A turbine engine nozzle extending about an axis and including a vane assembly mounted between an outer shroud and an inner shroud; and a position retaining device configured to be placed in an elastically prestressed state between the inner shroud and an inner platform of the vane assembly to apply a stress on the vane assembly towards the outer shroud. The contact between the inner shroud and the inner platform is obtained by one same surface of the position retaining device.
A distributor for a turbine of a turbomachine. The distributor includes a radially inner platform, a blade and a root which is coated with a track made of an abradable material. According to the invention, the distributor includes a leakage gas reintroduction duct. The duct includes an inlet through the track made of an abradable material, an outlet opening through a downstream surface of the root, and a duct intermediate portion which extends from the inlet up to the outlet. The duct intermediate portion is orientated with a tangential component.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 11/02 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
3.
ASSEMBLY FOR A TURBOMACHINE BLADE, COMPRISING A FASTENER, DEFINING A POCKET, AND A SHIM ADAPTED TO BE RECEIVED IN THE POCKET AT THE SAME TIME AS A ROOT OF THE BLADE
Assembly for a turbomachine blade, comprising: a fastener (9) defining a pocket (10) for receiving a blade root (11), and comprising two flanks defining between them a passage leading into the pocket (10) and forming stops preventing the root (11) from exiting from the pocket (10) via the passage, a shim (26a, 26b) adapted to be received in the pocket (10) with the root (11), the shim (26a, 26b) having a first support surface (28a, 28b) to bear on the root (11) and a second support surface (30a, 30b) opposite the first support surface (28a, 28b) to bear against the fastener (9), the support surfaces being fixed relative to each other and oriented such that movement of the shim (26a, 26b) in the pocket (10) relative to the fastener (9) varies the value of force exerted by the shim (26a, 26b) on the root (11).
INSTITUT SUPERIEUR DE L'AERONAUTIQUE ET DE L'ESPAC (France)
SAFRAN (France)
SAFRAN AIRCRAFT ENGINES (France)
Inventor
Gonidec, Patrick
Aguirre, Miguel Angel
Duplaa, Sébastien
Robic, Bernard
Rodriguez, Benoit
Tantot, Nicolas Jérôme Jean
Abstract
A propulsion unit including: a gas generator including a compressor, a combustion chamber, a turbine, at least two compressed-air propulsion modules, each propulsion module including: a fan, a compressed-air turbine configured to drive the rotation of the fan, a manifold allowing the respective turbines of the compressed-air propulsion modules to be supplied with compressed air, wherein the manifold is configured to collect and mix: at least a portion of the flow that has passed through the combustion chamber of the gas generator, typically as it leaves the turbine, and at least one bypass flow, the bypass flow being a flow of air which is not passed through the combustion chamber of the gas generator.
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type
F04D 17/02 - Radial-flow pumps specially adapted for elastic fluids, e.g. centrifugal pumps; Helico-centrifugal pumps specially adapted for elastic fluids having non-centrifugal stages, e.g. centripetal
5.
METHOD AND MODULE FOR DETECTING THE STATE OF A COUPLING DEVICE, AND ASSOCIATED TURBINE ENGINE AND AIRCRAFT
This method for controlling the operation of a coupling device (18) for a turbine engine (4) of an aircraft (2) comprises the steps of: - a first comparison (48) of the speed of rotation of an electric machine (16) which is determined with a profile of evolution with respect to time, the first comparison (48) being implemented to determine an operating deviation; - a second comparison (50) of the operating deviation with a detection threshold; and - according to the result of the second comparison step (50), an identification of the existence or absence of a failure (52) of the coupling device (18); the second comparison step (50) comprising the determination of the value of the detection threshold from at least one control parameter (P), the operating deviation and the rotational speed of the electric machine (16).
An accessory gearbox for an aircraft turbine engine having a radial shaft connecting a high-pressure body of the turbine engine to a gear train having at least one mixed shaft coupled to a high-speed device. The accessory gearbox has a clutch system having a low-speed motor mounted thereon and configured to mate with the mixed shaft such that: in an engaged position of the clutch, the low-speed motor is coupled to the mixed shaft so as to drive the high-pressure body at low speed when the turbine engine is stationary and, in a disengaged position of the clutch, the low-speed motor is decoupled from the mixed shaft so as to be protected when the turbine engine is in operation.
A method for generating an alarm is implemented by an alarm-generating device installed in an aircraft. It comprises: a step of obtaining information delivered by a sensor of said device and representative of an oil flow rate inside an oil filter of the aircraft; a step of obtaining a temperature delivered by a sensor of said device and representative of a temperature of the oil inside said filter; a step of obtaining a differential pressure at said filter; and a step of generating an alarm if the differential pressure exceeds an alarm threshold obtained on the basis of said information and of said temperature.
The turbomachine subassembly comprises:—at least one tow-pressure nozzle ring, the nozzle ring comprising a root (38) integral with the nozzle ring;—a clamp (40) configured so as to fasten the root to a casing of the turbomachine; and—at least two pads (150) interposed between the root and the clamp, each pad being assembled in a form-fitting manner with one of either the root or the clamp.
The invention relates to a device (B) for regulating a system (A) for supplying hydrogen to a turbine engine comprising a combustion chamber, the hydrogen supply system comprising, in particular arranged in series, at least: - a pressurising member (2), and - a metering member (5) advantageously connected to the combustion chamber, characterised in that it also comprises - a hydrogen accumulator (4) arranged between the pressurising member (2) and the metering member (5), - a first regulator (21) that is able to control the pressurising member (2) depending on a hydrogen pressure in the hydrogen accumulator (4) and on a setpoint hydrogen pressure in the hydrogen accumulator (4), and - a second regulator (31) that is able to control the metering member (5) depending on a flow rate of hydrogen injected into the combustion chamber and on a setpoint flow rate (CQ) of hydrogen to be injected into the combustion chamber.
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
F02C 7/224 - Heating fuel before feeding to the burner
F02C 9/40 - Control of fuel supply specially adapted to the use of a special fuel or a plurality of fuels
10.
TURBOJET ENGINE EQUIPPED WITH MEANS FOR MAINTAINING ITS OPERATION IN THE EVENT OF FAILURE OF ITS MAIN FUEL SUPPLY CIRCUIT
The invention relates to a turbojet engine comprising: a combustion chamber and a jet pipe; a main circuit (11) including a main centrifugal pump (14) and a main positive displacement pump (16) supplying fuel to the injectors (18) of the combustion chamber, via a main regulator (17); an auxiliary circuit (13) including an auxiliary positive displacement pump (24) for supplying pressurized fuel to the actuators (25) of the jet pipe via an auxiliary regulator (26); a post-combustion circuit (12) including a post-combustion centrifugal pump (21), a post-combustion regulator (22) and post-combustion injectors (23) supplied by the post-combustion pump (21) via the post-combustion regulator (22); according to the invention, a backup system (27) connected to the auxiliary regulator (26) and to the injectors (18) is provided so as to supply fuel to these injectors (18) using the auxiliary pump (24) in order to restart the turbojet engine in the event of a failure of the main circuit (11).
F02K 1/17 - Control or regulation conjointly with another control with control of fuel supply
F02C 7/236 - Fuel delivery systems comprising two or more pumps
F02K 3/10 - Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof by after-burners
11.
FUEL CONDITIONING SYSTEM AND METHOD FOR POWERING A TURBOMACHINE
The invention relates to a fuel conditioning system (SC) configured to power an aircraft turbomachine (M) with fuel (Q) from a cryogenic tank (R), the conditioning system (SC) comprising: a fuel circuit (1); a distribution valve (5) configured to divide a main fuel flow (Q) into a primary fuel flow (Q1) flowing through a primary branch (11) of the fuel circuit (1) and a secondary fuel flow (Q2) flowing through a secondary branch (12); a manifold (6) configured to mix the primary fuel flow (Q1) and the secondary fuel flow (Q2) and reform the main fuel flow (Q); a third mechanical pump (23) and a recirculating heat exchanger (4) installed in the secondary branch (12).
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
F02C 7/224 - Heating fuel before feeding to the burner
F02C 9/40 - Control of fuel supply specially adapted to the use of a special fuel or a plurality of fuels
12.
SYSTEM AND METHOD FOR CONTROLLING THE MODIFICATION OF THE PITCH OF THE BLADES OF A TURBINE ENGINE
A system for controlling the modification of the pitch of the blades of a fan of a turbine engine, in particular for an aircraft. The turbine engine comprising blades mounted radially in a drive shaft and a setting device configured to modify the pitch of the blades on the basis of an axial force applied to said setting device. The control system comprises a hydraulic actuator, a hydraulic pump, a connecting ring connected mechanically to the hydraulic pump such that the flow rate of the hydraulic pump is proportional to the relative speed between the drive shaft and the connecting ring, and a magnetic coupling device designed to control the drive speed of the connecting ring in order to control the pitch of the blades independently of the speed of the drive shaft.
