A mould for manufacturing a turbomachine fan casing made of a composite material with fibrous reinforcement that is densified by a matrix, includes an impregnation mandrel around which a fibrous preform is to be wound and angular counter-mould sectors assembled around the external contour of the impregnation mandrel, which are intended to close the mould. The impregnation mandrel includes a main body with an annular shape, and an annular barrel which is arranged around the main body and around which the fibrous preform is to be wound, the barrel and the angular sectors defining a moulding cavity intended to receive the fibrous preform. The main body and the angular sectors are made of a first material having a first thermal expansion coefficient, the barrel being made of a second material that has a second thermal expansion coefficient, the second thermal expansion coefficient being greater than the first thermal expansion coefficient.
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 45/14 - Injection moulding, i.e. forcing the required volume of moulding material through a nozzle into a closed mould; Apparatus therefor incorporating preformed parts or layers, e.g. injection moulding around inserts or for coating articles
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
A system for controlling the pitch of a propeller blade for an aircraft turbine engine, including a cup having an annular wall extending about an axis intended to be a pitch setting axis of the blade, and a bottom wall configured so as to cooperate in a form-fitting manner with a free end of the root so that the cup is prevented from rotating relative to the root about the axis, and a locking ring that is configured so as to be mounted within the cup and to cooperate respectively with the root and the annular wall of the cup in order to ensure axial retention of the root in the cup.
An assembly comprising a propeller blade and its system for angularly adjusting the pitch of the blade including a bowl which is radially delimited by an annular wall (44)-extending about an axis for adjusting the pitch of the blade, the bowl having a bottom wall, a free lower end of the root being fitted axially into a complementary housing of the bottom wall in order to rotatably connect the bowl and the blade about the pitch-adjustment axis. The root of the blade comprises a first limiting face that engages with a first abutment face of the bowl to limit the rotation of the blade in the event of breakage of the lower end of the root.
B64C 27/78 - Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement in association with pitch adjustment of blades of anti-torque rotor
F01D 7/00 - Rotors with blades adjustable in operation; Control thereof
An assembly including a propeller blade and a pitch adjustment system for setting the pitch of the blade. A root of the blade has a bulb and the pitch adjustment system for setting the pitch of the blade includes a cup having an upper opening; a retaining ring, which extends around the bulb, is restricted at least in terms of axial movement towards the opening relative to the cup and has a bearing face to inhibit axial movement of the root towards the opening The pitch adjustment system being further includes a lower seat for the root to engage the cup, the seat and/or the retaining ring being mounted so as to be capable of axial translation relative to the cup by at least one clamping mechanism.
A rolling bearing assembly includes an inner ring and an outer ring that are coaxial about an axis of rotation and between which rolling elements are arranged. An annular metal rolling bearing cage is centered about the axis of rotation and arranged radially between the inner ring and the outer ring, the annular cage having recesses for retaining rolling elements, an outer annular face facing the outer ring, and an inner annular face facing the inner ring. At least one insert at least one tubular portion engaged in a recess of the annular cage, a tubular portion of the insert being connected at one radial end to a rim arranged in abutment against the inner annular face or the outer annular face of the annular cage. A latching means on the insert cooperates with a latching means in the annular cage to hold the insert in a recess.
Fuel pump for an aircraft engine, comprising an inducer and a centrifugal impeller fixed together and having an axis of rotation, an annular space spacing axially the inducer and the impeller, and two contact portions between the inducer and the impeller disposed radially outside the annular space, a first plenum chamber and a second plenum chamber, in which the inducer and the impeller are spaced axially from each other, each being disposed between the two contact portions, the plenum chambers being symmetrical to each other with respect to the axis of rotation and in fluid communication with the annular space, the centrifugal impeller comprising a plurality of axial balancing holes distributed about the axis of rotation and opening out into the annular space at one end, and into a downstream space of the impeller at the other end.
