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 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
7.
METHOD FOR REPAIRING AN END PLATE OF A TURBOMACHINE ROTOR
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.
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.
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
11.
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
13.
IMPROVED FERRULE FOR COUNTER-ROTATING TURBINE IMPELLER
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
14.
METHOD FOR MANUFACTURING A VANE MADE OF A COMPOSITE MATERIAL WITH INTEGRATED ATTACHMENT LUGS AND PLATFORMS
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.
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).
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
19.
ADJUSTMENT OF TWIST TO OPTIMIZE MOULDS AND THE SHAPING OF TEXTILE ARCHITECTURES
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
20.
METHOD AND DEVICE FOR CHECKING THE COMPLIANCE OF A WORKPIECE
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)
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.
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
24.
IMPROVED ACOUSTIC ATTENUATION DEVICE FOR AN AIRCRAFT PROPULSION UNIT
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.
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
27.
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
28.
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.
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
34.
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 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 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
37.
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.
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.
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.
A device for setting the pitch of blades for a turbine engine, including a control system acting on a linkage mechanism and including an actuator having two chambers; —a supply member coupled to a fluid transfer bearing for transferring a pressurised fluid from a main supply source in a fixed frame of reference to the chambers; —an auxiliary supply circuit which is arranged in a rotating frame of reference and is connected to the supply member and to the main supply source, the supply member being configured to: —in normal operation, allow fluid to pass from the main supply source to the chambers and to the auxiliary supply circuit, and —in the event of a break in the supply to the chambers from the main supply source, allow pressurised fluid to pass from the auxiliary supply circuit to the chambers.
A method is provided for additive manufacturing of a wall for a turbine engine, the wall including a first cooling aperture. The manufacturing method includes additively manufacturing the wall by selective melting or selective sintering on a powder bed. The wall is manufactured at least partially around the first cooling aperture with at least one lower zone that has a minimum length between 0.01 and 0.4 mm and/or with a lower zone that has a total thickness between 0.06 and 0.22 mm.
The invention relates to a blade comprising a shield (14) attached to the upstream end of the body of the blade, the shield (14) comprising an upstream end forming a leading edge of the blade, the shield (14) further comprising a nose (141) upstream from which the leading edge is located, an pressure-face fin and an suction-face fin laterally attached on the blade, the fins extending from the nose, the blade comprising a defrosting air passage duct (15) arranged inside the nose and extending radially inside the nose, the duct (15) having an inner end (151) emerging opposite the root (125) and a radially external end emerging from the nose between the leading edge and its junction at the pressure-face fin.
An aircraft turbomachine, including a first shaft having external splines oriented parallel to an axis of rotation of the shaft, a second shaft having internal splines complementary to the external splines, this second shaft being fitted axially on the first shaft and these shafts rotating as one by engagement of their splines, and a system for axially locking the shafts with respect to one another, wherein the system includes screws that are oriented radially with respect to the axis, each of these screws being screwed into a first orifice in one of the shafts and having a free end for engaging a second orifice in the other of the shafts.
A part made of composite material, in particular for an aircraft turbine engine, includes a body produced from fibers woven in three dimensions and immersed in a resin, in particular an organic or polymer resin. The part has at least one crack sensor capable of detecting a crack in the part, at least one link element connected to the sensor and configured to transmit a signal (S), in particular to a member outside of the part, when a crack in the part is detected. The crack sensor has at least one conductive wire woven with the fibers of the body, and the link element is configured to measure the resistance of the conductive wire and to wirelessly transmit the signal.
A method for manufacturing a dual-material turbine engine disc, includes the following operations: providing a rough bore made of a first material, mounting the rough bore about an axis of rotation of a rotating device, rotating the rough bore, spraying a second material under solidification conditions, thereby generating a column-like or monocrystalline microstructure, which is different from the first material, on an outer surface of the rough bore in order to produce a dual-material part, and machining the dual-material part to produce a turbine engine disc.
F01D 5/34 - Rotor-blade aggregates of unitary construction
C23C 4/12 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
C23C 4/08 - Metallic material containing only metal elements
An aircraft turbine engine assembly, including a lubricating oil circuit and a fuel supply device, the assembly also including a passive bypass valve enabling the oil to bypass an exchanger, the valve including a valve body and a piston arranged to move in a sliding space defining a first actuating chamber supplied with oil from the oil circuit, as well as a second actuating chamber supplied with fuel from the device, the valve being configured so that when the differential pressure between the oil pressure in the first actuating chamber and the fuel pressure in the second actuating chamber drops below a predetermined value of differential pressure, the piston moves from a normal operating position to an exchanger bypass position.
