comprising a combustion chamber, a high-pressure turbine situated downstream from the combustion chamber and that rotates a high-pressure shaft, and a low-pressure turbine situated downstream from the high-pressure turbine and that rotates a low-pressure shaft, characterised in that the method comprises the following steps: - Injecting (100) mechanical power onto the high-pressure shaft until the speed of the high-pressure shaft reaches a target value, then - Drawing (200) mechanical power from the high-pressure shaft in order to maintain the speed at the target value.

IPC Classes  ?

• F02K 5/00 - Plants including an engine, other than a gas turbine, driving a compressor or a ducted fan
• F01D 15/10 - Adaptations for driving, or combinations with, electric generators
• F02C 7/26 - Starting; Ignition
• F02C 7/266 - Electric

Abstract

The invention relates to an assembly (10) for supporting and guiding a drive shaft (1) for an aircraft turbine engine, comprising a bearing support (3) having a cylindrical housing (30) for receiving an outer ring (21) of a bearing (2), and a device for rotationally retaining (5) a first portion (210) of the outer ring which is detached from its attachment to the bearing support; and when the above-mentioned detachment occurs: - the free outer peripheries (521) of the first teeth (52) of the first portion (210) are configured to cooperate by wedge effect in the tangential direction with complementary inner surfaces (532) of the housing (30) of the bearing support ( 3), - and/or the free inner peripheries (531) of the second teeth (53) inside the housing (30) are configured to cooperate by wedge effect in the tangential direction with complementary outer surfaces (522) of the outer ring (21).

IPC Classes  ?

• F01D 21/04 - Shutting-down of machines or engines, e.g. in emergency; Regulating, 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 bearings; Supporting or mounting bearings in casings
• F04D 29/053 - Shafts
• F04D 29/057 - Bearings hydrodynamic

Abstract

A system REG for regulating a physical parameter (y) of a real turbomachine system F(p) from a physical setpoint parameter (yc), the regulation system REG comprising a system OPTK for optimising the parameterisation gain K during the regulation, the optimisation system OPTK comprising a stability correction module (2) determining a first gain component K1, a response time correction module (3) determining a second gain component K2, the stability correction module (2) being designed to inhibit the response time correction module (3) when an instability is detected during the regulation of the physical parameter (y), and a determination module (4) configured to determine the parameterisation gain K as a function of the previously determined first gain component K1 and second gain component K2.

IPC Classes  ?

• G05B 11/01 - Automatic controllers electric
• G05B 13/02 - Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric

Abstract

Turbine for a turbomachine, comprising an annular row (20) of movable blades (20a) surrounded by a support ring (26) which is made of an abradable material (24) and borne by a casing (10), and a nozzle (30) which is mounted downstream of the annular row (20) of movable blades (20a) and comprises a means for hooking on a casing (10) support means, the hooking means comprising a radially outer spoiler (38) bearing radially inwardly on a cylindrical wall (12) of the support means, the support means further comprising an annular wall (14) which extends radially inwardly from the cylindrical wall (12) and the radially inner end of which is engaged in an annular groove (28) of the ring (26), and wherein a free annular space (40) is formed between a radially outer face (26b) of the ring (26) and the cylindrical wall (12) of the support means.

IPC Classes  ?

• F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
• F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings

Abstract

The invention relates to a method for estimating a dead zone at closing of a flap (P1) of a discharge valve (VBV1) of a turbomachine, in which the actual position of the actuator (V1) is measured for a setpoint closing signal of the flap (P1), a static angle is determined from the actual position, a closing dead zone is determined, corresponding to the fact that the joint (J1) is compressed so as to not let the primary flow pass through the orifice (01) in the secondary flow, by the fact that the dead zone is equal to a reference dead zone, which has been predetermined on a reference turbomachine operating on the ground, to which the measured static angle has been added and from which a reference static angle, which has been predetermined on the reference turbomachine operating in flight, has been subtracted.

IPC Classes  ?

• 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
• F02C 9/16 - Control of working fluid flow
• 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

Abstract

The invention concerns a turbomachine, the upstream part of the annular combustion chamber (110) of which is attached to the external casing (138), by means of pins, each having a ball-joint coupling. For each pin: - the ball-joint coupling (69) is established between the pin (160) and a part (86) of the external casing (138), and - the pin can also slide radially, or with a radial component, by means of a sliding action (72) established between the pin and a tubular part (70b) of the outer annular ferrule (114) and/or the bottom annular wall (154) of the chamber (118).

IPC Classes  ?

• F23R 3/60 - Support structures; Attaching or mounting means
• F23R 3/50 - Combustion chambers comprising an annular flame tube within an annular casing

Abstract

The present intention relates to a method for resetting the static pressure model (mod_Ps3(PCN25R)), called "Ps3 model", upstream of a combustion chamber in a turbine engine comprising a compressor (3), the Ps3 model being used to arbitrate between two acquisition channels (V10, V20) of the static pressure (Ps3), called "Ps3 pressure", upstream of the combustion chamber, the two acquisition channels (V10, V20) using two sensors (10, 20), the model expressing the pressure Ps3 as a function at least of the speed (PCN25R), called "PCN25R speed", of the compressor (3), and comprising the following steps: E1: measuring a value of the pressure Ps3 using one of the two sensors (10, 20); E2: resetting the Ps3 model using the measurement of the value of Ps3.

IPC Classes  ?

• B64D 33/00 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for
• B64F 5/60 - Testing or inspecting aircraft components or systems

Abstract

Turbomachine (1) comprising a cone (40) positioned at an upstream end and secured to rotate as one with the low-pressure shaft (20), a system for deicing said cone (40) comprising resistive heating elements (52) positioned in the cone (40), an energy transfer system (54), a source of electrical power (56) coupled to the high-pressure shaft (20) and connected to the energy transfer system (54) by a set of switches (58), and a computer (60) configured to define a setpoint for the power for the deicing of the cone (40) as a function of ambient pressure and temperature data and of a regime of operation of the low-pressure shaft (20), and as a function of the electrical power supplied by the source of electrical power (56), and to define a duty cycle for the switching of the set of switches (58) in order to deliver electrical power to the resistive heating elements (52).

IPC Classes  ?

• B64C 11/14 - Spinners
• F02C 7/047 - Heating to prevent icing

Abstract

The invention relates to a mould (1) intended to be used for manufacturing a monocrystalline component by the pouring of metal and epitaxial growth, the mould comprising a cavity (10) in which the component is intended to be formed, and a housing (12) of elliptical cross section in which a monocrystalline seed (2) is placed, the seed having an elliptical cross section defined by a minor axis and a major axis, the housing being in fluidic communication with the cavity via an opening (13) of circular cross section via which molten metal is intended to flow, the monocrystalline seed and the opening being centred on the one same vertical axis (Z), wherein the minor axis and the major axis of the cross section of the seed are oriented according to the secondary crystallographic orientations of the monocrystal that forms the monocrystalline seed. The invention also relates to the corresponding manufacturing method.

IPC Classes  ?

• B22C 9/04 - Use of lost patterns
• B22D 27/04 - Influencing the temperature of the metal, e.g. by heating or cooling the mould
• C30B 11/14 - Single-crystal-growth by normal freezing or freezing under temperature gradient, e.g. Bridgman- Stockbarger method characterised by the seed, e.g. its crystallographic orientation

Abstract

The invention relates to a method for estimating and using a dead zone at closing of a flap (P1) of a discharge valve (VBV1) of a turbomachine, in which the actual position of the actuator (V1) is measured for a setpoint signal for closing the flap (P1), a static angle is determined from the actual position, a closing dead zone is determined from the static angle, corresponding to the fact that the joint (J1) remains compressed and seals the closing of the orifice (O1), the dead zone is recorded and the signal (SC1) for controlling the actuator (V1) is generated by the computer as a function of the dead zone.

IPC Classes  ?

• 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
• 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

Abstract

The invention relates to a method for monitoring an engine control unit having at least two separate paths, each of the paths comprising at least: - means for carrying out a given application task AS, the application task AS consisting in part of a plurality of calculations which are performed consecutively and between which there are periods of latency; - a first component capable of performing the calculations; and - a second component capable of storing data. The application tasks AS of the paths are capable of communicating, and the method comprises the following steps: a) detecting a period of latency; b) performing, during said period of latency, a test of the operational state of at least one of the components; and c) determining a state of the component corresponding to a failure state or a healthy state.

IPC Classes  ?