A turbine engine module having longitudinal axis X including an unducted propeller for rotating about the longitudinal axis X; —at least one flow straightener wiht a plurality of stator vanes extends substantially along a radial axis Z, each stator vane having a root and a blade rising radially from the root; and—a pitch change system for changing the pitch of the stator vanes. At least two adjacent stator vanes are connected to each other by at least one retaining member coupled to the blades of the stator vanes by at least one pivot shaft and mounted radially from the root of the stator vanes. The pivot shaft extends along a pivot axis B coaxial with the pitch adjustment axis A to enable the stator vanes to pivot about the pivot axis B. An anti-vibration unit dampens vibration of the pivot shaft.
A propulsion system includes a drive shaft movable in rotation about an axis of rotation, a low-pressure compressor driven in rotation by the drive shaft, the low-pressure compressor having a mean radius, a fan shaft, a fan driven in rotation by the fan shaft, a reduction mechanism coupling the drive shaft and the fan shaft, having a reduction ratio, and an inlet channel which extends between the fan and the low-pressure compressor, the inlet channel having an inlet adjacent to the fan and an outlet opposite the inlet and adjacent to the low-pressure compressor, the inlet having a mean radius. A first ratio between a ratio of a mean radius of the inlet channel and the mean radius of the low-pressure compressor, and the reduction ratio of the reduction mechanism, is strictly less than 0.35.
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
9.
CONFLUENCE STRUCTURE OF A PRIMARY STREAM AND A SECONDARY STREAM IN A BYPASS TURBINE ENGINE
A confluence structure of an aircraft bypass turbine engine which includes a confluence plate with a downstream end supported by a portion that is movable in the direction of the axis by a control mechanism which can optionally be adjusted in flight. A mobile portion of a sleeve delimiting the secondary stream on the outside, and an inner projection of the outer casing can also slide axially in certain embodiments. This provides a wide range of options for modifying the gas dilution and operating conditions of the engine.
F02K 3/075 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type controlling flow ratio between flows
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F02K 3/10 - Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof by after-burners
A turbomachine blade includes a blading made of composite material with a fibrous reinforcement densified by a matrix and a metal leading edge formed by a metal foil, the foil having an intrados fin and an extrados fin which extend respectively over intrados and extrados faces of the blading by conforming to an airfoil of the blade, wherein the blade also includes at least one unidirectional fabric ply made of composite material on the leading edge between the blading and the metal foil, each unidirectional fabric ply extending at least partially over the intrados and extrados faces of the blading.
A system for controlling the pitch setting of a propeller blade for an aircraft turbine engine is provided. The system generally including a blade having an airfoil connected to a root; a cup having an annular wall extending about a pitch axis of the blade and a lower axial end enclosed by a bottom wall; a locking ring that extends around the root and inside the cup; a safety element that ensures the retention of the root relative to the cup. The at least one safety element can have a generally elongate shape and can pass through aligned holes in the bottom wall of the cup and in the free end of the root.
Annular combustion chamber for an aircraft turbomachine, said chamber having two coaxial annular walls, an inner annular wall and an outer annular wall, respectively, which are connected upstream by an annular bottom wall of the chamber, wherein an injection device passes through an axis and comprises an air injection system and a frustoconical bowl which is flared downstream and has air passage openings, the chamber further having an annular deflector placed downstream of the bottom wall substantially parallel to the latter; and wherein the air injection system, the bottom wall, the deflector and the bowl are integrally formed.
The turbine (12) for a turbine engine (100) extends around a major axis (X-X) and comprises: - a casing (4) comprising an annular hook (38), - a movably mounted impeller (16), - a ring (30) extending opposite the impeller in a direction radial to the major axis (XX), - a distributor (14) comprising a blade provided with a platform (11), the platform being extended radially outwards by a spoiler (22, 24), the spoiler (22, 24) being radially mounted on the hook (38), and - foils (40) each having a profiled trough shape in a direction circumferential to the axis, the foils extending in succession in the circumferential direction, each foil (40) extending between the spoiler (22, 24) and the hook (38). The turbine comprises stops (162) to prevent the foils (40) from moving in the circumferential direction. Each foil (40) comprises a protrusion (162) arranged so as to extend circumferentially opposite another protrusion (162) of another adjacent foil (40).