An air inlet of a nacelle of an aircraft propulsion assembly comprising an inner wall and an outer wall which are connected upstream by an air inlet lip. The air inlet lip comprising thick portions and thin portions distributed alternately over the circumference of the air inlet lip. Each thick portion comprising an upstream leading edge to separate an upstream air flow into an outer air flow guided by the outer wall and an inner air flow guided by the inner wall during a thrust phase. Each thin portion comprising a downstream leading edge situated longitudinally downstream of each upstream leading edge so as to detach a reverse air flow at the air inlet lip during a thrust reversal phase.
A turbomachine component such as a blade, includes a wall provided with at least one cooling hole having a metering portion and a tapered diffusion portion. The diffusion portion includes a central edge forming a bottom of the diffusion portion that is oblique with respect to a flow axis of the cooling hole.
A ventilation ring for a bearing support member of an aircraft turbine engine includes two or more tubular walls and one or more spacer walls. The two or more tubular walls extend opposite each other, and one or more of the tubular walls includes a main portion and a base, the base having a thickness greater than a thickness of the main portion. The one or more spacer walls connect the two or more tubular walls. The one or more spacer walls have one or more apertures extending into the base and a rim around the aperture. An entirety of the rim has a thickness greater than a thickness of the spacer wall at a distance from the rim.
A fuel conditioning system (SC) configured to supply an aircraft turbine engine with fuel (Q) coming from a cryogenic tank (R), the conditioning system (SC) comprising: a fuel circuit (1); a first heat module (31) mounted in an aircraft frame of reference (REF-A) and a second heat module (32) mounted In a turbine engine frame of reference (REF-M); a first circulation loop (41) for a first heat transfer fluid (F1) mounted in the aircraft frame of reference (REF-A), the first heat transfer fluid (F1) being configured to heat the fuel flow (Q) in the first heat module (31); a second circulation loop (42) for a second heat transfer fluid (F2) mounted in the turbine engine frame of reference (REF-M) and distinct from the first circulation loop (41), the second heat transfer fluid (F2) being configured to heat the fuel flow (Q) in the second heat module (32).
F02C 3/22 - Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being gaseous at standard temperature and pressure
F02C 7/224 - Heating fuel before feeding to the burner
F02C 9/40 - Control of fuel supply specially adapted to the use of a special fuel or a plurality of fuels
17.
VARIABLE PITCH VANE FOR A TURBOMACHINE FAN HAVING A STIFFNESS GRADIENT IN THE ROOT
Disclosed is a variable pitch vane for a turbomachine fan comprising a blade (60) and a root (62). The root (62) comprises a bulb (72) and an upright support (74) connecting the bulb (72) to the blade (60), the bulb (72) being connected to the upright support (74) by a neck (82) defining a local minimum of the cross-section of the root (62) along a plane orthogonal to the pitch axis. At least one section of the root (62) including the neck (82) and at least a portion of the upright support (74) comprises a surface layer and an inner layer between the pitch axis and the surface layer, the inner layer having a stiffness strictly greater than the stiffness of the surface layer.
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
B64C 11/06 - Blade mountings for variable-pitch blades
A propulsion assembly for an aircraft comprising a dual-flow turbine engine equipped with a fan, an aerodynamic outer shroud acting as a nacelle as well as a mounting pylon, the propulsion assembly having a secondary flow path defined by an outer radial defining surface formed by the shroud, the turbine engine including stator vanes, and the mounting pylon comprising a part housed in the secondary flow path, referred to as upstream part. According to the invention, the upstream part of the pylon extends radially from the inner radial defining surface, along a radial pylon height strictly less than a total radial height of the secondary flow path, and the upstream part of the pylon extends in the downstream direction from a root part of one of the stator vanes.
A rotor of a turbine extending around a longitudinal axis includes a rotor disc bearing at its periphery one or more rotor blades, each blade including an airfoil provided at its radially outer end with a platform equipped with an upstream lip and a downstream lip, the platform having a front lateral edge and a back lateral edge with respect to the normal direction of rotation of the rotor about the longitudinal axis, and a suction face portion which extends from the front lateral edge to the suction face of the airfoil. The upstream lip and the downstream lip each includes sawtooth-shaped cut-outs such as to form a front tooth and a back tooth which are contiguous, in that the back tooth has a cutting edge and in that the front tooth of each lip is disposed on the suction face portion of the platform.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 5/20 - Specially-shaped blade tips to seal space between tips and stator
20.
TREATMENT OF NON-AXISYMMETRIC CASING WITH CONTROLLED OPENING
The invention relates to a casing (200) of a turbine engine compressor, comprising openings (201) which are made in the thickness of the casing from an inner face (2001) of the casing and arranged next to one another on a circumference of the casing, characterised in that it also comprises a movable ring (220) which is formed by at least two annular portions, is present on an outer face (2002) of the casing opposite the openings and is able to move along the circumference of the casing so as to open and close the openings of the casing in order to activate or deactivate a casing treatment, the movable ring comprising the same number of slits (202) as there are openings in the casing.
A method for repairing an end plate of a turbomachine rotor, the end plate including a crown having at least two collar fastening holes and at least one balancing-weight fastening hole located between the two collar fastening holes, each collar fastening hole being equipped with a crimped nut, the crown having a damaged portion between two collar fastening holes, the method including the steps of removing the two crimped nuts located on either side of the damage; removing the damaged crown portion, for example by machining; putting in place a strip in the form of a crown portion closing off the removed crown portion, the strip having at least one balancing-weight fastening hole and two mounting holes; fastening the strip to the crown with two crimped nuts, which penetrate the mounting holes, by crimping each nut in a collar fastening hole.
The invention relates to a system comprising: a harness for a weaving loom and a device for monitoring the manufacture of a preform woven by the loom, the monitoring device comprising: a plurality of sensors, each sensor of the plurality of sensors being connected to one of the plurality of strings and being configured to measure a deformation of the string, the deformation being induced by a force exerted by the loom on the string to which the sensor is connected; and a processing unit connected to the plurality of sensors and configured to analyse the deformation so as to identify a manufacturing abnormality on at least one string.
METHOD FOR PREDICTING THE WEAR OF A MECHANICAL PART, AND METHOD FOR DETERMINING THE OPERATIONAL RISK ASSOCIATED WITH AN AIRCRAFT OR WITH A FLEET OF AIRCRAFT
A method for predicting the wear of a mechanical part and uncertainty of this prediction in a profile of use, the profile of use taking into account environmental conditions associated with environmental data and the time of use of the mechanical part under each of these environmental conditions, the method including determining operational data associated with a plurality of mechanical parts of the same type as the mechanical part; on the basis of the operational data determined, determining a plurality of predictive models of the wear of the part, the mechanical part being able to be divided into a plurality of elements, each element being modelled using at least one model of the plurality of models; and, for each model, determining a weighting coefficient, determining a wear prediction of the mechanical part, and determining a statistical quantity representative of dispersion of the predictions of the plurality of models.
One aspect of the invention relates to a method (2) for monitoring a turbomachine (1) comprising a stator (11), a rotor (12) comprising blades (13), a strain gauge (15) attached to the stator (11) and a position sensor (14) for detecting the position of the blades (13), the gauge (15) being configured to capture a first signal comprising two components, the position sensor (14) being configured to capture a second signal representing the position of each blade (13), the method (2) comprising, on each revolution of the rotor (12), with one revolution being detected on the basis of the second signal: acquiring (21) the first and second signals; angularly resampling (22) the first signal; extracting (23) the first component from the resampled first signal by sliding-filtering the resampled first signal; and, on the basis of the filtered resampled first signal, calculating (24) at least one contact indicator for each blade (13) for the revolution.
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
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
25.
TOOLING AND METHOD FOR MANUFACTURING A COMPOSITE BLADE FOR AN AIRCRAFT ENGINE
Tooling (40) for manufacturing a blade (10) made of composite material for a turbomachine, in particular of an aircraft, comprising: - a mould (30) and a counter-mould (34) which define between them a cavity (32) configured to receive a woven preform, the cavity (32) having a first part (Z1) configured to receive a shield (22) and at least one edge of the preform, and a second part (Z2) configured to receive at least part of the remainder of the preform, and - elements (42, 44, 46) for managing the temperature of the cavity (32), which elements are configured to heat the first and second parts (Z1, Z2) of the cavity (32) at different temperatures during at least one step of a method for manufacturing the blade (10).