• G06F 11/16 - Error detection or correction of the data by redundancy in hardware
• G06F 11/22 - Detection or location of defective computer hardware by testing during standby operation or during idle time, e.g. start-up testing

Abstract

Disclosed is a cascade-type thrust reversal device for a turbomachine (1) of an aircraft, comprising a cascade (80) having first partitions (82), second partitions (83) intersecting the first partitions (82), and cavities (84), and a casing (71) having a housing (75) into which the cascade (80) can be inserted in a first direction (DA), the casing (71) and the cascade (80) being able to move in translation relative to one another in the first direction (DA). The casing (71) comprises a perforated wall (72) that is intended to be in contact with an air flow (F) and has orifices (722), and wall strips (724) that do not have orifices and are intended to face the first walls (82) of the cascade (80) when the device (70) is in a first position in which the cascade (80) is disposed entirely in the housing (75).

IPC Classes  ?

• F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
• F02K 1/82 - Jet pipe walls, e.g. liners
• G10K 11/168 - Plural layers of different materials, e.g. sandwiches
• G10K 11/172 - Methods or devices for protecting against, or for damping, noise or other acoustic waves in general using resonance effects

Abstract

A cascade-type thrust reversal device (70) for a turbomachine (1) of an aircraft, comprising a thrust reversal cascade (80) and a casing (71), the cascade (80) having first cavities (84) and the casing (71) comprising an opening (76) defining a housing (75) into which said cascade (80) can be inserted in a first direction (DA), and the casing (71) and said cascade (80) being able to move in translation relative to one another in the first direction (DA) between a first position of the device (70), in which position the cascade (80) is disposed entirely in the housing (75), and a second position of the device (70), in which position said cascade (80) is at least partially outside said housing (75). The casing (71) comprises an acoustic treatment panel (74) having second cavities (744) extending in a second plane parallel to the first plane, each first cavity (84) facing a second cavity (744) when the device (70) is in the first position so as to form an acoustic treatment cell (710).

IPC Classes  ?

• F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
• F02K 1/82 - Jet pipe walls, e.g. liners
• G10K 11/168 - Plural layers of different materials, e.g. sandwiches
• G10K 11/172 - Methods or devices for protecting against, or for damping, noise or other acoustic waves in general using resonance effects

Abstract

A cascade (80) for a thrust reversal device intended to be mounted on a turbomachine of an aircraft, the cascade (80) comprising first partitions (82) extending in a first direction (D_A), second partitions (83) extending in a second direction (D_C) orthogonal to the first direction (D_C), the second partitions (83) extending in a third direction (DR) secant to a plane comprising first and second directions (D_A and D_C) between first and second ends (85 and 86). Each second partition (83) forms, with reference planes (Pr) parallel to the second direction (D_C) and to the third direction (DR), a plurality of different angles (?1 to ?4) at separate positions (H1 to H4) of the height (H) of the second partitions (83) separating the first end (85) from the second end (86) of the second partitions (83), each angle (?1 to ?4) being formed between a reference plane (Pr) and the tangent to the second partition (83) taken at the intersection between said corresponding reference plane (Pr) and the second partition (83) at the corresponding position (H1 to H4) of the height (H) of the second partition (83).

IPC Classes  ?

• F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
• F02K 1/82 - Jet pipe walls, e.g. liners

Abstract

The invention relates to a turbine engine separate flow mixer (2) centred on a longitudinal axis (X-X), comprising an exhaust housing (6), a shroud (4) directly connected to the exhaust housing (6) and intended to mix the flows originating from the turbine engine, said shroud (4) comprising a metal sheet formed by a series of first and second longitudinal strips (4a, 4b) circumferentially distributed around the longitudinal axis of the mixer by circumferentially placing the second strips (4b) on either side of the first strips (4a), the first and second strips being configured to form the shroud (4) at rest and by making grooves in the shroud, the grooves being defined by an alternation of internal lobes and of external lobes.

IPC Classes  ?

• F02K 1/38 - Introducing air inside the jet
• F02K 1/48 - Corrugated nozzles

Abstract

A cascade (80) for a thrust reversal device intended to be mounted on a turbomachine of an aircraft, the cascade (80) comprising first partitions (82) extending in a first direction (DA), second fixed partitions (83) extending in a second direction (DC) orthogonal to the first direction (DA), and a frame (81) within which the first and second partitions (82, 83) extend, the frame (81) comprising at least two fixed walls (810, 815) extending in the first direction (DA), and at least part of each first partition (82) extending between two second partitions (83). At least one first partition (82) is movable in the second direction (DC) between a first position in which said first partition (82) is distant, in the second direction (DC), from said fixed walls (810, 815) so as to form a plurality of resonant cavities (84) with the first partitions (83) and/or the fixed walls (810, 815), and a second position in which said one first partition (82) is in contact with a fixed wall (810, 815) or another first partition (82).

IPC Classes  ?

• F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow

Abstract

The invention relates to a turbine vane comprising a root carrying a blade terminated by a tip in the form of a squealer tip, the blade having an intrados wall and an extrados wall, as well as a leading edge, a trailing edge, and a tip wall delimiting a bottom of the squealer tip, by which the intrados wall is connected to the extrados wall, said blade also comprising: - a serpentine median circuit (28), including a first radial pipe (41) that collects air at the root and is connected by a first bend (46) to a second radial pipe (42) that is connected by a second bend (47) to a third radial pipe (43); - a cavity (36) under the squealer tip running along the extrados wall (21) and extending from a central region of the tip (S) to the trailing edge (17); - a central radial pipe (34) collecting air at the root and extending between at least two of the three pipes (41, 42, 43) of the median circuit (28) and directly supplying the cavity (36) under the squealer tip.

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades
• B22C 9/10 - Cores; Manufacture or installation of cores
• B22C 9/24 - Moulds for peculiarly-shaped castings for hollow articles

Abstract

Counter-rotating turbine (C) for a turbomachine (10) extending around an axis of rotation (X) and comprising: an internal rotor (22) turning about the axis of rotation (X) and comprising at least one internal mobile blade (220) rotatably supported by a first shaft (26); an external rotor (20) turning around the axis of rotation (X) in a direction opposite that of the internal rotor (22) and comprising at least one external mobile blade (200) rotatably supported by a second shaft (24) that is coaxial to the first shaft (26), the first and second shafts (24, 25) extending axially from upstream to downstream of the turbine (C), wherein the first shaft (26) is guided in rotation by a first bearing (62) arranged between the first shaft and an upstream casing (60) of the turbine, and the second shaft (24) is guided in rotation by a second bearing (72) arranged between the second shaft (24) and the upstream casing (60) of the turbine.

IPC Classes  ?

• 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

Abstract

The invention relates to a method for repairing a vane (1) made of composite material, which extends along a so-called axial direction (X), said vane (1) comprising a root (2) and a blade (3), wherein the axial end thereof, opposite the root (2), has an area to be repaired, said blade (3) comprising an intrados surface (4) and an extrados surface (5).

IPC Classes  ?

• B29C 73/02 - Repairing of articles made from plastics or substances in a plastic state, e.g. of articles shaped or produced by using techniques covered by this subclass or subclass using liquid or paste-like material
• B29C 73/24 - Apparatus or accessories not otherwise provided for
• F01D 5/00 - Blades; Blade-carrying members; Heating, heat-insulating, cooling, or antivibration means on the blades or the members

Abstract

An aircraft turbojet engine (1) comprising a fan (11) configured to provide a reverse thrust and a nacelle comprising an air intake (2), the air intake (2) comprising at least one circulation duct (3) in the annular cavity (20) opening, on the one hand, at the air intake lip (23) and, on the other hand, at the inner wall (21) and/or the outer wall (22) so as to promote a reverse thrust phase, the air intake (2) comprising at least one cover member (31, 32) mounted to move between a covered position, in which the cover member (31, 32) closes the circulation duct (3) at the air intake lip (23) and an uncovered position, in which the cover member (31, 32) opens the circulation duct (3) at the air intake lip (23).

IPC Classes  ?

• B64D 29/00 - Power-plant nacelles, fairings, or cowlings
• B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
• F02C 7/042 - Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
• F02K 1/64 - Reversing fan flow

Abstract

Disclosed is a method for using an aircraft turbojet engine (1) comprising an air inlet (2) comprising a plurality of rectifier vanes (3), each rectifier vane (3) being mounted such that it can move between a retracted position to assist the thrust phase and a deployed position (B) in which the rectifier vane (3) protrudes from the inner wall (21) in a radially inward direction in order to rectify the reverse air flow (F-INV) of the inner wall (21) to assist a thrust-reverse phase, in which method at least one rectifier vane (3) is in the retracted position during a turbojet engine thrust phase (1), the method comprising, during a thrust-reverse phase of the turbojet engine (1), a step of moving the rectifier vane (3) to the deployed position.