This device for disconnecting and braking a rotor of an electric machine comprises an inner axis driven in rotation by the rotor and longitudinally movable between an engagement position and a disconnection position, the disconnection and braking device further comprising a longitudinally fixed nut and a means for braking the rotation of the nut so that braking of the rotation of the nut urges the inner axis into its disconnection position and brakes the rotation of the rotor.
F16D 65/18 - Actuating mechanisms for brakes; Means for initiating operation at a predetermined position arranged in or on the brake adapted for drawing members together
The manufacturing method comprises the following steps: —extracting a sub-batch (18) from a batch (16) of powder, the batch having a ratio X0 of a predetermined component, —manufacturing parts (22) using powder-bed based additive manufacturing until all of the sub-batch is used up (18); then—performing the following cycle a number nmax of times: —recycling the powder and continuing manufacture until all of the recycled powder is used up; —determining a contact surface Sn between the powder and a material fused into the parts (22), n designating the cycle number, and a mass Mn of accumulated powder used since the start of the method, and—measuring a ratio Xnmax of a predetermined component in at least one of the parts (22) or a test piece (24) manufactured during the last cycle, then—determining a quantity R such that: Formula (I); —performing the following cycle at least once: —recycling the powder and continuing manufacture until all of the recycled powder has been used up at least once, then—determining the contact surface Sn, and the mass Mn of accumulated powder, and—calculating, by means of Sn, Mn and R, a ratio Xn of the component in one of the parts (22) manufactured during the cycle.
B33Y 50/02 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
B33Y 40/00 - Auxiliary operations or equipment, e.g. for material handling
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 12/90 - Means for process control, e.g. cameras or sensors
16.
METHOD FOR SUPPLYING VALUES OF PARAMETERS OF A HEAT SOURCE INTENDED TO CREATE A WELDED SEAM BETWEEN TWO PLATES, CORRESPONDING COMPUTER PROGRAM AND CORRESPONDING DEVICE
This method comprises: receipt of a desired value of at least one spatial characteristic of the welded seam; determination of several samples of the parameters of the heat source; for each sample, determination of a value for each spatial characteristic of the welded seam for this sample; several successive iterations of the following steps: the determination of extrapolated points, from simulated points, determination of a target point, of the function, at which each spatial characteristic of the welded seam exhibits a value close to the desired value, and determination of a value for each spatial characteristic of the welded seam from the values of the parameters of the heat source for the target point, so as to obtain a new simulated point; and the supply of the values of the parameters of the heat source of the target point obtained in the last iteration.
B23K 9/095 - Monitoring or automatic control of welding parameters
B23K 31/12 - Processes relevant to this subclass, specially adapted for particular articles or purposes, but not covered by any single one of main groups relating to investigating the properties, e.g. the weldability, of materials
G06F 30/20 - Design optimisation, verification or simulation
A tooling for injecting a polymer resin into a fibrous preform for the manufacture of a revolution part in composite material including a barrel shape with an inside diameter of smaller diameter delimiting the revolution part into an upstream portion and a downstream portion of the inside diameter, the upstream portion including a back-draft intermediate portion, injection tooling wherein, to allow a demolding of the barrel-shaped part once the injection and the polymerization of the polymer resin are carried out, the injection tooling includes on the one hand a frustoconical drum comprising a first drum portion in direct contact with an inner surface of the portion of the revolution part downstream of the inside diameter and a second drum portion and on the other hand a segmented insert whose outer surface matches an inner surface of the portion of the revolution part upstream of the inside diameter.
B29C 33/42 - 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 characterised by the shape of the moulding surface, e.g. ribs or grooves
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 33/48 - 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 with means for collapsing or disassembling
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
Architecture for collecting and transmitting an aircraft flight data report including at least one data acquisition unit installed on board the aircraft, means for transmitting the flight data report installed on board the aircraft and a station for collecting the data report on the ground, including means for calculating and taking into account a cost index for optimizing a choice of a best means for transmitting the flight data report and its content with respect to the cost index.