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 33/02 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor with incorporated heating or cooling means
B29C 35/02 - Heating or curing, e.g. crosslinking or vulcanising
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
The invention relates to a turbine engine comprising: - a first module comprising a speed reducer (3) having an input shaft (12), - a second module comprising a low-pressure compressor which is connected to the first module, - a third module comprising a low-pressure shaft which is centred on a longitudinal axis X and which comprises an upstream end connected to the input shaft, and - a locking device configured to axially immobilise the second module with respect to the first module and to the second module, the locking device comprising: - a first nut (60) screwed onto a thread of the low-pressure shaft and in contact with an annular projection of the second module, - an anti-displacement member (65) configured to axially lock the first nut, - a second nut (73) screwed onto a thread of the second module and configured to axially immobilise the anti-displacement member, and - a third nut (93) screwed onto the anti-displacement member.
A turbine stator assembly including a guide vane assembly and a sealing ring bearing an abradable element of a dynamic sealing ring. The assembly includes pins secured to the sealing ring and cooperating with respective oblong openings formed in the guide vane assembly so as to allow a radial movement of the ring with respect to the guide vane assembly in order to compensate for differential thermal expansions. A seal is arranged so as to compensate for circumferential spaces between ring sectors forming the guide vane assembly.
An acoustic treatment panel including acoustic absorption cells each comprising a bottom wall, a porous inlet wall, an enclosure extending in an axial direction between the bottom wall and the inlet wall, and a first acoustic horn extending inside the enclosure between a first opening and a second opening that is smaller than said first opening, the first opening facing said inlet wall. Each cell comprises a second horn extending inside the enclosure in the first axial direction between a first opening of the second horn and a second opening of the second horn that is smaller than said first opening.
F02K 1/34 - Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow for attenuating noise
F02C 7/045 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression
29.
METHOD AND SYSTEM FOR DETERMINING AIRCRAFT LANDING RUNWAY CONDITIONS
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 (C1) derived from the modified weighting coefficients.
Discharge valve of an aircraft turbomachine, including a discharge gate intended to be mounted pivotably about an axis between a position for closing an air passage orifice and an open position of said orifice, the orifice being formed in an intermediate casing of the turbomachine, the discharge valve including a sealing device having a seal configured to be movably mounted on the intermediate casing and a movement system configured to move the seal between a first position in which the seal is intended to bear against the discharge gate when the gate is in a closure position, and a second position in which the seal is retracted with respect to the air passage orifice when the discharge gate is in an open position.
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
F16K 3/10 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing with flat sealing faces; Packings therefor with pivoted closure members with special arrangements for separating the sealing faces or for pressing them together
31.
DRAINAGE MEMBRANE FOR THE MANUFACTURE OF COMPOSITE MATERIALS
The invention relates to a method for manufacturing a part made of composite material, said method comprising: - the arrangement of a fibrous preform (10) in a mould comprising an impregnation chamber (201), the impregnation chamber (201) being closed by a membrane (230) separating the impregnation chamber (201) from a compaction chamber (202), - the injection of an impregnation fluid (5) into the impregnation chamber (201), and the injection of a compression fluid (6) into the compaction chamber (202) so as to apply pressure to the membrane (230), - the aspiration of the compression fluid (6) present in the compaction chamber (202), the method being characterized in that the surface (230b) of the membrane (230) present on the side of the compaction chamber (202) comprises a plurality of grooves (235a, 235b).
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
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
B28B 1/26 - Producing shaped articles from the material by slip-casting, i.e. by casting a suspension or dispersion of the material in a liquid-absorbent or porous mould, the liquid being allowed to soak into or pass through the walls of the mould; Moulds therefor
C04B 35/80 - Fibres, filaments, whiskers, platelets, or the like
The invention relates to a turbomachine comprising: a first splitting edge capable of separating an incoming air flow into a radially internal air flow and a secondary air flow; a second splitting edge capable of separating the radially internal air flow into a primary flow and a tertiary flow which traverses a tertiary flow path that is radially external to a primary flow path traversed by the primary flow; a heat exchanger (18) arranged in the tertiary flow path; and an inner casing (28); characterized in that the exchanger comprises a body (32) and a flange (32.1) which extends radially inside and projects from the body, the flange being attached to the inner casing, the exchanger comprising, downstream of the flange, a downstream portion (40) to which a fire wall (46) forming a heat shield is attached.
F02K 3/077 - 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 the plant being of the multiple flow type, i.e. having three or more flows
F02C 7/14 - Cooling of plants of fluids in the plant
33.
THREE-STREAM AXIAL TURBINE ENGINE WITH A SEALED HEAT EXCHANGER IN THE THIRD STREAM
The invention relates to a turbine engine comprising: a first splitting edge capable of separating an incoming air stream into a radially internal air stream and a secondary air stream; a second splitting edge capable of separating the radially internal air stream into a primary stream and a tertiary stream which traverses a tertiary-stream flow path that is radially external to a primary-stream flow path traversed by the primary stream; a heat exchanger (18) arranged in the tertiary-stream flow path; and an inner casing (28); characterized in that the exchanger comprises an upstream portion (50) comprising a collar (52) which projects from a body (32) of the exchanger in the upstream direction, the collar being accommodated in a floating manner in a groove (54) that runs circumferentially in the inner casing.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
34.
INTER-TURBINE CASING FOR A TURBINE ENGINE, TURBINE ENGINE ASSEMBLY, TURBINE ENGINE AND METHOD FOR POSITIONING A SPLITTER VANE WITHIN AN INTER-TURBINE CASING
The invention relates to an inter-turbine casing (8) for a turbine engine of the turbine vane frame type, which casing acts as a turbine distributor in such a turbine engine, the inter-turbine casing comprising an inner shroud (9), an outer shroud (10), a plurality of arms (11) extending between the inner shroud and the outer shroud, and at least one set of N splitter vanes (12) positioned circumferentially between two successive arms. Each splitter vane has a mid-height axial chord shorter than the mid-height axial chord of the arms. The two successive arms define a reference position for each of the N splitter vanes. These reference positions are evenly spaced circumferentially between the two successive arms. At least one splitter vane (12-1) of the set of N splitter vanes is circumferentially offset from the reference position thereof. The invention further relates to a turbine engine assembly comprising the inter-turbine casing and a turbine (7) extending downstream of the inter-turbine casing (8) and comprising at least one movable vane (13) extending radially. The invention further relates to a turbine engine comprising the turbine engine assembly. The invention further discloses a method for positioning a splitter vane within an inter-turbine casing.
The invention relates to a sealing device (1) comprising: a ring (10) that comprises a main sealing lip (12) and a secondary sealing lip (14), the main sealing lip having an inner surface (11), the secondary sealing lip (14) having a radial surface (13), and a casing (20) comprising a cylindrical portion (22) and a radial portion (24), the radial portion (24) having a contact surface (23), the ring (10) being biased so as to apply the radial surface (13) of the secondary sealing lip (14) against the contact surface (23) of the casing (20), wherein the sealing device (1) has at least one leakage path (30) that extends between the secondary sealing lip (14) and the casing (20).
The invention relates to a method for detecting an anomaly in the operation of a component of an aircraft associated with at least one status indicator, the method comprising, for at least one acquired time sequence (Seq1, Seq2) comprising acquired values of said at least one indicator, steps of: - determining (E20) an approximated time sequence (Seq1*, Seq2*) of approximated values of said acquired values by means of an approximation module (AE1, AE2); - determining (E30) an overall abnormality score (A1, A2) for the acquired time sequence (Seq1, Seq2) on the basis of differences between the acquired values (Seq1, Seq2) and the approximated values; and - detecting (E40) an anomaly in the operation of the component on the basis of a comparison of the overall abnormality score with a first threshold, the approximation module being configured to minimize approximation errors between time sequences of reference values and time sequences of approximated values.
The invention relates to a method for manufacturing a steel part (10) for an aircraft, the method comprising the welding of at least two steel elements (12, 14) to one another so as to form this part (10), the application of a resistance treatment to the part (10) which comprises austenitization (30) followed by quenching (32) and at least one step of tempering (34), and the local application, to a weld bead (16) between the at least two elements (12, 14), of a softening treatment comprising heating the weld bead (16) to a softening temperature lower than the temperature Ac1 at which the steel starts to transform from ferrite into austenite.
B23K 20/12 - Non-electric welding by applying impact or other pressure, with or without the application of heat, e.g. cladding or plating the heat being generated by friction; Friction welding
A method for manufacturing a turbomachine vane made of composite material, the method including forming, by three-dimensional or multilayer weaving, a fiber blank separated across its thickness into first, second and third parts in two non-interlinked areas present at the longitudinal ends of the fiber blank, the first part located between the second and third parts to which it is connected by weaving outside non-interlinked areas, forming, from the fiber blank, a preform of the vane to be produced, by unfolding and shaping, at each longitudinal end and on either side of the first part, the segments of the second part and the segments of the third part not interlinked with the first part to form preform parts for a platform of the vane, the segments of the first part not interlinked with the segments of the second and third parts extending along the longitudinal axis, and densifying the preform.