IPC Classes  ?

• B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
• F02C 7/042 - Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
• F02K 1/66 - Reversing fan flow using reversing fan blades

Abstract

Disclosed is a method for treating an internal defect (10) in a part (1) made of a material, said method involving: a) detecting and locating the internal defect (10) in the part; b) defining, inside the part, at least one target volume (4) which at least partially includes the defect; c) for each target volume (4), simultaneously irradiating the target volume by at least two beams (2) which converge in the target volume and are continuous, whereby a treated area is obtained. The energy applied to the target volume by each beam is less than a threshold energy for sintering the material, and the sum of the energies applied to the target volume by each of the beams is greater than or equal to a transformation threshold energy that corresponds to the threshold energy for sintering or melting the material; the material of the part is partially transparent to said beams.

IPC Classes  ?

• B29C 73/34 - Apparatus or accessories not otherwise provided for for local pressing or local heating for local heating
• B23K 26/53 - Working by transmitting the laser beam through or within the workpiece for modifying or reforming the material inside the workpiece, e.g. for producing break initiation cracks
• B23P 6/04 - Repairing fractures or cracked metal parts or products, e.g. castings
• B29C 35/08 - Heating or curing, e.g. crosslinking or vulcanising by wave energy or particle radiation

Abstract

The present invention relates to a bypass air/fluid heat exchanger (2) for a turbofan engine. According to the invention, this exchanger (2) comprises: - an annular outer shroud (3) with two walls, an inner wall (32) and an outer wall (31), - an annular inner shroud (4) concentric with the outer shroud (3), - a series of OGV guide vanes (5) which connect said outer shroud to said inner shroud, - and a circulation circuit (6) for circulating said fluid, the two shrouds delimiting a bypass air flow path, the fluid circulation circuit (6) is formed in the body of the outer shroud (3) and in the body of at least one of the OGV guide vanes (5), this circulation circuit (6) opening at the two respective ends thereof into an inlet opening (34) and into an outlet opening (35), formed through said outer wall (31) of the outer shroud, and the two shrouds (3, 4), the OGV guide vanes (5) and the circulation circuit (6) of said fluid are integral.

IPC Classes  ?

• F28D 7/00 - Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall
• F01D 9/06 - Fluid supply conduits to nozzles or the like
• F01D 25/26 - Double casings; Measures against temperature strain in casings
• F02C 7/14 - Cooling of plants of fluids in the plant
• F28F 1/26 - Tubular elements or assemblies thereof with means for increasing heat-transfer area, e.g. with fins, with projections, with recesses the means being only outside the tubular element and extending transversely the means being integral with the element

Abstract

A process for using an air input (2) of a turboreactor nacelle (1) of an aircraft, comprising an air input lip (23) which comprises at least one fixed portion and at least one portion (3, 3', 4, 4', 5, 5', 6, 6', 7) which can be moved between a first position, in which the air input lip (23) has an aerodynamic profile so as to guide the internal air flow over the internal wall (21) in order to promote a thrust phase, and a second position (B), in which the portion (3, 3', 4, 4', 5, 5', 6, 6', 7) is displaced in relation to the fixed portion so that the air input lip (23) has a second radial thickness (EB) in the second position (B) which is less than the first radial thickness (EA) in the first position so as to promote a reverse thrust phase.

IPC Classes  ?

• B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
• F02C 7/042 - Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
• F02K 1/66 - Reversing fan flow using reversing fan blades

Abstract

L'invention concerne un ensemble (10) pour turbomachine comprenant une conduite (12) reliée à une extrémité (12a) à des moyens (14) d'alimentation en air pressurisé et à une extrémité (12b) opposée à un équipement (16) de la turbomachine, la conduite (12) comportant une portion coudée (12c) destinée à former en utilisation un point bas de la conduite (12), ladite portion coudée (12c) comportant un trou (18) de purge premier d'eau de la conduite (12), dans lequel il comprend en outre un orifice (20) agencé en aval du trou (18) de purge.

IPC Classes  ?

• F01D 25/32 - Collecting of condensation water; Drainage
• F01D 9/06 - Fluid supply conduits to nozzles or the like

Abstract

The invention relates to a contra-rotating turbine (C) of a turbine engine (10) extending about an axis (X) and comprising an inner rotor configured to rotate about the axis of rotation (X), and comprising an inner drum to which an inner moving blading (22) is attached, an outer rotor configured to rotate about the axis of rotation (X) in a direction opposite to the inner rotor, and comprising an outer drum (50) to which an outer moving blading (20) is attached, the outer moving blading (20) comprising at least one attachment rod (212) extending through an opening (51) of the outer drum (50), the outer moving blading (20) being attached to the outer drum (50) via a clamping means (100) attached to the attachment rod (212) from an outer face of the outer drum (50), a set ring (80) being disposed around the attachment rod (212) in the opening (51) of the outer drum (50).

IPC Classes  ?

• F01D 1/26 - Non-positive-displacement machines or engines, e.g. steam turbines characterised by counter-rotating rotors subjected to same working-fluid stream without intermediate stator blades or the like traversed by the working-fluid substantially axially
• F01D 5/03 - Annular blade-carrying members having blades on the inner periphery of the annulus and extending inwardly radially, i.e. inverted rotors
• F01D 5/26 - Antivibration means not restricted to blade form or construction or to blade-to-blade connections
• F01D 5/30 - Fixing blades to rotors; Blade roots
• 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

Abstract

The present invention concerns a method for monitoring an actuating system of a movable component, in particular a movable component of a turbomachine such as a nozzle or a blade, the actuating system comprising a control device configured to deliver a position instruction to a first cylinder and a second cylinder, each cylinder being configured to deliver a position feedback measurement in response to the position instruction, the method being implemented in a monitoring system and comprising, - a first monitoring mode in which the deviations between the position feedback measurements of the two cylinders are detected; - a second monitoring mode in which the deviations between the position feedback measurements of the two cylinders are not detected; method in which the second mode is selected when at least one of the two position feedback measurements is in a transient phase.

IPC Classes  ?

• F01D 17/10 - Final actuators
• F02K 1/15 - Control or regulation
• F02K 1/76 - Control or regulation of thrust reversers

Abstract

Aircraft turbomachine blade (102), said blade comprising at least one inner cavity (100) for circulating a ventilation air flow, said cavity comprising a wall (104) which comprises first projecting elements (106) oriented in a first direction and forming air flow disrupters, said wall further comprising at least a second projecting element (108) oriented in a second direction different from the first direction, said second element and at least one of the first elements overlapping each other in one area, the blade being characterised in that the or each first element which overlaps the second element has a height (H2, H4') which is greater than that (H1) of the second element in the area and greater than that (H3) of the other first elements of the wall, in order to retain its disruptive function along the entire length thereof.

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades

Abstract

Method for controlling a first, a second and a third variable of a turbomachine as a function of a first, a second and a third turbomachine control quantity, each of which can become saturated as a function of the turbomachine operating parameters. The method comprises a first multi-variable correction (120) delivering a first value for the three control quantities, a selection (130) of the first control quantity to be delivered as a function of a minimum value, of a maximum value and of the value determined by the first correction, a second multi-variable correction (140) delivering a second value for the second and third control quantities, and a selection (150) of the values of the second and third control quantities to be delivered according to the values determined during the first correction and those determined during the second correction.

IPC Classes  ?

• F02C 9/48 - Control of fuel supply conjointly with another control of the plant
• B64C 11/48 - Units of two or more coaxial propellers
• F02K 3/02 - 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

Abstract

The invention relates to a tool for preforming a fibrous preform, comprising: - an inflatable first membrane (11) intended to accept the fibrous preform, - a second membrane (18) intended to attach to the first membrane (11) via a fixing system (20) in such a way as to form a fluidtight internal cavity (19) between the first and second membranes, and - an evacuation device (25) for creating a vacuum in the internal cavity between the first membrane (11) and the second membrane (18).

IPC Classes  ?