An air intake for an aircraft propulsion assembly extending along a longitudinal axis and comprising a turbine engine that comprises a primary flow path and a secondary flow path for respectively guiding a primary air flow and a secondary air flow during a thrust, and thrust reversal means for changing the secondary airflow into a reverse airflow during a thrust reversal. The air intake comprising a peripheral external enclosure comprising, in each plane radial to the axis, a point of maximum curvature for detaching the reverse airflow, an osculating circle defining a radius of curvature that is defined at each of the points of maximum curvature. The average value of the radii of curvature being less than a product of 0.028 times an internal radius of the intake at the fan.
A turbine ring assembly extending around a longitudinal axis and comprises:
a plurality of ring angular sectors made of CMC material circumferentially arranged in such a way as to form a turbine ring, each sector comprising a base from which an upstream leg and a downstream leg extend radially,
a ring support structure comprising:
a spacer extending around the sectors and comprising a downstream flange in such a way that the downstream leg of each sector is held against the downstream flange,
a upstream force-absorbing plate extending circumferentially around the longitudinal axis in such a way as to be in contact with the upstream leg of each sector and fixed to an upstream area of the spacer which extends in extension of the upstream leg.
A rear part for a turbojet engine includes a variable-geometry nozzle in which a convergent flap is provided with a lever supporting a bearing roller arranged axially between an upstream bearing wall and a downstream bearing wall, which are rigidly secured to a movable part capable of moving axially on command with respect to an upstream stator structure in such a way that the bearing roller is free to move with respect to the upstream and downstream bearing walls in a radial direction. During a downstream movement of the movable part, the upstream bearing wall pushes the lever to pivot a downstream end of the convergent flap towards the longitudinal axis. Operation of the drive mechanism for moving the movable part can thus be optimized while limiting the size and the mass of the mechanism for controlling the flaps.
Disclosed is a main shaft for an aircraft turbine engine drive line extending longitudinally from front to rear along an axis and having a rear portion configured for receiving a front portion of a rear shaft by interlocking, the main shaft having: —a front retaining member and a rear retaining member together defining a prison and each having a central opening, —a connecting ring, located in the prison, having a rear end configured to be screwed to a front end of the rear shaft extending into the central opening of the rear retaining member, the connecting ring having a gripping member formed on the inner surface and accessible from a front portion of the main shaft by a tool extending via the central opening of the front retaining member.
A method for automatically searching for at least one given textile pattern in a composite material reinforcement including a plurality of textile patterns, each textile pattern including a plurality of reinforcing yarns arranged according to a textile topology, the method including acquiring a three-dimensional image of the composite material reinforcement, and searching for the given textile pattern in the acquired three-dimensional image, using an artificial neural network trained on a training database to detect the given textile pattern in a three-dimensional image of a composite material reinforcement.
The invention relates to a method for controlling a turbine engine having an electric motor forming a device for injecting torque on the high-pressure rotation shaft, in which method a setpoint for fuel flow into the combustion chamber and a setpoint for torque supplied to the electric motor are determined, the control method having: a step of determining a temperature correction value as a function of a temperature parameter of the gases leaving the turbine engine and a maximum value of the temperature parameter of the gases leaving the turbine engine, a step of determining a torque correction value as a function of the temperature correction value and a step of determining the torque setpoint as a function of the torque correction value.
An aircraft turbomachine having a longitudinal axis includes a high-pressure (HP) body, a low-pressure (LP) body, a fan, and a reduction gear having an epicyclic gear train. The turbomachine includes three modules: a first module having the LP turbine and the LP shaft, a second module having the LP compressor and a journal secured to the rotor of the LP compressor, and a third module having an input shaft of the reduction gear, this input shaft having an upstream end for coupling to a sun gear of the reduction gear and a downstream end comprising first splines that are oriented parallel to the axis and are configured to be engaged in second complementary splines of said second module.
F02C 3/06 - 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
56.