An oil jet for lubricating members such as bearings inside a turbine engine, includes a main pipe for supplying oil, at least one secondary oil discharge pipe through which the oil is sprayed onto the member, and a connection area which is for connecting the secondary pipe to the main pipe and which includes a junction angle between the main pipe and the secondary pipe, wherein the connection area includes an inner wall in contact with the oil, the inner wall being at least partially curved at the junction between the secondary pipe and the main pipe such that the junction angle inside the connection area is at least partially rounded.
An inner ring for supporting variable-pitch blades of an axial compressor stator of a turbomachine includes two or more ring sectors, assembled to form the inner ring, each ring sector including plural housings for receiving a pivot axis of one of the blades. At each junction plane between the ends of two successive ring sectors, an assembly pin and a cavity for receiving the assembly pin are engaged. The assembly pin is fixed on one end one of the two successive ring sectors and the cavity is formed at the end of the other of the two successive ring sectors. The cavity receives the assembly pin with a mounting clearance allowing a mounting by sliding fit, and the assembly pin is made of a material having a coefficient of thermal expansion greater than a coefficient of thermal expansion of a material constituting the successive ring sectors.
A counter-rotating turbine for a turbine engine comprising an inner rotor having an inner drum to which is attached a plurality of inner impellers rotatably supported by a first shaft, an outer rotor comprising an outer drum to which is fastened a plurality of outer impellers rotatably supported by a second shaft coaxial with the first shaft, the outer rotor comprising a downstream impeller having a plurality of downstream moving blades extending between an outer shroud and an inner shroud, one upstream end of the outer shroud being attached downstream of said outer drum, the inner shroud being attached to the second shaft, at least one of the inner shroud and of the outer shroud comprising at least one flexible transition part configured to allow elastic deformation of said shroud in the radial direction.
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
The invention relates to manual deburring tool (1) comprising: - a consumable carbide plate (10) having deburring edges (12), and - a casing (20) which is configured to be grasped manually and has an internal cavity (22) for housing the plate, the casing (20) comprising a body (25) and a cover (27) removably attached to the body so as to allow the insertion of the plate (10) into the cavity (22) and the removal of said plate from the cavity, the housing comprising at least one opening (26) that opens into the cavity and exposes a deburring edge (12).
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE PAUL SABATIER TOULOUSE III (France)
Inventor
Rousselle, Mélanie, Laura, Josépha
Fradet, Guillaume
Estournes, Claude, Gilbert, Jean-Pierre
Ansart, Florence
Herisson De Beauvoir, Thomas
Abstract
The invention relates to a method for producing an abradable ceramic composite coating on a substrate, the method comprising: obtaining (E1) a composition (30) in powder form comprising a matrix powder and a ceramic filler hydrated precursor powder having a lamellar crystallographic structure, wherein the ceramic filler powder represents from 5 to 40% of the combined volume of the matrix powder and the ceramic filler powder; compressing the prepared powder composition at a pressure greater than 150 MPa; and a step of reactive sintering (E2) the obtained powder composition, during which the pressure is maintained at a temperature of less than 550°C, and the particles of the matrix powder in the sintered powder composition have an aspect ratio of 2 or greater. The invention also relates to an abradable ceramic coating obtained according to the method. The invention also relates to a superalloy part for a turbomachine, for example a turbine part, comprising such a coating.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
C23C 24/08 - Coating starting from inorganic powder by application of heat or pressure and heat
F01D 25/00 - Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
C04B 111/00 - Function, property or use of the mortars, concrete or artificial stone
44.
PROPELLER BLADE OR VANE HAVING A HOLLOW COMPOSITE ROOT
The invention relates to a propeller blade or vane (10) for a turboprop engine, made from composite material and comprising a matrix-densified fibrous reinforcement, the propeller blade or vane comprising, in the direction of its span (DL), a root (12) and an aerodynamic profile (11). The fibrous reinforcement comprises a fibrous preform having three-dimensional weaving, with a root preform portion and an aerodynamic-profile preform portion. The fibrous preform comprises a disconnected portion delimiting a housing that forms a cavity (14) extending both into the root (12) and into the aerodynamic profile (11). A spar (130) is present in the cavity (14), the spar comprising an aerodynamic-profile shaping portion (131) positioned in a first portion (142) of the cavity (14) and a root shaping portion (132) positioned in a second portion (141) of the cavity. The root (12) has a rotationally symmetrical shape.
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
B29C 70/72 - Encapsulating inserts having non-encapsulated projections, e.g. extremities or terminal portions of electrical components
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
The invention relates to a cluster of wax models, around which a mould (1) for manufacturing a plurality of bladed elements by lost-wax casting is intended to be formed, the model having a longitudinal axis (X) and comprising: a replica of a casting cup (2), a replica of a central sprue (3), a plurality of replicas of mould elements (4), and a plurality of replicas of source feed ducts (5) of the mould elements (4). According to the invention, each replica of a source feed duct (5) is connected to a replica of a mould element (4) via replicas of fluid communication elements (6a, 6b, 6c).
A turbine ring assembly extending around an axis, including a plurality of ring sectors made of ceramic matrix composite material forming a turbine ring and a ring support structure held by a turbine casing, each ring sector including a base from which an upstream tab and a downstream tab extend radially outward, spaced axially from one another, wherein the assembly also includes a cross-member mechanically connected to the ring support and including a first radial flange bearing against the upstream latching tabs of the ring sectors, and, for each ring sector, at least two transverse pins and a radial spring, each transverse pin passing through the upstream latching tab and the downstream latching tab of the ring sector and the ring support to hold the ring sector and the ring support secured to one another.
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
47.
HYBRIDIZATION OF THE FIBERS OF THE FIBROUS REINFORCEMENT OF A FAN BLADE
The invention relates to a fan blade (3) made of composite material the fibrous reinforcement of which comprises first strands (12) having a first stiffness, second strands (13) having a second stiffness and third strands (14) having a third stiffness lower than the first stiffness and greater than the second stiffness, the blade comprising a first portion (15) comprising only first strands (12), a second portion (16) comprising only second strands (13) and a third portion (17) which is located between the first and second portions (15, 16) and which comprises both first, second and third strands (12, 13, 14); there is a gradual transition of properties at the interface between the different portions (15, 16, 17).
Sector (6) of a guide vanes assembly for a turbine of an aircraft turbomachine, the sector (6) comprising an outer platform (7) and an inner platform (8) which are coaxial along an axis (X), the sector (6) further comprising at least one blade (9) which connects the outer and inner platforms (7, 8) together, characterized in that the blade (9) is variable-pitch about an axis of rotation (Y) of the blade (9), the blade (9) comprising an aerodynamic body (11) delimited radially by a head and a root, the head being placed with a first functional clearance in an opening of the outer platform (7), the root being placed with a second functional clearance in a cavity of the inner platform (8), the axis of rotation (Y) of the blade (9) being located downstream of the head and the root.
The invention relates to a device for guiding a shaft of an aircraft turbine engine, the device comprising: - a rolling bearing (10, 11); - an annular bearing support (22); and - an oil film compression damping system (19) comprising: - two annular sealing segments (26, 27); and - a first oil supply circuit (33), this first circuit being connected to an annular supply groove (34), characterized in that the damping system further comprises: - a second oil supply circuit (41), this second circuit being connected to at least one annular temperature control groove (40) which is independent of the supply groove (34) and which runs around one of the sealing segments (26, 27).
The invention relates to a guide vane (33B) for attachment to a stator shroud of a gas turbine engine, comprising: - a profiled part (36B) intended to extend in a gas flow in order to guide the gas flow, the profiled part (36B) having a pressure-side surface (43B) and a suction-side surface (44B), and - a platform (34B) having a guide surface (45B) from which the profiled part (36B) extends, a first lateral surface (46B) and a second lateral surface (47B), the second lateral surface (47B) being able to be arranged facing a first lateral surface (46A) of an identical adjacent guide vane (33A), forming a raised portion which prevents a parasitic transverse flow of the gas flow, while creating a depression which at least partially compensates for the obstruction of the gas flow by the raised portion.
An installation for compacting strands or a preform including strands, the strands or the preform including at least a first portion in which the strands are twisted at a first angle and a second portion in which the strands are twisted at a second angle different from the first angle, the installation including a mould which includes the strands or the preform, and wherein the mould has at least a first part in which the first portion of the strands or of the preform is disposed, having a first thickness and a second part in which the second portion of the strands or of the preform is disposed, having a second thickness different from the first thickness.
B29C 43/02 - Compression moulding, i.e. applying external pressure to flow the moulding material; Apparatus therefor of articles of definite length, i.e. discrete articles
B29C 70/20 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in a single direction, e.g. roving or other parallel fibres
52.