• B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
• B29C 70/34 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression
• 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

Abstract

A weaving installation (400) comprises a loom (100 intended to produce a woven texture by weaving together a plurality of threads, at least some of the threads of the plurality of threads being carbon threads (210, 211, 212, 213, 214, 215), the carbon threads each being stored individually on one package of a plurality of carbon thread storage packages (220, 221, 222, 223, 224, 225) present upstream of the loom. The installation also comprises a plurality of pairs of first and second electrical contacts (301, 302; 303, 304; 305, 306; 307, 308; 309, 310; 311, 312) present between the storage packages (220, 221, 222, 223, 224, 225) and the loom (100). Each pair of first and second electrical contacts is present on the path of a carbon thread, the first and second contacts of each pair being intended to be in electrical contact with a given carbon thread. The contacts of each pair of first and second contacts are also connected to an open-circuit detection circuit (230).

IPC Classes  ?

• D03D 51/28 - Warp stop motions electrical

Abstract

The invention relates to a silicon heating mat (10), said mat (10) being formable and deformable and comprising a matrix (14) made of elastic material, in which at least one cavity (16) is arranged that fully passes through the matrix (14), said at least one cavity (16) being intended to accommodate a resistive filament (18) connected to a heating cycles management unit (12). Furthermore, said at least one cavity (16) has an undulating layout, with each resistive filament (18) being able to move inside said at least one cavity (16), and each resistive filament (18) having a zigzag or spiral shape.

IPC Classes  ?

• H05B 3/34 - Heating elements having extended surface area substantially in a two-dimensional plane, e.g. plate-heater flexible, e.g. heating nets or webs

Abstract

The invention relates to a turbine engine vane (20) comprising a blade (21) extending along a radial axis and a first cooling circuit (28) arranged inside the blade, the first cooling circuit (28) comprising a first cavity (34) and a second cavity (35) disposed downstream of the first cavity in a direction of circulation of a coolant in the blade, the first and second cavities radially extending inside the blade and being at least partly separated by a first radial partition (36) having a radially internal free end (37), which at least partly demarcates a first coolant passage (40) connecting the first and second cavities. According to the invention, the radially internal free end (37) is enlarged by having a general transverse section substantially in the form of a keyhole.

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades

Abstract

The invention relates to a turbine engine blade (20) comprising: - an airfoil (21) with a pressure-side wall and a suction-side wall which are connected upstream by a leading edge (26) and downstream by a trailing edge (27), - a cooling circuit (28) which comprises an internal cavity extending inside the airfoil and a plurality of outlet openings each oriented substantially along a longitudinal axis X, each outlet opening communicating with the cavity and being arranged in the vicinity of the trailing edge, and - a calibration device (33) arranged in the cavity and provided with calibration conduits (34) which are arranged substantially opposite the outlet openings. According to the invention, the calibration conduits (34) each comprise an oblong transverse section which is substantially perpendicular to the longitudinal axis.

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades

Abstract

Disclosed is a method for controlling the braking of the wheels of an aircraft, in which method the braking of the wheels of the aircraft is controlled by a wheel braking controller (14) actuating wheel brakes (10) of the aircraft as a function of both a deceleration control request (16) and a request for deployment of the thrust reversers (24).

IPC Classes  ?

• B64C 25/42 - Arrangement or adaptation of brakes
• B64C 25/34 - Alighting gear characterised by elements which contact the ground or similar surface wheeled type, e.g. multi-wheeled bogies

Abstract

The invention relates to a tubular ventilation sleeve (1) for a turbomachine distributor, in particular for an aircraft, the sleeve having a generally elongate shape along an axis (A-A) and comprising a perforated tubular wall (1a) around said axis, one of the axial ends of the sleeve being open and the other being closed by a bottom wall (1b), characterized in that it further comprises support beams (7) when the sleeve (1) is made by additive manufacturing, said beams (7) extending inside the sleeve (1) between the tubular wall (1a) and the bottom wall (1b) and having a longitudinal cross-section with a generally triangular shape, two sides of which are respectively connected to the tubular wall (1a) and the bottom wall (1b) and the last side of which is free and extends inside the sleeve, perforations (2) in the tubular wall (1a) being provided between the support beams (7).

IPC Classes  ?

• F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades

Abstract

The invention relates to a method for regulating a temperature of the exhaust gases of a turbomachine (1), the method comprising the following steps: - regulating the injection of fuel into a combustion chamber (5) of the turbomachine (1) such that the turbomachine (1) generates a target thrust; - regulating the injection of mechanical power by an electric motor (10) to a shaft (8, 9) driven to rotate by a turbine (6, 7), the electric motor (10) being activated when a clearance between a casing (62) and blades (61) of the turbine (6, 7) exceeds a threshold value.

IPC Classes  ?

• F02C 7/275 - Mechanical drives
• F01D 15/10 - Adaptations for driving, or combinations with, electric generators
• F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
• F02C 9/48 - Control of fuel supply conjointly with another control of the plant

Abstract

The invention relates to an assembly for a turbomachine extending along an axis (X) and comprising: - a ferrule (32) designed to define a fan duct (5) of a gas stream of the turbomachine, - a fan casing (2) radially surrounding the ferrule (32) and defining with the ferrule (32) the fan duct (5), - a rectifier (6) comprising a plurality of vanes (7) comprising a first vane (7a) and a second vane (7b) adjacent to the first vane (7a), the vanes defining between them a converging flow channel (13) designed to direct and accelerate the stream by means of an inlet section (14a) included in a plane non-perpendicular to the axis of the turbomachine and an outlet section included in a plane (14b) perpendicular to the axis (X) of the turbomachine, the first vane (7a) and the second vane (7b) each having an unducted downstream portion which forms a trailing edge.

IPC Classes  ?

• 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

Abstract

The invention relates to a method for detecting a possible fuel leak in an oil circuit of an aircraft engine, said aircraft having at least one pair of identical engines equipped with respective oil circuits, said pair of engines being associated with at least one quadruplet of measurements acquired beforehand at a measuring time during operation of the engines of the pair, said measurements corresponding to a pressure measurement and a temperature measurement of the fluid contained in each of the oil circuits of the engines of the pair. Moreover, the method includes: - a step (100) of determining a quantity Q representative of a possible difference in operation between the engines of the pair, depending on the quadruplet of measurements, - a step (200) of comparing the quantity Q with a threshold value determined beforehand, so as to obtain a comparison result, - a step (300) of detecting a possible leak of fuel in the oil circuit of one of the engines of the pair, depending on the comparison result.

IPC Classes  ?

• F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
• F01D 25/18 - Lubricating arrangements

Abstract

The invention concerns a bypass turbomachine (1) with a primary flow path and a secondary flow path, comprising: - a low-pressure body comprising a low-pressure compressor (120) connected to a low-pressure turbine (122) via a low-pressure shaft (124), - a high-pressure body comprising a high-pressure compressor (130) connected to a high pressure turbine (132), via a high-pressure shaft (134), - a low-pressure power take-off system (220) comprising an electrical generator (226), configured to take power (W12) from the low-pressure body, wherein - the turbomachine comprises a debris removal system (500), located between the two compressors (226, 236), - the low-pressure power take-off system (220) is configured to take power (W12) from the low-pressure shaft (124) using the resistive torque of the electrical generator (226), in order to avoid a risk of surging.

IPC Classes  ?

• F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
• 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

Abstract

Method for repairing a blade (10) of a turbomachine rotor, particularly a turbomachine fan, this blade comprising a vane (12) made from an organic-matrix composite and a metallic shield (14) bonded to a leading edge of the vane, the metallic shield (14) having a damaged region (20), the method comprising a step of resetting the leading edge of the vane, a step of cooling the entire vane, followed by a step of refilling the damaged region (20) by adding a metallic material to this damaged region (20) and melting the material using a laser, and a step of tailoring the damaged region (20) by machining the molten material.

IPC Classes  ?

• F01D 5/14 - Form or construction
• F01D 5/00 - Blades; Blade-carrying members; Heating, heat-insulating, cooling, or antivibration means on the blades or the members
• F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion

Abstract

Disclosed is a method for repairing a composite material part (10), comprising: - producing a hollowed-out portion (30) by removing the composite material in a damaged area (20), - producing grooves (110, 120) extending on the two faces (11, 12) of the part (10) from the edges (32, 34) of the hollowed-out portion, - producing a plurality of orifices (130) in the thickness of the part, opening into the grooves, - three-dimensional weaving of a fibrous preform of a filling part (40, 50) comprising a central portion (41, 51) having a shape corresponding to the shape of the hollowed-out portion and a plurality of fibre bundles (420, 450) extending from the central portion (41, 51), - placing the filling preforms (40, 50) respectively on one or other face of the part (10) with the fibre bundles present in the grooves present on the faces of the part and in the orifices of the plurality of orifices (130) so that the fibre bundles open into the grooves present on the opposite face of the part, - impregnating the fibrous filling preforms (40, 50) with a resin precursor of a matrix, - transforming the resin into a matrix in order to obtain a filling part made of composite material (60) comprising a fibrous reinforcement densified by a matrix.