An annular intermediate casing for a turbomachine through which an aerodynamic airstream circulates, including a radially internal shroud (31), a radially external shroud, at least one stator blade including a vane with a leading edge, and at least one aero-dynamic member which is formed of a structural arm extending radially at least in part between the radially internal shroud and the radially external shroud. The aerodynamic member includes a leading edge which is aligned with the leading edge of the stator blade in one plane and which is borne by a profiled portion positioned upstream of the structural arm in the direction of circulation of the aerodynamic airstream, the profiled portion, the structural arm, the radially internal shroud and the radially external shroud being monobloc.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE NORMALE SUPERIEURE PARIS- SACLAY (France)
Inventor
Schneider-Die-Gross, Julien Paul
Belhaj Saad, Intisar
Mendoza Quispe, Arturo
Parra, Estelle Marie Laure
Roux, Stéphane
Abstract
The invention relates to a method for characterizing, from a volume image, a fibrous structure having a three-dimensional weaving between a plurality of warp yarns extending along a first direction and a plurality of weft yarns extending along a second direction perpendicular to the first one, the method comprising: a first processing (E10) of the volume image by filtering along a third direction perpendicular to the first and second directions so as to attenuate the periodic patterns along the third direction, obtaining (E20) a two-dimensional image corresponding to an intermediate plane along the third direction of the filtered volume image, a second processing (E31, E41) of the two-dimensional image by filtering along the first or second direction so as to attenuate the periodic patterns, obtaining (E32, E33) a one-dimensional profile representing the positions of warp or weft columns and corresponding to an intermediate line along the first or second direction of the filtered two-dimensional image, and comparing (E33, E43) the one-dimensional profile with a reference profile.
An intermediate casing includes an inner wall having an outer surface extending along a longitudinal axis, the outer surface having a first profile in a first plane, an outer wall having an inner surface a second profile in the first plane, and first through fourth arms each extending radially. The outer and inner surfaces and the first and second arms delimit a first cavity, the first cavity having a first area and the outer and inner surfaces being separated by a first distance. The outer and inner surfaces and the third and fourth arms delimit a second cavity, the second cavity having a second area and the outer and inner surfaces being separated by a second distance. The first and second profiles are such that the first area is substantially identical to the second area and the first distance is different from the second distance.
A method for manufacturing a composite part reinforced with nanotubes, includes stacking a plurality of composite plies of prepreg and at least one composite ply integrating nanotubes, the at least one composite ply integrating nanotubes being positioned in an inter-ply space between two composite plies of prepreg, wherein the at least one composite ply integrating nanotubes is a ply having a thermoplastic component, the nanotubes being integrated in the thermoplastic component.
B29C 70/02 - Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising combinations of reinforcements and fillers incorporated in matrix material, forming one or more layers, with or without non-reinforced or non-filled layers
B29C 70/18 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length in the form of a mat, e.g. sheet moulding compound [SMC]
B29C 70/34 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression
B29C 70/50 - Shaping or impregnating by compression for producing articles of indefinite length, e.g. prepregs, sheet moulding compounds [SMC] or cross moulding compounds [XMC]
B29C 70/88 - Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts characterised primarily by possessing specific properties, e.g. electrically conductive or locally reinforced
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
23.
PLATFORM FOR A FAN ROTOR OF AN AIRCRAFT TURBOMACHINE
Platform for an aircraft turbo machine fan rotor, the platform being configured to be secured to a fan disc between two adjacent fan blades. The platform further including a longitudinal wall defining an aerodynamic external face. The wall includes a honeycomb structure interposed between two skins which are respectively an internal skin and an external skin, with the external skin defining the aerodynamic external face.
A device for filtering a flow of cooling air for cooling a low-pressure turbine of a turbomachine, includes a duct having a geometry configured to centrifuge the flow of cooling air passing through the duct, the duct having openings dimensioned to enable a separation of the solid particles contained in the flow of cooling air being centrifuged.
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
B01D 45/16 - Separating dispersed particles from gases or vapours by gravity, inertia, or centrifugal forces by centrifugal forces generated by the winding course of the gas stream
B01D 46/24 - Particle separators, e.g. dust precipitators, using rigid hollow filter bodies
B01D 46/48 - Removing dust other than cleaning filters
B01D 50/20 - Combinations of devices covered by groups and
25.