DEVICE AND METHOD FOR INJECTING A HYDROGEN-AIR MIXTURE FOR A TURBINE ENGINE BURNER
The invention relates to a device for injecting a combustible mixture, for a combustion chamber (100) of an aircraft turbine engine turbine, which comprises, about a longitudinal axis (X), a tubular central channel (1), a first annular channel (2) about the central channel and a second annular channel (3) about the first annular channel (2), the channels (1, 2, 3) opening into the combustion chamber at a first lip (9) of the central channel (10), of a second lip (10) of the first annular channel and of an end (11) of the second annular channel, the first annular channel comprising, upstream of the second lip (10), a device (5, 6, 7) for injecting dihydrogen into the first annular channel (2) in an air flow (8) passing along the longitudinal axis of the first annular channel so as to create a dihydrogen-air mixture flowing towards the combustion chamber.
Method for checking the compliance of a workpiece, comprising the following steps: estimating a risk of non-compliance of characteristic on the basis of a law of probability associated with the characteristic; and verifying whether the estimated risk of non-compliance satisfies a decision criterion and, if so, declaring that the workpiece is compliant for characteristic; if not, measuring a value of the characteristic, determining whether or not the workpiece is compliant based on the measured value, and updating the law of probability associated with characteristic based on the measured value.
G05B 19/418 - Total factory control, i.e. centrally controlling a plurality of machines, e.g. direct or distributed numerical control (DNC), flexible manufacturing systems (FMS), integrated manufacturing systems (IMS), computer integrated manufacturing (CIM)
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE BORDEAUX (France)
INSTITUT POLYTECHNIQUE DE BORDEAUX (France)
Inventor
Cavarroc, Marjorie, Christine
Peigney, Erwan
Aymonier, Cyril
Aubert, Guillaume
Poulon, Angélique, Nadine, Jeanne
Abstract
The present invention relates to a method for depositing a continuous coating of aluminium oxide on a metal substrate by means of induction heating-assisted pressurised, temperature-controlled chemical deposition, the method comprising a solvothermal synthesis step based on an aluminium oxide precursor dissolved in a water-co-solvent mixture heated by induction to a temperature of between 400°C and 700°C and a pressure of between 1 MPa and 25 MPa. It also relates to a device (100) for depositing an aluminium oxide on a metal substrate (104) by means of pressurised, temperature-controlled chemical deposition, which device comprises: - an chamber (102) delimited by walls forming a closed volume (V), the chamber being intended to contain a heated and pressurised fluid, the material of the walls of the chamber being transparent to electromagnetic radiation; - a support (106) transparent to electromagnetic radiation intended to support the metal substrate (104) located inside the chamber; - an induction heating device (109) surrounding the outside of the chamber (102) so as to be able to heat the metal substrate (104) positioned on the support (106); - an inlet (116) located in the upper portion of the chamber (102) and configured to allow a precursor material previously dissolved in water to be added to the chamber (102); - an inlet (120) located in the lower portion of the chamber (102) and configured to allow a fluid to be added to the chamber (102); - at least one outlet (124) configured to purge the volume (V); - a sapphire window (112) arranged in the upper portion of the chamber, allowing the temperature of the metal substrate (104) to be controlled by a bichromatic pyrometer (114) arranged outside the chamber; - a set of polymer seals (200); - a metal assembly (202) rigidly screwed together by metal columns (204) containing a circulating fluid maintained, by a cryostat (206), at a temperature of 20°C.
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coating; Contact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
C01B 13/36 - Methods for preparing oxides or hydroxides in general by precipitation reactions in solutions
C01F 7/448 - Dehydration of aluminium oxide or hydroxide, i.e. all conversions of one form into another involving a loss of water by wet processes using superatmospheric pressure, e.g. hydrothermal conversion of gibbsite into boehmite
The invention relates to a method for producing a moulded part and to the three-dimensionally moulded part itself, comprising a composite material including a fibrous preform and a resin, the part having a cavity (230') filled with a cellular material (50). The cellular material comprises an expanded cellular polymeric material.
B29C 44/14 - Incorporating or moulding on preformed parts, e.g. inserts or reinforcements the preformed part being a lining
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
B29C 33/52 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
B29C 70/16 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length
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
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
B29C 43/36 - Moulds for making articles of definite length, i.e. discrete articles
B64F 5/10 - Manufacturing or assembling aircraft, e.g. jigs therefor
56.
NICKEL-BASED SUPERALLOY, SINGLE-CRYSTAL BLADE AND TURBINE ENGINE
ECOLE NATIONALE SUPERIEURE DE MECANIQUE AEROTECHNIQUE (France)
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE POITIERS (France)
Inventor
Menou, Edern
Cormier, Jonathan
Rame, Jérémy
Abstract
The invention relates to a nickel-based superalloy comprising, in weight percentages, 5.0 to 6.0% aluminium, 6.5 to 8.5% tantalum, 0 to 1.0% titanium, 1.0 to 4.0% cobalt, 5.0 to 8.0% chromium, 0 to 0.5% molybdenum, 3.0 to 4.0% tungsten, 3.75 to 5.75% rhenium, 3.5 to 5.0% platinum, 0.05 to 0.25% hafnium and 0 to 0.15% silicon, the remainder being made up of nickel and inevitable impurities. The invention also relates to a single-crystal blade (20A, 20B) comprising such an alloy and to a turbine engine (10) comprising such a blade (20A, 20B).
C22C 19/05 - Alloys based on nickel or cobalt based on nickel with chromium
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
57.
SHIM FOR A TURBINE ENGINE TURBINE DISC, ASSEMBLY FOR A TURBINE ENGINE ROTOR, TURBINE ENGINE ROTOR, TURBINE ENGINE, AND METHOD FOR INSTALLING AN ASSEMBLY FOR A TURBINE ENGINE ROTOR
The invention relates to a shim (400) for a turbine engine turbine disc (40), comprising a metal sheet forming an envelope that at least partially surrounds a tooth (42) of the turbine disc, the tooth separating a first and a second cavity (41) of the disc, each cavity being adapted to receive a root (20) of a movable blade (10) of the turbine. The envelope is adapted to at least partially cover: - each of the side walls (44) of the tooth, and - a downstream rim (46) of the tooth (42). The invention also relates to an assembly for a turbine engine rotor, to a turbine engine rotor, and to a method for installing a turbine engine rotor assembly.
A turbine ring assembly having ring segments made of ceramic matrix composite material and each having first and second attachment tabs and a cavity for the circulation of air flow, a metal support having a first bracket and a second bracket bearing axially upstream against the second tab, a first metal flange arranged upstream of the first bracket and having an inner periphery bearing axially downstream against the first tab and an outer periphery fastened to the first bracket, and air passage orifices formed in the inner periphery of the first flange and/or in the second bracket, the orifices configured to ensure that the air flow passes from the cavity to the outside of the assembly.
A hybrid turbofan engine for an aircraft, comprising a fan, an electric motor/generator and a gas generator, the engine comprising, between the fan and a splitter that separates the flows, an internal wall for delimiting an air duct, situated upstream of inlet guide vanes of a primary duct, and, upstream of outlet guide vanes, an internal upstream wall for delimiting a secondary duct, the motor/generator having a stator carried by a stator support fixed to a stator part of the engine. According to the invention, a plurality of heat pipe for cooling the motor/generator are provided, each heat pipe having an evaporation section fixed to the stator support, and a condensation section fixed to the internal wall or the internal upstream wall.
The present application relates to a blade (3) of a static vane assembly (2) of a turbine engine (1) comprising: - a blade root (8) comprising a fibrous reinforcer (13) comprising a first skin (12) and a second skin (12) separated by a cavity (14) that opens onto a lower face (15) of the blade root (8); - a fastener (6) configured to receive the blade root (8) and attach it to a hub (4) of the vane assembly (2), the fastener (6) comprising a platform (16) and a rib (17) protruding from the platform (16), the rib (17) being configured to penetrate the cavity (14) so as to extend between internal faces (12a) of the first skin (12) and of the second skin (12); and - an attachment system (18) configured to mechanically attach the first skin (12) and the second skin (12) to the rib (17) of the fastener (6).
1233) across its width (L) by means of weaving between a plurality of warp yarns (30) and a plurality of weft yarns (31), the loom comprising: a plurality of heald wires (22), which are connected to a mechanism (11) capable of moving the heald wires in a vertical direction (Dv) between a reference position (PH) and a first position for opening the warp yarns, each heald wire being further provided with an eye (23) through which a warp yarn (30) passes, a rapier (40) arranged downstream of the heald wires (113), capable of drawing a weft yarn (31) from a bobbin (50), a device (90) for holding the fabric arranged downstream of the plurality of heald wires (22) and of the rapier (40), comprising two bearing surfaces (93, 94) arranged on either side of the fabric (T) to compress the fabric between the bearing surfaces (93, 94), characterised in that at least a first bearing surface (93) is shaped so as to be at least partially complementary to a surface of the fabric.