IPC Classes  ?

• B29C 73/14 - Repairing of articles made from plastics or substances in a plastic state, e.g. of articles shaped or produced by using techniques covered by this subclass or subclass using preformed elements using elements composed of two parts joined together after having been placed one on each side of the article
• B64F 5/40 - Maintaining or repairing aircraft
• B29C 73/04 - Repairing of articles made from plastics or substances in a plastic state, e.g. of articles shaped or produced by using techniques covered by this subclass or subclass using preformed elements
• 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 73/26 - Apparatus or accessories not otherwise provided for for mechanical pretreatment

Abstract

The invention relates to a drive shaft (50) of an aircraft turbine engine, comprising: - a first portion (52) and a second portion (54), - connecting means (56) connecting said first and second portions and being configured to transmit a torque from said second portion to said first portion, said connecting means comprising at least one bellows (70) comprising: - a first section (72) having a diameter greater than the diameters of said first and second portions, - second sections (74) flanking said first section, said drive shaft being characterised in that said first section includes at least one fusible section (76) comprising at least one through-hole (78) and being configured to break when the value of a torque applied to said first portion exceeds a predetermined threshold value.

IPC Classes  ?

• F16D 9/08 - Couplings with safety member for disconnecting by breaking due to shear stress over a single area encircling the axis of rotation, e.g. shear necks on shafts
• F01D 21/04 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
• F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
• F16C 3/00 - Shafts; Axles; Cranks; Eccentrics
• F16D 3/12 - Yielding couplings, i.e. with means permitting movement between the connected parts during the drive adapted to specific functions specially adapted for accumulation of energy to absorb shocks or vibration

Abstract

Method for controlling a fuel-metering device with a mobile metering element, comprising at least two iterations of the following steps: - detecting (E1) a potential change in operational status of one of two metering-element position sensors, if no change in operational status is detected, determining (E2_1) the position of the metering element from a mean of the measurements from the sensors, or else determining (E2_2) said position from the sensor that is not defective, - determining (E4) a fuel flowrate setpoint, converting (E5) the fuel flowrate setpoint, - determining (E6) a metering-element movement command, commanding (E7) the position of the metering element and, if a change in operational status is detected, calculating an instantaneous fuel flowrate from the position of the metering element and, during the second iteration of the method, determining the flowrate setpoint as a function of the instantaneous flowrate in order to make the position setpoint correspond to the position of the metering element.

IPC Classes  ?

• F02C 9/32 - Control of fuel supply characterised by throttling of fuel
• F02C 9/26 - Control of fuel supply

Abstract

The invention relates to a combustion chamber comprising an end wall (4) comprising at least one opening (5), at least one sleeve (12) mounted upstream of the end wall (4) and fastened to the end wall (4), a closure ring (13) delimiting an annular groove (11) with the sleeve (12) and fastened to the sleeve (12), at least one air- and fuel-injection system (6), having an axis (A), mounted in the opening (5) of the end wall (4), the injection system (6) comprising an annular rim (10) extending radially with respect to said axis (A) and being mounted in said groove (11) with a radial clearance, and a deflector (14) situated downstream of the end wall (4) and fastened to the sleeve (12) and/or to the end wall (4).

IPC Classes  ?

• F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
• F16B 5/10 - Joining sheets or plates to one another or to strips or bars parallel to them by means of bayonet connections
• F23R 3/50 - Combustion chambers comprising an annular flame tube within an annular casing
• F23R 3/60 - Support structures; Attaching or mounting means

Abstract

The invention relates to a blade pivot (2) of adjustable orientation for a turbomachine fan hub, comprising a block (4) having retaining means (10) configured to retain a fan blade root (12) and coupling means (16) for the transmission of a torque, a ball-type rolling bearing (26) for taking up centrifugal forces having an inner ring (28), a clamping nut (42) screwed onto an external thread (44) of the block in order to clamp the inner ring of the ball-type rolling bearing for taking up centrifugal forces onto the block, a rolling bearing (34) for taking up transverse forces, a wedging transmission ring (18) positioned inside the inner radial end of the block and provided with coupling means (20) cooperating with the coupling means (16) of the block, and means (22, 24) for locking the wedging transmission ring on the block.

IPC Classes  ?

• B64C 11/06 - Blade mountings for variable-pitch blades
• B64C 11/32 - Blade pitch-changing mechanisms mechanical
• F01D 7/02 - Rotors with blades adjustable in operation; Control thereof having adjustment responsive to speed
• F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps

Abstract

The invention relates to a turbomachine blade (54) comprising a body (60) that extends mainly in a plane defined by a main axis B and a longitudinal direction, which is defined by a lower surface wall (66), an upper surface wall (68), a leading edge (70) located at a first longitudinal end of the body (60) and a trailing edge (72) located at a second longitudinal end of the body (60), characterized in that the body (60) of the blade (54) comprises a plurality of first pipes (74) that extend mainly along the direction of the main axis B, for circulation of a gas flow, and a plurality of second pipes (76) that extend mainly along the longitudinal direction, for circulation of a second gas flow.

IPC Classes  ?

• 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
• F01D 5/14 - Form or construction
• F01D 9/06 - Fluid supply conduits to nozzles or the like
• F01D 25/12 - Cooling
• F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
• 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
• F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange

Abstract

A method for electrical discharge drilling of a hole (18) in a part (16) made of electroconductive material, in particular for an aircraft turbomachine, the method using an electrical discharge machine comprising a head (12) that can move relative to the part and has a consumable EDM electrode (10), which is of elongated shape and is translationally moved along its elongation axis, the method comprising the following steps: a) advancing the EDM electrode towards the part in order to drill a hole in the part; and b) retracting the EDM electrode and removing the EDM electrode from the hole, characterised in that it further comprises the following steps: c) laterally moving the head; d) advancing the EDM electrode towards the part in order to scan for the part; e) computing the effective depth of the hole drilled in step a).

IPC Classes  ?

• B23H 9/10 - Working turbine blades or nozzles
• B23H 9/14 - Making holes
• B23H 1/00 - Electrical discharge machining, i.e. removing metal with a series of rapidly recurring electrical discharges between an electrode and a workpiece in the presence of a fluid dielectric

Abstract

The invention relates to a reducing assembly (26) comprising a planet carrier (27) that is flexible and fixed, borne by a casing (32) which surrounds it by means of two distinct connections with this casing (32), with means (44) for applying a torsional prestress to the planet carrier (27) between these two connections. Prestressing the flexible planet carrier (27) makes it possible to limit its level of mechanical stress when it is in service, in order to simplify its design and dimensioning.

IPC Classes  ?

• F16H 57/08 - General details of gearing of gearings with members having orbital motion

Abstract

The invention relates to a method for monitoring the state of health of at least two vibration sensors of a twin-spool turbomachine comprising a low-pressure spool and a high-pressure spool, one vibration sensor being located at the front of the turbomachine, and another vibration sensor being located at the rear of the turbomachine, each of the sensors being configured to measure the vibrations of the low-pressure spool and high-pressure spool at the front and at the rear of the turbomachine, the method being carried out in a processing unit (20) of the turbomachine, which processing unit is in communication with each of the sensors, the method comprising the following steps: - receiving the low-pressure speed (NBP) and high-pressure speed (NHP) of the turbomachine and, when the speeds are simultaneously in predetermined ranges, - receiving the front and rear vibration levels of the low-pressure and high-pressure spools which are recorded by each sensor; - establishing the mean of the values of the vibration levels of the low-pressure and high-pressure spools received over a predetermined receiving time; - establishing the state of health of the at least first and second vibration sensors from a comparison between the established mean values of the vibration levels of the low-pressure and high-pressure spools and predetermined thresholds.

IPC Classes  ?

• F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
• F01D 17/02 - Arrangement of sensing elements
• F01D 25/04 - Antivibration arrangements
• G01H 1/00 - Measuring vibrations in solids by using direct conduction to the detector
• G01H 3/00 - Measuring vibrations by using a detector in a fluid

Abstract

The invention relates to a method for separating a first mechanical part (12) from a second mechanical part (14), wherein the second mechanical part (14) is bonded to the first mechanical part (12) by an adhesive film (16) along a connecting area, the first mechanical part (12) having a first specific thermal conductivity and the second mechanical part (14) having a second thermal conductivity that is higher than the first thermal conductivity, characterised in that it comprises at least one cooling step during which the second mechanical part (14) is cooled to a negative temperature and at least one stressing step during which the second mechanical part (14) is subjected to mechanical stress in order to cause the adhesive film (16) to break.