ACOUSTIC TREATMENT COATING COMPRISING A FOLDED ORDERED METAPOROUS MATERIAL
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DU MANS (France)
Inventor
Boulvert, Jean
Cavalieri, Théo
Groby, Jean-Philippe
Ross, Annie
Fotsing Roland, Edith
Mardjono, Jacky Novi
Romero Garcia, Vicente
Gabard, Gwénaël
Costa Baptista, Josué
Abstract
An acoustic treatment coating includes a metamaterial, wherein the metamaterial includes a solid structure including at least one free space forming a folded cavity and at least one porous material placed in the free space of the solid structure.
A scoop for an aircraft turbine engine, this scoop including a body which is movable between at least two positions, including a first position in which it is configured to capture part of a first air stream flowing in a first direction, wherein the body is movable into a second position in which it is configured to capture part of a second air stream flowing in a second direction which is opposite to the first direction, and wherein the body is mounted to move freely between the at least two positions so as to automatically adopt the first position when the first air stream flows, and to automatically adopt the second position when the second air stream flows.
A method for diagnosing an engine control unit of an aircraft uses a device having an autonomous and mobile electronic diagnostic unit and connecting means for connecting the unit to an engine control unit of an aircraft. The connecting means advantageously includes a single connection cable.
G01R 31/00 - Arrangements for testing electric properties; Arrangements for locating electric faults; Arrangements for electrical testing characterised by what is being tested not provided for elsewhere
B64F 5/60 - Testing or inspecting aircraft components or systems
59.
AIRCRAFT DUAL-FLOW TURBINE ENGINE ASSEMBLY, THE ASSEMBLY BEING EQUIPPED WITH INCIDENCE-CONTROLLED AIR DISCHARGE FINS
An aircraft dual-flow turbine engine assembly includes: an internal shroud for externally delimiting a primary flow path of the turbine engine gases; an external shroud for internally delimiting a secondary flow path of the turbine engine gas; and at least one air discharge duct extending between the internal shroud and the external shroud, the air discharge duct opening into the secondary flow path through an outlet orifice equipped with discharge fins. At least some of the discharge fins are movably mounted so as to be able to be incidence-control between a propulsion position, and a reverse thrust 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
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 1/66 - Reversing fan flow using reversing fan blades
F02K 1/74 - Reversing at least one flow in relation to at least one other flow in a plural-flow engine
F01D 17/12 - Final actuators arranged in stator parts
60.
COMPOSITE VANE FOR AN AIRCRAFT TURBOMACHINE AND METHOD FOR THE MANUFACTURE THEREOF
A composite vane for a turbomachine, in particular an aircraft turbomachine, this vane including an aerofoil having a pressure side and a suction side connected together by a leading edge and by a trailing edge, the aerofoil been formed from a fibrous preform obtained by weaving fibres in three dimensions, which is embedded in a polymer matrix, the vane further including a first metal shield extending over and along the leading edge of the aerofoil, the vane further including at least one cover element extending over and along the trailing edge of the aerofoil.
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
B29C 70/46 - 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
61.
Stator support for a fan shaft driven by a reduction gearbox in a turbomachine
Stator structure (22) extending around an axis of the turbomachine and comprising: —a support (50) having an inner surface centred on the axis and—a flange (60) defining an air chamber (A2) and having an outer surface centred on the axis, the support (50) extending around the flange (60) such that the inner and outer surfaces are opposite to each other, the structure (22) defining an oil circuit and an air circuit which are formed by upstream channels (64, 65) and downstream channels (54, 55), —each upstream channel (64, 65) defining an outer opening in the outer surface, —each downstream channel (54, 55) defining an inner opening in the inner surface, each circuit being oriented between the outer and inner openings in a direction comprising a component radial to the axis.
A device and a method for diagnosing an engine control unit of an aircraft, the device including an autonomous and mobile electronic diagnostic unit and connecting means for connecting the unit to an engine control unit of an aircraft, the connecting means advantageously including a single connection cable.
G01R 31/00 - Arrangements for testing electric properties; Arrangements for locating electric faults; Arrangements for electrical testing characterised by what is being tested not provided for elsewhere
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
63.
Fuel supply circuit for a combustion chamber of a turbomachine
A fuel supply circuit including a supply duct in which a first flow (Fl) is able to circulate, at least one supply pump for circulating the fuel from a tank to a metering unit, and a recirculation duct for a second flow upstream of the pump, wherein the ducts (20, 60) discharge in different directions upstream of the pump; the circuit includes a dispensing device including an internal flow duct for the first flow (Fl), an external flow channel for the second flow and at least one passage orifice for the second flow, which communicates fluidically with the channel; and the at least one orifice (810) and the duct are coaxial and respectively discharge upstream of the pump.