GENERAL ELECTRIC DEUTSCHLAND HOLDING GMBH (Germany)
GENERAL ELECTRIC COMPANY (USA)
Inventor
Delbosc, Philippe
Reigner, Pierre-Alain Jean Philippe
Osama, Mohamed
Zatorski, Darek
Abstract
A turbine engine is disclosed, which comprises a first electric machine coupled to a low pressure shaft of the turbine engine and a second electric machine coupled to a high pressure shaft of the turbine engine. The turbine engine includes one or more engine loads and an external electrical connection for communicating electrical power between the turbine engine and an electrical system of an aircraft. The turbine engine includes a power management system including a high voltage DC busbar configured to communicate electrical power between the first and second electric machines and the engine loads and the external electrical connection. The power management system includes a control system configured to control power offtake from and injection to the first and second electric machines, and controls power delivered to the one or more engine loads and the external electrical connection.
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 15/10 - Adaptations for driving, or combinations with, electric generators
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
SYSTEM AND METHOD FOR PROVIDING NAVIGATION ASSISTANCE TO A MOBILE SYSTEM BY MEANS OF A MODEL FOR PREDICTING THE TRAVERSABILITY OF THE MOBILE SYSTEM ON A TERRAIN
The invention relates to a navigation assistance method for a mobile system, the method comprising: - generating training data, comprising: o obtaining a sequence of images of an environment acquired by a camera (2) installed on board a vehicle (1) moving through the environment; o determining a footprint (EN) of the vehicle on the ground during the acquisition of an image in the sequence; o marking, as an area traversable by the mobile system, an image region (RI) corresponding to the projection of the footprint in another image in the sequence; - training, by machine learning, a model for predicting the traversability of the mobile system on a terrain by means of the training data; - obtaining images of a scene that are acquired by a camera installed on board the mobile system; and - predicting the traversability of the scene by supplying images of the scene to the model trained by machine learning.
Propulsion assembly for an aircraft comprising a turbomachine having at least one rotating part rotating about an axis of rotation, an attachment strut, and a structural element carrying the turbomachine via the attachment strut, the rotating part being disposed upstream of the structural element and of the attachment strut such that an air jet emerging from the rotating part, in the wake of thereof, impacts the structural element and the attachment strut, a leading edge of the structural element and/or of said attachment strut locally comprising at least one acoustic attenuation device disposed at least partly in the wake of the rotating part, the acoustic attenuation device being a local modification of the structure and/or of the profile of the leading edge.
A device for shaping at least one fibrous preform of a bladed part of a turbine engine, the device including a mould formed of multiple parts nested inside one another, the mould defining an internal cavity for enclosing the preform entirely, the cavity having two platform zones and a blade zone extending between the two platform zones, wherein the mould includes at least a lower shell, an upper shell, a side shell, and end shells, and in that each of the side and end shells includes three elements, respectively lower, intermediate and upper.
b), on which a fastening tab is located (34), said fastening tab being configured to be fixed to a fan disc (2). The invention is characterized in that it comprises the steps of: a) preparing fabrics or sheets which are pre-impregnated with a resin, b) depositing the fabrics or sheets in a mold, c) positioning a metal reinforcement (36) in the mold on the fabrics or sheets, the reinforcement being integrally formed with said fastening tab, d) depositing the fabrics or sheets on a part of the reinforcement, and e) closing and heating the mold for solidification of the assembly formed by the fabrics or sheets and the reinforcement.
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
B29C 70/72 - Encapsulating inserts having non-encapsulated projections, e.g. extremities or terminal portions of electrical components
B29K 705/00 - Use of metals, their alloys or their compounds, for preformed parts, e.g. for inserts
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
The present invention relates to a turbomachine assembly (1), comprising: - a high-pressure compressor (5) configured to be driven by a high-pressure shaft (12); - a low-pressure compressor (4) configured to be driven by a low-pressure shaft (11), the low-pressure shaft (11) being configured to be driven at a lower speed than the high-pressure shaft (12); - a first power converter (21) configured to be driven by the high-pressure shaft (12); - an air-circulation system comprising a first air bleed port (9) positioned in the high-pressure compressor (5) and a second air bleed port (3) positioned upstream of the first air bleed port (9); - means (50) for determining an operating condition of the turbomachine assembly (1); and - a controller (40) configured to send, to the first power converter (21), instructions to transfer power from the high-pressure shaft (12) to the low-pressure shaft (11) or power from the low-pressure shaft (11) to the high-pressure shaft (12), according to the operating condition.
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 5/00 - Plants including an engine, other than a gas turbine, driving a compressor or a ducted fan
The invention relates to an assembly screw comprising at least a threaded part (12A) and a non-threaded part (12C), the non-threaded part being provided with a strain gauge (22) that is sensitive to the instantaneous elongation of the non-threaded part and with an RFID tag (24) connected to the strain gauge by a wired link, the assembly allowing the wireless transmission of an electrical resistance value of the strain gauge representative of the instantaneous elongation.
Disclosed is a stator element (200) comprising: - an electrical conductor (202) that is elongate and hollow in order to define a flow channel (308) allowing a coolant to flow from one end to the other of the conductor (202), the conductor (202) being designed to pass an electric current; and - a connection element (208; 902) located at one of the ends of the conductor (202), comprising a fluid connection terminal (210) designed to allow coolant to enter the channel (308) or indeed coolant to exit from the channel (308), and an electrical connection terminal (212) designed to electrically connect the conductor (202). The conductor (202) and the connection element (208; 902) are formed of a single unitary part.
A device for moulding at least one bladed part of a turbine engine, including a base; a mould formed from a plurality of parts nested inside one another, this mould being applied to the base. A first seal is mounted between the mould and the base; and a bell is mounted on the mould and around the mould, this bell being applied to the base. A second seal is mounted between the bell and the base, this bell being configured to be held tight against the base and having interior surfaces engaging by wedge effect with complementary exterior surfaces of the mould in order to apply a clamping force on the parts of this mould.
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 33/00 - SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING - Details thereof or accessories therefor
B29C 61/06 - Making preforms having internal stresses, e.g. plastic memory
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
71.
AIRCRAFT TURBINE ENGINE ASSEMBLY COMPRISING A HOLDER FOR FITTINGS
Assembly for an aircraft turbine engine, comprising an intermediate casing hub, an accessory gearbox, and a transfer case including a power transmission shaft configured to transmit mechanical power, taken off an engine shaft of the turbine engine, to at least one accessory mounted on the accessory gearbox, in which the transfer case comprises a structural element surrounding the power transmission shaft and rigidly attaching the accessory gearbox to the intermediate casing hub.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
F02C 7/20 - Mounting or supporting of plant; Accommodating heat expansion or creep
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
72.
TURBOMACHINE MODULE EQUIPPED WITH AN ELECTRIC MACHINE, AND TURBOMACHINE EQUIPPED WITH SUCH A MODULE
An aircraft turbomachine module, including an accessory gearbox, a transfer shaft, and a drive shaft rotatably connected with one another by a power transmission device. The power transmission device can be housed in a gearbox and the module can further include an electric machine rotatably connected to the transfer shaft. The electric machine can be interposed between the transmission housing and at least one part the transfer shaft.
Assembly including an aircraft turbine engine and a pylon for mounting the turbine engine to an element of the aircraft. The pylon includes members for suspending the turbine engine, the members being connected to the turbine engine in at least one plane which is perpendicular to the axis and which is located upstream of the combustion chamber of the turbine engine, such that the turbomachine is cantilevered to the pylon, The assembly further includes at least one damper which connects the turbomachine to the pylon and which is located in a plane perpendicular to the axis located downstream of the combustion chamber, the damper being configured to limit the relative movements between the turbine engine and the pylon without transmitting force.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
Schneider-Die-Gross, Julien Paul
Jailin, Clément
Mendoza Quispe, Arturo
Roux, Stéphane
Abstract
Method for processing a plurality of X-ray tomography volume images each associated with a part, the plurality of volume images comprising a reference volume image, including:
a step of correlating volume images to obtain a displacement field between each image and the reference image, to obtain a plurality of displacement fields minimizing the difference between the volume images,
a processing by a dimensionality reduction method of the plurality of the image displacement fields to express them according to eigenmodes, and
a statistical analysis of the fields expressed according to the eigenmodes.
G06T 15/00 - 3D [Three Dimensional] image rendering
75.
ASSEMBLY COMPRISING A RING AND A PIVOTING SUPPORT SLEEVE FOR VARIABLE PITCH BLADE ROOTS, TURBINE ENGINE PROVIDED WITH SUCH AN ASSEMBLY AND METHOD FOR DISMANTLING SUCH AN ASSEMBLY
The invention relates to an assembly for a turbine engine, in particular of an aircraft, the assembly comprising a ring (2) with a rotationally symmetrical axis (O). According to the invention, the assembly comprises a plurality of support sleeves (12) each having a rotationally symmetrical axis (O) and each intended to receive a stud (8) of a variable pitch blade (3), each support sleeve (12) comprising a bore (30) passing right through the support sleeve (12) along the rotationally symmetrical axis (O), each support sleeve (12) being mounted on the ring (2) so as to pivot about a pivot axis (C, C') transverse to the radial axis (Z) and to the rotationally symmetrical axis (O).