IPC Classes  ?

• 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
• B26F 3/06 - Severing by using heat
• B32B 43/00 - Operations specially adapted for layered products and not otherwise provided for, e.g. repairing; Apparatus therefor
• C09J 5/06 - Adhesive processes in general; Adhesive processes not provided for elsewhere, e.g. relating to primers involving heating of the applied adhesive

Abstract

The invention relates to an aerodynamic arm (100) for an aircraft turbine engine casing, characterised in that it comprises: - a tubular outer shell (110) having a generally elongate shape extending substantially along an axis (A-A), the shell (110) comprising axial ends (112) for connecting to a turbine engine casing; - an electrically conductive core (120) extending inside the shell (110) and having ends for electrically connecting to each of the ends (112) of the shell (110); and - an insulating material (130) configured to occupy a space provided between the core (120) and the shell (110).

IPC Classes  ?

• F04D 29/54 - Fluid-guiding means, e.g. diffusers
• F01D 5/14 - Form or construction
• F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
• F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
• F01D 9/06 - Fluid supply conduits to nozzles or the like
• F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
• F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep
• 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 29/02 - Selection of particular materials

Abstract

An injector nozzle (43) for a turbomachine comprises a primary fuel circuit ending in a fuel ejection pipe (66), and a secondary fuel circuit comprising an annular end portion for ejecting fuel (68) arranged around the fuel ejection pipe. A portion upstream from the primary fuel circuit comprises an annular channel (70), which extends around the secondary fuel circuit and is defined by an external wall (72) of the injector nozzle. The injector nozzle comprises air intake channels (126) extending through the annular channel (70) and having inlets opening into the external wall (72) and outlets (130) opening into an annular air injection channel (124) arranged radially in the interior in relation to the end portion for ejecting fuel, around the fuel ejection pipe, and cooperating with the end portion for ejecting fuel in order to form an aerodynamic secondary injector.

IPC Classes  ?

• F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
• F23R 3/34 - Feeding into different combustion zones

Abstract

The invention relates to a system and method for controlling an aircraft turbine engine. According to the invention, the control system (100) comprises: - a nominal-mode processing chain (110) comprising a global corrector (21) designed to control a speed of rotation of the turbine engine by delivering a position setpoint (C_WF) for a fuel metering device (11, 12), and a local corrector (23) designed to control a position of the fuel metering device by delivering a nominal-mode control current (I_nom), - a degraded-mode processing chain (120) comprising a direct corrector (122) designed to control the speed of rotation of the turbine engine by delivering a degraded-mode control current (I_dég), and - a mode management module (130) designed to deliver, to the fuel metering device, the nominal-mode control current (I_nom) in the absence of failure of a position sensor measuring a position of the fuel metering device, and the degraded-mode control current (I_dég) in the case of failure of the position sensor.

IPC Classes  ?

• F02C 9/32 - Control of fuel supply characterised by throttling of fuel
• F02C 9/46 - Emergency fuel control

Abstract

Tooling (3) for placing a propulsive assembly (2) from a horizontal position to a vertical position, comprising: - a front tool (4) comprising: a first assembly comprising a frame (19) configured to be positioned and maintained with respect to a front casing (14), said frame (19) comprising a first connection interface (20) configured to be connected to a first lifting system (8) such that said frame (19) is movable about an axis of rotation (R); a second assembly comprising a shaft configured to be positioned and maintained with respect to a rotor; a member for connection between said first assembly and said second assembly; - a rear tool (6) configured to be positioned and maintained with respect to said rear casing (7), said rear tool (6) comprising a second connection interface (23) configured to be connected to a second lifting system (10).

IPC Classes  ?

• B64F 5/50 - Handling or transporting aircraft components
• F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
• F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep

Abstract

The invention relates to a propulsion system (100) for an aircraft, comprising at least one rotor (110) and a nacelle cowling (120) extending around the at least one rotor (110), said nacelle cowling (120) being sectored and comprising at least one sector (130a, 130b) which is fixed and sectors (141a, 141b, 142a, 142b) which are retractable in the peripheral direction (F1, F2) relative to an axis of rotation (X) of the rotor (110), characterised in that the retractable sectors (141a, 141b, 142a, 142b) comprise at least a first series of sectors (141a, 142a) which are telescopically retractable in or on at least one fixed sector (130a) and at least a second series of sectors (141b, 142b) which are telescopically retractable in or on at least one fixed sector (130b), the at least one fixed sector (130a, 130b) having an angular extent around the axis (X) which is less than or equal to 90°.

IPC Classes  ?

• B64C 11/00 - Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
• B64D 29/06 - Attaching of nacelles, fairings, or cowlings

Abstract

The invention relates to a fire resistance device (50) that is designed to be placed between a mounting strut of a double-flow aircraft turbomachine and a connecting cowling with which said turbomachine is equipped, the connecting cowling being designed to connect an upstream ring, delimiting an inter-flow compartment, to an arm that extends radially across a secondary flow of said turbomachine. According to the invention, the device (50) comprises two contacting lips (52a, 52b) which extend along different lines (72a, 72b), a first lip (52a) with a C-shaped section being integrated inside a contact structure (73), and a second lip (52b) having a protruding blocking portion (63) at the end thereof, protecting said contact structure (73).

IPC Classes  ?

• F02C 7/25 - Fire protection or prevention
• B64D 29/06 - Attaching of nacelles, fairings, or cowlings
• F01D 9/06 - Fluid supply conduits to nozzles or the like
• F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
• F02C 7/28 - Arrangement of seals

Abstract

The invention relates to a turbomachine rotor blade which is characterized in that: - the ratio between the maximum thickness and the chord at 30% of the height of the blade is between 20% and 42% of the ratio between the maximum thickness and the chord at the blade root, - the ratio between the maximum thickness and the chord at 70% of the height of the blade is between 10% and 30% of the ratio between the maximum thickness and the chord at the blade root, - the ratio between the maximum thickness and the chord at 90% of the height of the blade is between 10% and 30% of the ratio between the maximum thickness and the chord at the blade root, - the ratio between the maximum thickness and the chord at the blade head is between 3% and 21% of the ratio between the maximum thickness and the chord at the blade root.

IPC Classes  ?

• F01D 5/14 - Form or construction

Abstract

The invention relates to a turbomachine rotor blade comprising a plurality of blade sections stacked along an axis between a blade root and a blade head defining between them the height of the blade, each blade section comprising a chord and a maximum sweep defined by the maximum length of a segment perpendicular to the chord line and connecting a point of the chord line and a point of a camber line formed of all points located equidistant from the extrados and the intrados in the section, and verifying that the ratio between the maximum sweep and the chord at mid-height of the blade and the same blade head ratio is between 25% and 40% of the ratio between a maximum sweep and a blade root chord.

IPC Classes  ?

• F01D 5/14 - Form or construction

Abstract

The invention relates to a fan module (3) comprising variable-pitch blades, said fan module (3) including: a fan rotor (20) bearing the blades (23) of the fan (4, 400a, 400b), each mounted to pivot about a pitch axis (A); a fan shaft (21) extending along a longitudinal axis X inside the fan rotor and driving the fan rotor in rotation; a power shaft (11) driving the fan shaft in rotation by means of a planetary gear speed reducer (50); at least a first and a second bearing (69, 71) for guiding the rotation of the fan rotor, housed inside a lubrication chamber (60); and a system (32) for varying the pitch of the fan blades, comprising a connecting mechanism (33) connected to the blades of the fan and a control means (34) acting on the connecting mechanism (33). According to the invention, the first bearing (69) is disposed upstream of the speed reducer (50) and the second bearing (71) is disposed downstream of the speed reducer (50), the fan rotor (20) being connected to the fan shaft (21) by means of an annular trunnion (59) extending at least along an upstream portion of the lubrication chamber (60).

IPC Classes  ?