A lubrication system for a turbomachine which includes a reduction gear. The lubrication system includes a main branch and a bypass branch which is arranged in parallel with the main branch. The main branch includes a heat exchanger. The bypass branch includes a hydraulic resistance that is lower than the hydraulic resistance of the main branch. The reduction gear is located downstream of the main branch and the bypass branch. The lubrication system includes at least one bypass device that is configured to supply the bypass branch with lubricant when a pressure value of the lubricant is below a threshold value.
A turbine engine assembly including a first rotor element and a second rotor element extending around a longitudinal axis X and coupled to one another by a gear coupling, the first rotor element including teeth and the second rotor element including complementary teeth which extend along the longitudinal axis and form the gear coupling. The first and second rotor elements each include a first and a second radial flange which are annular and arranged facing one another, and the assembly includes a sealing device configured to ensure that the gear coupling is sealed at the first and second flanges.
A connecting rod includes a hollow body with a ring at each end and a centre, a thickness wall e, the wall defining an outer perimeter pe and an inner perimeter pi, a surface section s being contained between the outer and inner perimeters, wherein the outer perimeter pe increases from the ends of the hollow body to the centre of the connecting rod, the hollow body maintaining a constant surface section s, the thickness e decreasing from the end to the centre of the connecting rod.
A method for repairing a part for an aircraft turbine engine, the part including a lower panel, an upper panel and a core having a honeycomb structure arranged between the lower panel and the upper panel, the part having an unimpaired portion and an at least partially impaired portion, the repair method including the following steps: (a) removing at least one portion of the lower panel or the upper panel from an area to be repaired; (b) removing at least one portion of the core from the area to be repaired; (c) reforming the core in the area to be repaired directly on the part by additive manufacturing; (d) reforming the lower panel or the upper panel in the area to be repaired, directly on the part.
B29C 64/118 - Processes of additive manufacturing using only liquids or viscous materials, e.g. depositing a continuous bead of viscous material using filamentary material being melted, e.g. fused deposition modelling [FDM]
B29C 64/393 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
68.
DEVICE FOR PRESSURIZING TURBOMACHINE DOWNSTREAM ENCLOSURE, AND CORRESPONDING TURBOMACHINE
A device for pressurizing a turbomachine downstream enclosure, the device being fed by means of a pressurizing air flow and including: a first shaft portion, a second shaft portion coupled to the first shaft portion via internal splines and external splines and extending, at a distance, inside the first shaft portion. The device includes at least a first passage made through the splines, and at least one orifice passing right through the second shaft portion, the orifice being arranged upstream of the first passage, in fluidic communication with the first passage, and opening out inside the second shaft portion so as to allow the pressurizing air flow to circulate from the first passage, between the first and second shaft portions, towards the downstream enclosure.
The present invention relates to a blade (3) of a fan (1), made of composite material, of a turbomachine. On a portion (15) of the blade (7) which extends from a lower limit (16) located at a predetermined distance (d) from the shank (6) of the blade equal to at least 80% of the predetermined height (h) to the top (11) of the blade (7), a dihedral angle (D) measured at at least one predefined point on the chord of the blade (3), located on the chord of the blade (3) at the upstream end (13a) of the nose (13) of the shield (12), is greater than or equal to −3° and less than or equal to 0°.
A device for distributing at least one fuel flow for a circuit for supplying fuel to a combustion chamber of a turbo machine, in particular of an aircraft, the device including a rotating main body extending about a longitudinal axis, the main body including an internal flow duct for a first flow of fuel, extending through the main body along the axis, a first annular portion configured to form an outlet passage for a second flow of fuel, and including at least one orifice for the second flow of fuel to pass through, and an external annular groove in fluidic communication with the at least one orifice in the first portion.
A blade comprising a blade body made of fiber-reinforced organic-matrix composite material and a leading edge shield made of a material having a better resistance to point impacts than the composite material of the blade body, the leading edge shield being assembled on the blade body and including a pressure side fin and an suction side fin connected by a thicker central part, the blade including an aerodynamic airfoil height and a chord length, the suction side fin having a first length projected onto the chord between 10 and 18% of the chord length, the first length being disposed between 70 and 80% of the aerodynamic airfoil height, a second length projected onto the chord between 18 and 26% of the chord length, the second length being disposed between 85 and 95% of the aerodynamic airfoil height.