The invention relates to a propeller (10) for an aircraft turbine engine, this propeller (10) having: - a hub (12), - vanes (14), - bearings (46, 48) for guiding the roots (18) of the vanes (14) into orifices (12a) in the hub (12), and - systems for retaining the roots (18) of the vanes (14) in the orifices (12a) in the hub (12) along the pitch axis (A), this system for retaining the root (18) comprising: + a ring (66) configured to be fitted in a groove (32) in the root (18), and + a screw-nut system (68) mounted on the root (18) between the ring (66) and one of the bearings (48). Figure for the abstract: Figure 4
The present invention relates to a protection element (45.1, 45.2) for a compressor drum (20), which is intended to be arranged at least in part between a blade root (25) and the compressor drum (20), comprising: - a first lateral wall (51.1, 51.2) comprising an inner portion (54.1, 54.2) intended to cover an inner rounded portion of a pocket (24), a bearing surface portion (55.1, 55.2) intended to cover a blade root bearing surface of the pocket (24), and an outer portion (56.1, 56.2) intended to cover an outer rounded portion of the compressor drum (20), - a connecting wall (53.1, 53.2) intended to cover an outer face of the compressor drum (20), and - a second lateral wall (52.1, 52.2) comprising a projecting portion (57.1, 57.2) intended to engage with a mounting groove (46.1, 46.2) formed in a vertical face of the compressor drum (20).
The invention relates to a turbine engine (1) for an aircraft, comprising an annular lubrication enclosure (17) in which a speed reducer (12) is arranged, a system for lubricating the speed reducer (12) comprising a main lubrication circuit (24) and an auxiliary lubrication circuit (25) connected to the lubrication enclosure (17), the auxiliary circuit (25) being connected to an auxiliary reservoir (31), an inner shroud (19) extending around the lubrication enclosure (17) and the auxiliary reservoir (31) being located outside the outer shroud (18), one of the arms, referred to as the arm at 6 o'clock (20), being tubular and located at the 6 o'clock position, the lubrication enclosure (17) comprising a deflector (22) connected to the radially inner end (20a) of the arm at 6 o'clock (20) and configured to deflect oil flowing by gravity from the lubrication enclosure (17) towards the upstream compartment (21a) of the arm at 6 o'clock (20).
The present disclosure relates to a method for regulating the voltage (V_AC) of an electrical bus (411_AC) of an electrical system of an electrically hybridized turbomachine, the bus (411_AC) being intended to transport an AC current, the method (E) comprising controlling a magnetic flux within a permanent-magnet synchronous machine (401) of the electrical system, the machine (101) being connected to the bus (411_AC), the control being carried out on the basis of a voltage (V_AC) measured on the bus (411_AC) and of a rotational speed (Vrot) of the rotor of the machine (401) in relation to the stator of the machine (401) by driving an inverter (421) of the electrical system, the inverter (421) being connected to the bus (411_AC).
H02J 7/14 - Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries for charging batteries from dynamo-electric generators driven at varying speed, e.g. on vehicle
An auxiliary tank for an aircraft turbine engine is provided, and in conjunction with a pump and associated auxiliary lubrication circuit, supplies oil to a reducer when, for example, a phase of free rotation of the fan is detected. Thus, the reducer is always lubricated, even during the phases of free rotation of the fan, thus ensuring a longer life of the reducer gears. The auxiliary tank is arranged with respect to the reducer so that oil is recovered from the reducer at least in part by the auxiliary tank.
A guide vane assembly of an aircraft turbine engine, the guide vane assembly being segmented into a plurality of segments arranged circumferentially next to one another about an axis. The guide vane assembly includes two annular rings connected together by vanes and the geometry of the inner ring is modified to reduce the vibrational responses of the guide vane assembly
A blade and a sealing element forming an assembly for a turbomachine turbine. The blade includes an airfoil and a platform to which the sealing element is connected. An internal channel passes through the platform so as to draw off fluid circulating in a primary duct that it delimits and to inject the thus drawn-off fluid into an inter-lip cavity delimited by the sealing element.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 11/10 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
83.
TURBOMACHINE ROTOR HAVING IMPROVED VIBRATORY BEHAVIOUR
Assembly comprising a body extending around a central axis, the body having an outer surface from which a plurality of blades extend, each of said blades having a blade root and a blade tip, defining an inner radial end and an outer radial end of the blade relative to the central axis, said blades having the same blade height measured radially relative to the central axis, characterized in that each of the blades is connected to the body by its blade root via a connection having a nonzero connection height, so that for the plurality of said blades, the connection height of two successive blades is different.
The present invention relates to a method for controlling an electrical system, comprising: processing a main command signal (IP) for commanding drawing of a mechanical power by the electrical system from an electrically hybridized turbomachine so as to extract from said signal a first frequency component (CF1) corresponding to a natural frequency of a first rotary body of the turbomachine and to generate a filtered first main signal (IF1) without the first frequency component (CF1); generating a first subsidiary command signal (CTRL_1) for commanding drawing of a mechanical power by the electrical system from the first rotary body based on the filtered first main signal (IF1); and generating a second subsidiary command signal (CTRL_2, CTRL_3) for commanding drawing of a power by the electrical system from an auxiliary power source based on the first frequency component (CF1).
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
INSTITUT NATIONAL POLYTECHNIQUE DE TOULOUSE (France)
UNIVERSITE PAUL SABATIER - TOULOUSE III (France)
Inventor
Azzopardi, Stéphane, Joseph
Le, Thanh Long
Richardeau, Frédéric
Jouha, Wadia
Barazi, Yazan
Rouger, Nicolas
Blaquiere, Jean-Marc
Vinnac, Sébastien
Picot, Mathis
Abstract
The invention relates to a device (100) for controlling and protecting a power transistor (102), comprising: - a nominal switching circuit (103) for the transistor; - a short-circuit detection circuit (105) which keeps the transistor in the conducting state and detects an increase or decrease in the voltage VGS of the transistor relative to reference voltages representative of a short-circuit of the transistor; - a protection circuit (107) which discharges the gate of the transistor after the detection of a short-circuit; and - a circuit (106) for measuring and controlling the nominal switching circuit, the short-circuit detection circuit and the protection circuit.
H03K 17/0812 - Modifications for protecting switching circuit against overcurrent or overvoltage without feedback from the output circuit to the control circuit by measures taken in the control circuit
H03K 17/18 - Modifications for indicating state of switch
H03K 17/16 - Modifications for eliminating interference voltages or currents
According to one aspect, the invention relates to a device for de-icing a turbine engine blade (1) of axis Y, comprising a blade (2) that is rotatably movable about a pivot (20) of radial axis X and comprising an electrical resistor (60), an electrical power supply (6) connected to the electrical resistor (60), and a control lever (3) for the blade (2) connected to a control ring (11), the device being characterised in that the control lever (3) comprises a pair of conductive strips (32) connected to the lever for conjoint rotation and located on the side opposite the control ring (11) with respect to the radial axis X, and in that each conductive strip (32) of the pair of conductive strips comprises a first end in contact with a circular conductive track (4) connected to the electrical power supply (6) and comprises a second end connected to the electrical resistor (60).
The present invention relates to a method for applying a coating to a substrate (100) made of a nickel-based superalloy and to a turbine blade made of a nickel-based superalloy with a coating applied according to this method. The coating application method comprises at least a first chemical vapour deposition step for depositing chromium and/or cobalt on a surface of the substrate (100) and a second chemical vapour deposition step, after at least partially diffusing, in an underlying layer (102) of the substrate (100), the chromium and/or the cobalt deposited in the first chemical vapour deposition step, for depositing aluminium on the underlying layer (102) of the substrate (100).
C23C 16/06 - Chemical coating by decomposition of gaseous compounds, without leaving reaction products of surface material in the coating, i.e. chemical vapour deposition (CVD) processes characterised by the deposition of metallic material
C23C 16/04 - Coating on selected surface areas, e.g. using masks
C23C 10/28 - Solid state diffusion of only metal elements or silicon into metallic material surfaces using solids, e.g. powders, pastes
C23C 28/02 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of metallic material
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
88.