• F04D 25/02 - Units comprising pumps and their driving means
• B64C 11/38 - Blade pitch-changing mechanisms fluid, e.g. hydraulic
• F01D 5/12 - Blades
• F01D 7/00 - Rotors with blades adjustable in operation; Control thereof
• F01D 15/12 - Combinations with mechanical gearing
• F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
• F01D 25/18 - Lubricating arrangements
• 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 3/068 - 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 being characterised by a short axial length relative to diameter
• F04D 29/054 - Arrangements for joining or assembling shafts
• F04D 29/056 - Bearings
• F04D 29/063 - Lubrication specially adapted for elastic fluid pumps
• F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps

Abstract

The invention relates to an engine assembly (200) for an aircraft comprising a bypass turbomachine as well as a turbomachine attachment pylon (202) comprising an air-oil exchanger system (30) arranged in an inter-ducts compartment (22) between the flow ducts, the compartment being delimited radially on the outside by an inter-ducts cowling (20), the exchanger system being supplied with air from a secondary flow duct (14b) of the turbomachine delimited radially on the inside by the inter-ducts cowling (20), and the exchanger system (30) being supported by a support (34) arranged in the inter-ducts compartment (22), this support (34) being mechanically connected to the attachment pylon by connecting means (30) passing through the inter-ducts cowling (20).

IPC Classes  ?

• F01M 5/00 - Heating, cooling, or controlling temperature of lubricant; Lubrication means facilitating engine starting
• F01D 25/18 - Lubricating arrangements
• F02C 7/06 - Arrangement of bearings; Lubricating
• F02C 7/14 - Cooling of plants of fluids in the plant
• F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
• F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange
• F16N 39/02 - Arrangements for conditioning of lubricants in the lubricating system by cooling

Abstract

The invention concerns an air-sealing device (40) intended to be inserted between an aircraft dual-flow turbine engine casing element (16) and a nacelle element (30), the sealing device comprising an attachment tab (42) at the end of which is located a sealing portion (44) having an outer surface intended to be contacted by the casing element and the nacelle element, and an inner surface defining a cavity. The inner surface defines at least one protuberance extending inside the cavity.

IPC Classes  ?

• F02C 7/28 - Arrangement of seals
• F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
• F16J 15/02 - Sealings between relatively-stationary surfaces

Abstract

The method for detecting a rotating stall comprises; - a step (E90, E100) of determining a level of variation of a static pressure in a combustion chamber of the turbojet engine around an average value of this static pressure; - a step (E11) of comparing the level of variation of the static pressure relative to a first threshold (THR.1); - a step (E140) of comparing a temperature measured at the outlet of a turbine of the turbojet engine relative to a second threshold; and - if the level of variation of the static pressure is greater than the first threshold and the temperature at the outlet of the turbine is greater than the second threshold, a step (E170) of detecting the presence of a rotating stall.

IPC Classes  ?

• F04D 27/00 - Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
• F02C 9/00 - Controlling gas-turbine plants; Controlling fuel supply in air-breathing jet-propulsion plants

Abstract

The invention relates to a method for control of a cylinder (12), comprising the following steps: providing a cylinder that has a piston (22), a servo valve (14), and a measuring device (16) that comprises at least one first position sensor (28) and one second position sensor (30); taking position measurements (X1, X2) of the piston simultaneously with the first position sensor and with the second position sensor; determining at least one first travel speed (v1) of the piston on the basis of the position measurements of the piston taken with the first position sensor; determining at least one second travel speed (v2) of the piston on the basis of the position measurements of the piston taken with the second position sensor; and comparing each of the first and second determined travel speeds (v1, v2) of the piston with a modeled (vmod) or a predetermined travel speed of the piston so as to ascertain which is the more reliable position sensor.

IPC Classes  ?

• F15B 19/00 - Testing fluid-pressure actuator systems or apparatus, so far as not provided for elsewhere

Abstract

The invention relates to a computing environment system for monitoring aircraft engines, said system being connected to a cluster of servers, said system comprising: - an application interface (13) configured so as to receive user codes (19) that are developed independently from a distributed deployment system specifying the calculation of a set of indicators relating to an aircraft engine (3) for deployment on a fleet (12) of aircraft engines; - an extraction module (15) configured so as to extract said indicators by deploying parallel calculations on temporal flight data from the fleet of aircraft engines and stored in a database (9) distributed over said cluster (5) of servers (7); - a learning module (17) configured so as to use said indicators to construct, without supervision, from said indicators, a monitoring model (21) representative of the indicators by implementing predetermined learning functions.

IPC Classes  ?

• G06F 8/60 - Software deployment
• G06N 20/00 - Machine learning
• G05B 23/02 - Electric testing or monitoring
• B64D 45/00 - Aircraft indicators or protectors not otherwise provided for

Abstract

The invention relates to a turbojet engine comprising a central shaft (AC) surrounded by a coaxial and independent high-pressure body (CH), said turbojet engine comprising, from upstream (AM) to downstream (AV): - a fan (13) driven by the central shaft (AC) ;- a high-pressure compressor (16) and a high-pressure turbine (17) supported by the high-pressure body (CH); - an inter-turbine housing (18); - a low-pressure turbine (19); and - an exhaust housing (21); said turbojet engine additionally comprising: - a low-pressure rotor (RB) which extends downstream of the central shaft (AC) and supports the low-pressure turbine (19); - a rotor bearing (23) supported by the exhaust housing (21); - a reduction gear (22) by means of which the low-pressure rotor (RB) drives the central shaft (AC), said reduction gear being located upstream of the rotor bearing (23); and - a downstream shaft bearing (29) which is located upstream of the reduction gear (22).

IPC Classes  ?

• F02C 7/06 - Arrangement of bearings; Lubricating
• 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

Abstract

The invention relates to a dual-flow turbojet engine having a central shaft (AC) surrounded by a high-pressure body (CH) which rotate about the same longitudinal axis (AX) while being independent in rotation, and comprising: - a fan (13) driven by the central pressure shaft; - a high-pressure compressor (16) and a high-pressure turbine (17) mounted on the high-pressure body (CH); - an inter-turbine casing (18); - a low-pressure turbine (19) mounted on a low-pressure rotor (RB) surrounding the central shaft (AC); - an exhaust casing (21) on which an output cone (23) is mounted; -a reduction gear (22) by means of which the low-pressure rotor (RB) drives the central pressure shaft (AC); - two bearings (34) mounted on the exhaust casing (21) and respectively receiving the central shaft (AC) and the low-pressure rotor (RB); and - a bearing (26) mounted on the inter-turbine casing (18) and receiving the low-pressure rotor (RB).

IPC Classes  ?

• 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
• F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
• F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user

Abstract

The invention relates to a turbojet engine including a shaft surrounded by a low-pressure rotor surrounded by a coaxial and independent high-pressure spool, this turbojet engine including from upstream to downstream:- a fan driven by the shaft;- a low-pressure compressor carried by the rotor;- an inter-compressor casing;- a high-pressure compressor and a high-pressure turbine carried by the high-pressure spool;- an inter-turbine casing;- a low-pressure turbine carried by the rotor;- an exhaust casing;this turbojet engine including:- an upstream rotor bearing carried by the inter-compressor casing;- a downstream rotor bearing carried by the exhaust casing;-a gearbox downstream of the downstream bearing and through which the rotor drives the shaft;- a downstream shaft bearing downstream of the downstream rotor bearing.

IPC Classes  ?

• 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
• F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
• F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user

Abstract

The invention relates to a seal (20) configured to be mounted on a first member of a turbomachine (100), such as a door (16) of a discharge valve (10), and to come to bear against a second member (11, 12, 128) of the turbomachine, such as a hub (11) of an intermediate casing (12), said seal being characterized in that it comprises: latching means (21) for mounting said seal on said first member; a sealing lip (24) that seals with respect to air by contact and that is intended to come to bear against said second member (11, 12, 128); means (23) for preventing said sealing lip (24) from being turned back, increasing the stiffness of the seal (20) at the base of the sealing lip (24).

IPC Classes  ?

• B64C 1/00 - Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
• F01D 25/28 - Supporting or mounting arrangements, e.g. for turbine casing
• F02C 7/28 - Arrangement of seals
• F16J 15/02 - Sealings between relatively-stationary surfaces

Abstract

The invention relates to a propulsion system (10) intended to be mounted on an aircraft comprising a main body, said propulsion system (10) comprising: - a first rotating propulsive member (11) and a second rotating propulsive member (12) that are intended to be mounted on either side of said main body, - a transmission housing (3) connected to the first rotating propulsive member (11) via a first mechanical shaft (51) and to the second rotating propulsive member (12) via a second mechanical shaft (52), - a single gas generator (2) connected to said transmission housing (3) and configured to rotate the first rotating propulsive member (11) and the second rotating propulsive member (12), and - a single auxiliary turbomachine (4) configured to rotate the first rotating propulsive member (11) and the second rotating propulsive member (12) independently of the gas generator (2).