An aircraft turbomachine module having two tubular casings equipped with first and second annular flanges assembled by studs received in the first flange and receiving nuts on the back of the second flange, each stud having an intermediate stretch of hexagonal shape cooperating with a locking plate. The locking plate can have a first orifice received on the intermediate stretch by at least two opposite walls, and a transverse body which is immobilized to prevent the rotation of the plate and that of the stud. Each plate can be fitted on two immediately adjacent studs and the body can have, for this purpose, a second orifice opposite the first orifice and which is received on an intermediate stretch of an immediately adjacent stud and cooperates with the intermediate stretch of hexagonal shape of the stud.
A propulsion system includes a drive shaft that rotates about an axis, a fan, a fan shaft that drives the fan about the axis, a reduction mechanism coupling the drive and fan shafts, a compression section driven by the drive shaft, and an inlet channel that extends between the fan and the compression section. The inlet channel includes inner and outer ferrules delimiting an air inlet flow path, the inner ferrule having a minimum radius. The reduction mechanism includes two reduction stages including two or more planet gears circumferentially distributed around the axis, each planet gear including a first portion meshed with the drive shaft and a second portion meshed with the fan shaft and being mounted fixed with respect to the inner ferrule and having a maximum radius greater than the minimum radius such that the air inlet flow path extends partly between the two or more planet gears.
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F02C 7/04 - Air intakes for gas-turbine plants or jet-propulsion plants
74.
TURBINE BLADE FOR AN AIRCRAFT TURBINE ENGINE, COMPRISING A PLATFORM PROVIDED WITH A CHANNEL FOR PRIMARY FLOW REJECTION TOWARDS A PURGE CAVITY
A turbine blade for a turbine engine, including an aerofoil and a platform. The platform includes an internal channel having a suction opening which opens on a first surface of an upstream portion of the platform, this first surface defining a primary duct. The internal channel includes an ejection opening which opens on a second surface of a downstream portion of the platform, this second surface defining a purge cavity. The internal channel makes it possible to suck in a part of a fluid circulating in the primary duct so as to reduce the intensity of secondary flows which result from friction of the fluid on the first surface.
A bypass turbomachine is disclosed. The turbomachine includes a longitudinal axis having an upstream blower impeller and a downstream air-flow straightening assembly of a secondary annular passage delimited radially on the inside by a radially inner shroud and radially on the outside by a radially outer shroud, blades extending between the radially inner and outer shrouds and being attached at a first end portion to the radially inner shroud and at a second end portion to the radially outer shroud, the blades including a useful portion extending between the first and second end portions and defining a lower face and an upper face. For each blade, in a plane perpendicular to the longitudinal axis, each aerodynamically useful portion is bent into a C-shape in the circumferential direction.
The invention relates to an assembly for a turbine engine blade, the assembly comprising a fastener (9) defining a well (10) for receiving a blade root (7), and a clamping device (32) for clamping the blade root (7) against the fastener (9) when the blade root (7) is received in the well (10), wherein the clamping device (32) comprises: at least one clamping member (34; 36) suitable for bearing simultaneously on the blade root (7) and on the fastener (9) in the well (10) so as to exert a clamping force on the blade root (7), the clamping member (34; 36) defining a threaded hole, and a resilient layer (44) arranged to bear in the well on the blade root, a rod (38) comprising a thread engaging with the threaded hole so that a rotation of the rod (38) with respect to the clamping member causes a variation in the clamping force exerted by the clamping member on the root.
The invention relates to a method for modelling the behaviour of a circular rolling mill (1) intended for rolling a cylindrical component on the basis of a setpoint, the circular rolling mill comprising at least one tapered roller (3) configured to effect a translational movement in a first direction (Y), and a mandrel (2), configured to effect a translational movement in a second direction (X), the setpoint comprising a setpoint for the rate of increase of an outside diameter of said cylindrical component as a function of said external diameter, and a setpoint for the height of the cylindrical component in the first direction as a function of a thickness of the cylindrical component in the second direction.