METHOD FOR OBTAINING AN ANTI-OXIDATIVE COATING FOR A TITANIUM ALLOY PART
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE CENTRALE DE LYON (France)
UNIVERSITE CLAUDE BERNARD LYON 1 (France)
Inventor
Senani, Sophie
Toury, Bérangère
Hermange, Kurt, Gérard, Yves
Benayoun, Stéphane
Desgranges, Clara
Delfosse, Jérome
Abstract
The present invention relates to a method for obtaining an anti-oxidative coating for a titanium alloy part, comprising: - depositing a liquid composition on the part, comprising at least (i) a first sol-gel alkoxysilane precursor, (ii) a second precursor which is an organic-Inorganic sol-gel precursor of an element E1 or a salt of element E1, said element E1 being chosen from among silicon, aluminum, zirconium, titanium, tin, zinc, phosphorus and rare earths, (iii) a third precursor which is an organic-inorganic sol-gel precursor of an element E2 or a salt of element E2, said element E2 being chosen from among silicon, aluminum, zirconium, titanium, tin, zinc, phosphorus and rare earths, E1 and E2 not both being silicon, and (iv) water, and - hydrolyzing the first, second and third precursors of the deposited liquid composition and condensing same so as to form the anti-oxidative coating which comprises an interconnected mixed oxide network of the elements silicon, E1 and E2.
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coating; Contact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
The invention relates to a stator assembly for an aircraft turbomachine extending about an axis and including: a shroud ring, a plurality of bladed ring sectors, sliding assembly allowing relative movement between the shroud ring and each of the ring sectors, the sliding assembly including two outer annular flanges which are upstream and downstream of the shroud ring, respectively, and an inner flange of each of the ring sectors which is inserted axially between the two outer flanges, wherein each of the ring sectors includes a deflector which is supported by the inner flange of the sector and extends around the upstream annular flange of the shroud ring.
The invention relates to a system (3) comprising one or more light sources (21) which are each capable of delivering light pulses, one or more light detectors (31), an optomechanical head (160) and an optical fibre-connecting device (350) connecting the one or more light sources to the optomechanical head and connecting the one or more detectors to the optomechanical head, in which system the optomechanical head comprises at least three separate collimators optically associated with three optical fibres (51, 52, 53) of the connecting device, the three fibres being connected to the one or more light sources or being connected to the one or more light detectors.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE CENTRALE DE LYON (France)
UNIVERSITE CLAUDE BERNARD LYON 1 (France)
Inventor
Senani, Sophie
Toury, Bérangère
Hermange, Kurt, Gérard, Yves
Benayoun, Stéphane
Desgranges, Clara
Delfosse, Jérome
Abstract
The invention relates to a method for obtaining an anti-oxidative coating for a titanium alloy part, comprising: - depositing a liquid composition on the part, comprising at least (i) a first precursor which is a metal-organic sol-gel precursor of a metal element E1 or a salt of metal element E1, said metal element E1 being chosen from among aluminum and zirconium, (ii) a second precursor which is an organic-inorganic sol-gel precursor of an element E2 or a salt of element E2, said element E2 being chosen from among aluminum, zirconium, titanium, tin, zinc, phosphorus or rare earths and being different from metal element E1, and (iii) water, and - hydrolyzing the first and second precursors of the deposited liquid composition and condensing same so as to form the anti-oxidative coating which comprises an interconnected mixed oxide network of elements E1 and E2.
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coating; Contact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
37 - Construction and mining; installation and repair services
Goods & Services
Aircraft propulsion systems, as well as components and spare parts thereof, including the following products: Motors, Turbines, Pumps [machines], Propellers, Reactors, nacelles, thrust reverser, Engine air intakes, blower crankcases, exhaust cones; Compressors; Heat exchangers [parts of machines]; Couplings and transmission and propulsion components. Repair services, overhaul, Maintenance and Maintenance In connection with the following goods: propulsion systems for aeronautical vehicles, Parts of the aforesaid products and Spare parts for the previously mentioned goods.
93.
Method for manufacturing a turbomachine compressor blade by compacting
The invention relates to a method for producing a turbomachine compressor blade, comprising the following steps:—installing primary pins (26) comprising a material other than a titanium-based alloy in primary bores (20) of a core, the primary bores forming at least one polygon, and installing a secondary pin made of titanium-based alloy in a secondary bore of the core; —producing a stack (2) of a suction-face sheet (4), a core (14) and a pressure-face sheet (6); —compacting the stack; —removing the primary pins (26) from the primary bores (20); —removing the secondary pin from the secondary bore; and—taking the core (14) away from the stack.
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
B21D 53/78 - Making other particular articles turbine blades
B23K 20/02 - Non-electric welding by applying impact or other pressure, with or without the application of heat, e.g. cladding or plating by means of a press
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
94.
METHOD FOR SURFACE TREATMENT BY SELECTIVE REMOVAL OF A BONDING PRIMER ON A TITANIUM OR TITANIUM ALLOY SUBSTRATE
A method for treating the surface of a titanium or titanium alloy metal reinforcement of a blade made of composite material enables selective removal of a bonding primer with respect to the titanium or titanium alloy reinforcement. The method includes subjecting the metallic reinforcement to a thermal treatment performed at a temperature of between 250 and 350° C. for a period of between 1 hour and 10 hours in an oxidizing atmosphere. The method further includes subjecting the metallic reinforcement, after the thermal treatment, to a chemical pickling in an alkaline bath.
A turbine engine extends along an axis (X) and includes a flow path of a primary flow (F1) that has a compressor, a combustion chamber, and a turbine. The turbine engine further includes an outflow region of a secondary flow (F2) that surrounds the primary path, a blower or a propeller located upstream of the primary path and the outflow region of the secondary flow (F2), at least one arm extending radially through the primary path, and at least one fluid circulation pipe extending inside the arm. The arm includes an inlet for air from the primary path so as to cool the fluid circulating in the pipe.
A gear changing system for the driving of an apparatus by an accessory gearbox comprising: a train comprising a ring gear (5), a planetary gear (6) integral with an output shaft (4) and a planet carrier (7) integral with an input shaft (3), a clutch device (10) for the input shaft (3) and output shaft (4), and a locking device (9), the system comprising a first configuration in which the clutch device (10) is in the engaged position and the locking device (9) is in the disconnected position, to drive the output shaft (4) at the speed of the input shaft (3), and a second configuration (C2) in which the clutch device (10) is in the disengaged position (D) and the locking device (9) is in the connected position (B), to drive the output shaft (4) at a speed (N4) higher than the speed (N3) of the input shaft.
F16H 3/54 - Gearings having only two central gears, connected by orbital gears with single orbital gears or pairs of rigidly-connected orbital gears comprising orbital spur gears one of the central gears being internally toothed and the other externally toothed
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
F02C 3/113 - 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 with variable power transmission between rotors
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F16H 59/46 - Inputs being a function of speed dependent on a comparison between speeds
F16H 63/30 - Constructional features of the final output mechanisms
F16H 59/68 - Inputs being a function of gearing status
97.
CASE COMPRISING INTERNAL AND/OR EXTERNAL STIFFENERS
The invention relates to a turbine engine case (1), extending around an axis, said case (1) comprising: a hub comprising an outer wall (21) and an inner wall (22) between which an annular flange (23) extends; an outer shell (3); a plurality of arms (4) which each extend between the hub and the outer shell (3); the case (1) further comprising a plurality of main stiffeners (5), each being disposed in the extension of an arm (4) between the outer wall (21) and the inner wall (22) projecting from the flange (23), each main stiffener (5) comprising two ribs (51, 52) arranged opposite one another and which move away from one another from the outer wall (21) towards the inner wall (22).
A heat exchange system for a turbine engine is provided. The heat exchange system includes a cavity having an air intake, a heat exchanger arranged in the cavity and having a first circuit in which a first fluid can circulate, a movable flap mounted at the air intake and moving between two positions permitting or preventing, respectively, the circulation of air flow in the cavity, and a control device having a movable member configured to drive the movement of the movable flap. The control device can be arranged in the heat exchanger supply circuit and configured so as to permit or prevent the circulation of the first fluid to the heat exchanger and simultaneously move the movable flap between at least one of the two positions.
The present invention relates to a method and system for managing the drawing of mechanical power from a two-spool or three-spool turbine engine (1) for an aircraft, in which at least two electric machines (3, 4) are suitable for recovering mechanical energy, one from a shaft driven by one of the turbines of the turbine engine (1), the other from a shaft driven by another turbine, in which the distribution of the drawing between one and the other of the two electric machines (3, 4) is controlled dynamically according to the phases of flight.
An oil recovery device for a reducer of an aircraft turbomachine, the device having a gutter extending around an axis and being intended to be arranged facing oil ejection means formed in a ring gear of the reducer of the turbomachine. The gutter can include at least one oil reception chamber which extends around the axis and which is open towards the axis so as to receive the oil coming from the oil ejection means. The reception chamber can be delimited by a first wall and a second wall arranged facing each other. The first wall can have an axial oil evacuation orifice and the second wall can have a radial dimension relative to the axis which increases increasingly along the axis towards the axial orifice, so as to allow oil ejected by the reducer to be conveyed to the axial orifice.