IPC Classes  ?

• B64D 35/04 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission driving a plurality of propellers or rotors
• B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type

Abstract

The invention concerns a flow straightener unit (1) for a fan module of a turbomachine, the straightener unit (1) comprising a plurality of blades (2) distributed about an axis of rotation, each blade (2) is made of a composite material and comprises an aerofoil (21) and a root (22) intended to be assembled on a hub (4) of the turbomachine. The unit (1) comprises a centring and attachment plate (3) of the blade (2) on the turbomachine intended to be attached to the hub (4) at a determined azimuthal position and to the root (22) of the blade (2), the plate (3) is designed to be screwed to the hub (4) by screws (51a) that are longitudinal with respect to the axis of rotation of the unit (1) and screwed to the root (22) of the blade (2) by screws (52) that are radial with respect to the axis of rotation of the unit (1).

IPC Classes  ?

• F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
• F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector

Abstract

The invention relates to a device for monitoring the lifetime of at least one hydraulic apparatus of an aircraft that is subject to variations in hydraulic pressure during flight, comprising an interface for receiving measurement data which are representative of hydraulic pressure (P). The invention is characterised in that the device comprises a processing device, comprising a means for detecting a pressure (P) load (SOLLEND) of a damaging nature, which load is defined by the fact that the pressure (P) comprises a pressure increase (?PAUG) that is greater than a predetermined damage threshold (S?P), followed by a pressure decrease (?PDIM) that is greater than the threshold (S?P), a means for calculating a pressure variation magnitude that is equal to the maximum increase (?PAUG) and the maximum decrease (?PDIM), a means for projecting the magnitude onto a decreasing curve or straight line of a damage model in order to determine the permissible number of loads corresponding to the magnitude, a means for calculating a potential damage ratio that is equal to a number of reference loads divided by the permissible number, a means for increasing a count of accumulated ratios by said ratio.

IPC Classes  ?

• G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
• G01M 99/00 - Subject matter not provided for in other groups of this subclass
• B64F 5/60 - Testing or inspecting aircraft components or systems
• F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups ; Air intakes for jet-propulsion plants
• F02C 9/00 - Controlling gas-turbine plants; Controlling fuel supply in air-breathing jet-propulsion plants
• F28D 1/047 - Heat-exchange apparatus having stationary conduit assemblies for one heat-exchange medium only, the media being in contact with different sides of the conduit wall, in which the other heat-exchange medium is a large body of fluid, e.g. domestic or mo with the heat-exchange conduits immersed in the body of fluid with tubular conduits the conduits being bent, e.g. in a serpentine or zig-zag
• F28D 15/00 - Heat-exchange apparatus with the intermediate heat-transfer medium in closed tubes passing into or through the conduit walls
• F28D 21/00 - Heat-exchange apparatus not covered by any of the groups
• F28F 1/12 - Tubular elements or assemblies thereof with means for increasing heat-transfer area, e.g. with fins, with projections, with recesses the means being only outside the tubular element
• G05B 23/02 - Electric testing or monitoring

Abstract

The invention relates to a turbine shroud sector (2) made of ceramic matrix composite, of longitudinal axis X and which comprises a base (3) with a radially internal face (6), a radially external face (7) from which there extend upstream (4) and downstream (5) tabs for attachment to a shroud support structure, and at least one inter-sector lateral face (9), the base comprising a first slot (10) and a second slot (16), which is arranged radially on the outside of the first slot, which slots are each formed in the lateral face, a first and a second sealing strip (14, 17) respectively resting against a radially internal wall of these first and second slots. According to the invention, the first strip has the overall shape of an omega and the shroud sector exhibits a first clearance which is defined radially between a central part (14a) of the first strip and the radially internal wall of the first slot, and a second clearance which is defined radially between this central part and a radially external wall (15b) of the first slot.

IPC Classes  ?

• F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
• F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
• F01D 11/08 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator

Abstract

The invention concerns an aircraft turbomachine comprising an outer housing (2) delimiting, with an inner hub (3), a flow path (1) for a gas stream in which a low-pressure turbine is arranged, configured to rotate a low-pressure shaft; the turbomachine comprising, in the direction of flow of the gas stream, a first propeller (31); and a second propeller (32) downstream from the first propeller, the first propeller (31) being rotated by the low-pressure shaft and the second propeller being rotated by an electric motor (70), the second propeller (32) further being arranged at a distance of between 1.5 and 4 chord lengths (LC1) from the first propeller (31) defined between the respective pitch axes (A31, A32) of each of the first and second propellers.

IPC Classes  ?

• 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
• F01D 15/10 - Adaptations for driving, or combinations with, electric generators

Abstract

A method for controlling a turbomachine (T) comprising an electric motor (ME) forming a torque injection device on a high-pressure rotation shaft (22), in which method a fuel flow setpoint QCMD and a torque setpoint TRQCMD provided at the electric motor (ME) are determined, the control method comprising: a step of implementing a first fuel control loop in order to determine the fuel flow set point QCMD, a step of implementing a second, torque control loop in order to determine the torque setpoint TRQCMD comprising i. a step of determining a torque correction variable ?TRQCMD as a function of a transitory speed setpoint NHTrajAccelCons, NHTrajDecelCons and ii. a step of determining the torque setpoint TRQCMD as a function of the torque correction variable ?TRQCMD.

IPC Classes  ?

• F02C 9/48 - Control of fuel supply conjointly with another control of the plant
• F01D 15/00 - Adaptations of machines or engines for special use; Combinations of engines with devices driven thereby
• 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

Abstract

The invention relates to a profiled structure which is elongated in a direction in which the structure has a length exposed to an air flow, and transversely to which the structure has a leading edge (164) and/or a trailing edge, at least one of which is profiled and has, in the direction of elongation, clamps (28a) defined by successive teeth (30) and recesses (32). Along the profiled leading edge and/or trailing edge, the successive teeth (30) and recesses (32) extend over only a portion of the length exposed to the flow, over which portion the amplitude and/or spacing of the teeth varies monotonously with the exception of the few teeth closest to each end of said portion, a remaining portion (280) of said length being smooth.

IPC Classes  ?

• F01D 5/14 - Form or construction
• B64C 3/14 - Aerofoil profile
• B64C 11/18 - Aerodynamic features
• F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps

Abstract

Disclosed is a positioning tool (22) for an endoscopic inspection means for inspecting a manifold casing (12) of an aircraft turbine engine (10), this manifold casing comprising an annular wall (14) with an axis of rotation A that comprises through holes (16) and that is surrounded by an annular manifold (18) comprising at least one gas port (20a, 20b) emerging in an annular recess (21) delimited between the manifold and the wall, characterised in that the tool comprises: - a centring device (24) configured to be engaged in the port and comprising a distal end (24a) configured to bear against the wall, and a proximal end (24b) configured to be held securely on an opening (26) of the port, and - at least one mask (28a, 28b, 28c) configured to be engaged in the port and to be positioned and secured on the centring device, this mask comprising an endoscopic inspection window (30, 30', 30'') for inspecting a zone (Z) to be inspected inside the port and/or the recess.

IPC Classes  ?

• G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
• G01D 11/30 - Supports specially adapted for an instrument; Supports specially adapted for a set of instruments

Abstract

The invention concerns a suspension assembly (7) for a turbine engine, comprising a beam (10) intended to be attached to a pylon of an aircraft, and a cylindrical part (15) hinged to a ball-joint housing, the ball-joint housing comprising a body (11), and a ball-joint yoke (16) hinged to the body (11), the cylindrical part (15) being mounted so as to be able to pivot about its axis in the ball-joint yoke, characterised in that the cylindrical part (15) is intended to be secured to a fixed part (8) of the turbine engine (1), the body (11) of the ball-joint housing being secured to the beam (10).

IPC Classes  ?

• B64D 27/26 - Aircraft characterised by construction of power-plant mounting

Abstract

The invention relates to a turbomachine (1) extending along an axis (X), comprising a high-pressure body with a high-pressure compressor (4) coupled in rotation to a high-pressure turbine (6), and a low-pressure body including a low-pressure compressor (3) coupled in rotation to a low-pressure turbine (7), an upstream casing (14) located upstream of the high-pressure compressor (4), and means of suspension (12, 13) for the turbomachine (1), intended for attaching the turbomachine (1) to an aircraft pylon (11).