B21B 37/00 - Control devices or methods specially adapted for metal-rolling mills or the work produced thereby
B21B 1/18 - Metal rolling methods or mills for making semi-finished products of solid or profiled cross-section; Sequence of operations in milling trains; Layout of rolling-mill plant, e.g. grouping of stands; Succession of passes or of sectional pass alternations for rolling wire or material of like small cross-section in a continuous process
A turbojet engine, having a primary duct and a secondary duct surrounding the primary duct, a row of guide vanes extending in the secondary duct downstream of a fan, at least one servo extending in the secondary duct downstream of the guide vanes and housed in a profiled casing, and at least one air sampling scoop located in the region of the casing and equipped with a control valve. Where at least one guide vane located in the vicinity of the casing has a variable geometry zone, the geometry of which can be adjusted during operation.
F02C 7/042 - Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
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
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
79.
METHOD FOR ASSESSING THE SERVICE LIFE OF A TURBINE ENGINE PART
The invention describes a method for assessing the service life (DDV) of a turbine engine part, comprising the following steps: S1: determining average damage to the part over time (E_moy(t)) from stresses applied to the part on the basis of wear of the part (σ(u)) and a law of variation in the wear of the part over time (u(t)); S3: determining cumulative damage (E_cum) to the part corresponding to damage on breaking (E_rupt) of the part, the cumulative damage (E_cum) corresponding to the integral of the average damage over time (E_moy(t)) between an initial time (t_0) and an end time (t_rupt): E_cum=E_rupt=Formula (1); and S4: inferring the service life (DDV) of the part, the service life (DDV) corresponding to the end time (t_rupt).
An aircraft turbine engine with a main axis includes: an arm that is radial with respect to the main axis, extending through an air flow duct; a support wall belonging to the radial arm and which has two openings; a plurality of electric harnesses extending partially through the radial arm and for which a section of each harness is secured to the support wall of the arm; and two mounting blocks, in which each mounting block is associated with a group of sections of harness.
H02G 3/00 - Installations of electric cables or lines or protective tubing therefor in or on buildings, equivalent structures or vehicles
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
F16L 3/223 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets specially adapted for supporting a number of parallel pipes at intervals each support having one transverse base for supporting the pipes
81.
TURBOMACHINE MODULE PROVIDED WITH A PROPELLER AND OFFSET STATOR VANES
A turbomachine module having a longitudinal axis, having a shroudless propeller which is rotated about the longitudinal axis by a drive shaft, which is connected at least to a compressor rotor, and at least one distributor comprising a plurality of stator vanes which extend along a radial axis which is perpendicular to the longitudinal axis Z from a fixed casing, the distributor being arranged downstream of the propeller. The fixed casing can be an inter-compressor casing which is arranged downstream of a low-pressure compressor, along the longitudinal axis, the inter-compressor casing having a ring which has a longitudinal axis and which is provided with sleeves intended to support the stator vanes, the inter-compressor casing and the ring being monobloc.
A system for real-time control of the trajectory of an aircraft on a runway, includes at least one steering system configured to steer the aircraft on the ground, each steering system being associated with at least one use parameter; a computer configured for determining, from aircraft data and external data including runway condition data and ground meteorological data, each steering system intended to steer the aircraft according to a predetermined trajectory and each corresponding use parameter; and a control system configured to control each steering system determined according to each corresponding determined use parameter.
A blade made of composite material includes a fibrous reinforcement and an at least partially ceramic matrix, wherein the fibrous reinforcement includes a first structural portion formed by three-dimensional or multilayer weaving, and a second functional portion, distinct from the first portion, including short fibers.
A mechanical part includes a part body provided with a through-opening, and a through-fastening element including a rod engaged in the through-opening and connected to an inner surface of the through-opening by a support fastened to the rod of the through-fastening element and to the inner surface of the through-opening and configured to break the link between the rod of the through-fastening element and the inner surface of the through-opening, under the effect of a thrust applied on the through-fastening element so as to push the rod of the latter into the through-opening. Due to the fact that the through-fastening element is pre-integrated with the mechanical part, the logistical management prior to assembly of the mechanical part to another mechanical part is simplified, the risk of operator errors and loss of the fastening element are reduced.
A turbine engine module of longitudinal axis X, including an unducted propeller intended to be rotated around the longitudinal axis X by a power shaft which is connected at least to a rotor member, at least one flow straightener including a plurality of stator vanes (28)-extending along a radial axis Z, at least one first casing mounted upstream, along the longitudinal axis, of the rotor member and a second casing mounted downstream, along the longitudinal axis, of the rotor member. The stator vanes each include a root housed in a sleeve which is connected to the first casing and to the second casing
Methods, apparatus, systems and articles of manufacture are disclosed. An apparatus for mounting a gas turbine engine to a pylon, the gas turbine including an upstream section and a downstream section, the gas turbine defining a roll axis, a yaw axis, and a pitch axis, the apparatus including: a first mount to couple the upstream section of the gas turbine engine to the pylon; a second mount to couple the upstream section of the gas turbine engine to the pylon, the second mount downstream of the first mount; a thrust linkage to couple the upstream section to the pylon, wherein the downstream section is decouplable from the upstream section without decoupling the first mount, the second mount, and the thrust linkage.
The present invention relates to a fan blade of a turbomachine comprising a structure made of composite material comprising a fibrous reinforcement obtained by three-dimensional weaving of strands and a matrix in which the fibrous reinforcement is embedded, which comprises a first portion forming the suction-side wall and a second portion forming the pressure-side wall, the strands of the fibrous reinforcement comprising first strands comprising carbon or aramid fibres, the Young's modulus of which is greater than 250 GPa, and second strands produced from a viscoelastic material and having a Young's modulus less than 10 GPa, the first portion comprises uniquely first strands whereas the second portion comprises second strands.
An actuating device of a pump of a fuel pumping system of an engine, including a motor, a generator, an inverter, a switching member and a control member, the motor including a first rotor coupled to the pump and a first stator including at least one input stator winding, the generator including a second rotor coupled to a drive shaft of the engine, and a second stator including at least one output stator winding, the control member being configured to control the switching member in order to selectively connect each input stator winding: to a corresponding output stator winding if a speed of the engine is higher than or equal to a predetermined speed; to a corresponding output of the inverter, otherwise.
H02P 9/42 - Arrangements for controlling electric generators for the purpose of obtaining a desired output to obtain desired frequency without varying speed of the generator
An unducted fan having variable pitch blades for an aircraft or wind propulsion system is described. The fan includes a hub provided with housings pivotally receiving vanes about radial axes, and an outer casing of the hub into which circular openings of the housings open. Each vane includes a root which is rotatably mounted in the associated housing, a radially oriented blade and a disc-shaped platform extending from the blade to an edge of the circular opening. The blade includes at least one skin which includes a free lower edge facing the side of the root, wherein the platform is attached to the free lower edge of the skin.
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]
An assembly for an aircraft turbine engine, includes a fan, a speed reduction gear and an intermediate casing including an internal annular structure that houses an epicyclic gear train equipped with a planet carrier. The assembly includes a secondary retention device for the ring gear of the planet carrier, the device including: first projections that project radially inwards from the internal annular structure, each first projection having a recess that is open radially inwards, axially downwards and circumferentially in a first direction; and second projections that project radially outwards from the ring gear of the planet carrier and cooperate with the first projections in such a way that each second projection is partially accommodated in the recess of the first projection with which it is associated and downstream therefrom.
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
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
91.
ADDITIVE MANUFACTURING ON A TURBINE ENGINE PART BLANK INCLUDING A ROUGH AREA
A production of a structure for a turbomachine by additive manufacturing, including the following steps: a) providing an element having at least one localised rough region, in particular produced by machining, b) using powder bed fusion additive manufacturing in order to deposit a layer of material resulting from the fusion on the element.
B33Y 30/00 - ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING - Details thereof or accessories therefor
B22F 10/47 - Structures for supporting workpieces or articles during manufacture and removed afterwards characterised by structural features
An aircraft turbine engine including a gas generator, a fan, and an electrical machine, the stator of the electrical machine being connected to a power electronic circuit by at least one electrically conductive bar which extends substantially radially relative to the axis inside an IGV which forms part of a bladed angular segment, the bar being configured to be disassembled and removed from the turbine engine by disassembling and removing the segment.
INSTITUT SUPERIEUR DE L AERONAUTIQUE ET DE L ESPACE (France)
SAFRAN AIRCRAFT ENGINES (France)
Inventor
Tantot, Nicolas Jérôme Jean
Carbonneau, Xavier
Chanez, Philippe Gérard
Garcia Rosa, Nicolas
Martin Pleguezuelo, Alejandro
Sirvin, Nicolas Joseph
Abstract
The invention relates to an assembly for a turbomachine, comprising: —a nacelle comprising an inlet lip which defines an air inlet, and—a device for modifying the geometry of the air inlet, comprising: —a first spout, and—a second spout, the first spout and the second spout being translatably movable in relation the nacelle between: —a first configuration, in which the first spout forms the inlet lip and the second spout extends inside the nacelle, and—a second configuration, in which the first spout extends away from the inlet lip, and the second spout forms the inlet lip so as to define an air flow channel between a downstream surface of the first spout and an upstream surface of the second spout.
A mechanical reducer for a turbomachine for an aircraft, this reducer having a sun gear with internal splines and an external toothing, a ring gear which extends around the sun gear and which has an internal toothing, planet gears which are arranged between the sun gear and the ring gear and which each have at least one external toothing meshing with the external toothing of the sun gear and the internal toothing of the ring gear, and a planet carrier which carries first bearings for guiding the planet gears in rotation. The reducer can further include at least one second bearing for guiding the sun gear in rotation. The at least one second bearing can be mounted between the sun gear and the planet carrier.
A system for controlling the pitch of a propeller blade for an aircraft turbine engine is provided. The system generally includes a cup having an annular wall extending about a pitch setting axis of the blade, the annular wall having a lower axial end enclosed by a bottom wall, and an upper axial end that is open and configured so as to enable a root of the blade to be mounted in the cup; and a locking ring that extends around the pitch setting axis and is configured to be mounted around the root, the locking ring being a double dog ring that comprises two annular rows of outer dogs constituting active and safety dogs, respectively.
The present document relates to an annular assembly for a turbomachine turbine, in particular of an aircraft, said annular assembly extending about a longitudinal axis X and comprising: a distributor fixed to an external casing; a bladed disc mounted so as to rotate inside the external casing; said bladed disc being surrounded by a circumferentially segmented ring carried by the external casing and formed by a crown arranged radially on the outside of the bladed disc and by an annular deflector that is carried by an upstream edge of the crown and extends radially towards the inside from said upstream edge of the crown; characterized in that it comprises sealing members (56) between two circumferentially adjacent ring segments, these sealing members (56) comprising first circumferential sealing means (58) between two circumferentially consecutive deflector segments.
A turbomachine part of an annular assembly includes a structural body and a connecting edge integral with the structural body. Each connecting edge includes at least one groove intended to receive a sealing tab. Each connecting edge is made of composite material including a fibrous reinforcement consisting of randomly-oriented short fibres, the reinforcement being densified by a ceramic matrix.
Method for producing a casing made of a composite material for an aircraft turbine engine, the casing having an annular shape and including an annular fibrous preform formed by winding a fibrous cloth soaked in a polymer matrix, the method including the steps of: a) positioning an edge of the cloth on a molding drum, this cloth having a generally elongate shape and the edge being located at a first longitudinal end of the cloth, b) winding the cloth on the drum, over at least one rotation, so as to obtain the preform wound around the drum, c) molding of the preform and hardening of the resin which is injected into the preform or pre-impregnated on the preform, wherein the method includes, before step a), a step i) of mounting a removable stop on the drum, and in that step a) includes abutting the edge against this stop.
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
B29C 70/32 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core on a rotating mould, former or core
99.
METHOD FOR MANUFACTURING A COMPOSITE PART FOR A TURBOMACHINE
A method for manufacturing a part made of composite material for a turbomachine, in particular of an aircraft, includes the steps of: b) arranging a preform made of fibers in a mold, c) injecting polymerizable resin into the mold, d) machining the part, and e) visually checking the part. Step b) is preceded by a step a) in which at least one compliance coating is deposited in the mold. The compliance coating has a calibrated thickness (X) and at least one visual aspect identifiable by an operator. The coating is configured to cover at least one area of the preform and to be rigidly attached thereto by the resin at the end of step c) Step e) includes verifying, by the operator, the presence of the aspect in the area.
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
B29C 37/00 - Component parts, details, accessories or auxiliary operations, not covered by group or
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
100.
Turbine nozzle having blading made of ceramic matrix composite through which a metal ventilation circuit passes
A nozzle for a turbomachine turbine. The nozzle includes blades made of ceramic matrix composite material, and at least one metal liner passing through a respective blade. The liner is connected in a sealed manner to a metal internal shroud so as to guide a ventilation fluid through the blade. The nozzle can be configured to maintain effective sealing in spite of the different thermal expansions of the blade and of the internal shroud.
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