IMPROVED CMC GUIDE VANE FOR A TURBOMACHINE TURBINE
A turbomachine turbine nozzle extending around a central axis, including at least one radially outer shroud, at least one radially inner shroud, and at least one blade made of ceramic matrix composite material, distinct from the radially inner shroud and from the radially outer shroud, and extending radially between the radially inner shroud and the radially outer shroud, the blade being hollow and including a cavity opening at a radially inner end and at a radially outer end of the blade, the nozzle including at least one tubular mast arranged in the cavity of the blade and allowing routing the ventilation air passing through the cavity of the blade, the mast including a radially outer end attached to the radially outer shroud, and a radially inner end cooperating with a radial flange for positioning the radially inner shroud.
The invention proposes an aeronautical turbine engine assembly comprising an upstream casing (55) to which guide blading (48a) is fastened, and a downstream casing (58) to which a sealing element (62) provided with an abradable material for rotor blading is fastened. This assembly further comprises a shroud ring (66) placed between the upstream casing and the downstream casing and fastening means (68) for detachably fastening the shroud ring. In order to be fastened to the upstream casing, the guide blading (48a) of the turbine engine is mounted on a downstream hook (480b) of the upstream casing, without being hooked onto the shroud ring (66), and the downstream casing (58) has an upstream hook with which the sealing element (62) is engaged in order to be fastened to the downstream casing, or the shroud ring has an upstream hook on which the sealing element (62) is mounted so as to be fastened to the downstream casing.
F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
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 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
27.
TURBINE ENGINE WITH A FLOW SPLITTER HAVING A PROFILE WITH INCLINED SERRATIONS
An airflow profile structure having a leading and/or trailing edge profiled with a serrated profile having a succession of teeth and depressions. Along the leading and/or trailing edge, from a first location to a second location, the teeth of the serrated profile are individually inclined towards the second location.
A method for treating a surface of a metallic part in particular of a turbomachine includes a step of shot-peening a surface of the metallic part using metallic beads and a step of cleaning the surface of the metallic part after the shot peening. The cleaning comprises step includes applying a magnetic cleaning paste to the surface of the metallic part in such a way as to remove any residues of metallic beads.
A turbine nozzle for a turbine engine, including two annular walls extending about the same axis, the walls being connected to each other by blades having an aerodynamic profile, the nozzle being divided into sectors and including several nozzle sectors arranged circumferentially end-to-end about a longitudinal axis. The blades include first blades made from a first metal alloy, and second blades made from a second metal alloy different from the first alloy, each nozzle sector including first blades and second blades. The invention also relates to a turbine engine, in particular of an aircraft, including at least one such nozzle. The invention also relates to a method for manufacturing such a nozzle in which each of the sectors is produced, for example, by additive manufacturing.
Counter-rotating turbine of a turbomachine extending about an axis of rotation and comprising an inner rotor having 5 at least one inner moving blade supported in rotation by a first shaft, an outer rotor rotating in an opposite direction to the inner rotor, and comprising at least one outer moving blade supported in rotation by a second shaft coaxial with the first shaft, the first 10 and second shafts extending axially from upstream to downstream of the turbine, the first shaft being guided in rotation by a first bearing disposed between the first shaft and an upstream casing of the turbine, and the second shaft being guided in rotation by a second bearing disposed between the second 15 shaft and said upstream casing of the turbine, the first bearing and the second bearing being disposed upstream of a first stage of the inner rotor.
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
F01D 1/24 - Non-positive-displacement machines or engines, e.g. steam turbines characterised by counter-rotating rotors subjected to same working-fluid stream without intermediate stator blades or the like
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
F02K 3/072 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with counter-rotating rotors
31.
FIBROUS TEXTURE FOR A CASING MADE OF COMPOSITE MATERIAL WITH HYBRID WARP STRANDS
A fibrous texture has the shape of a strip extending in a longitudinal direction over a determined length between a proximal portion and a distal portion and in a lateral direction over a determined width between a first lateral edge and a second lateral edge. The fibrous texture has a three-dimensional or multi-layer weaving between a plurality of layers of warp strands extending in the longitudinal direction and a plurality of layers of weft strands extending in the lateral direction, the fibrous texture including first and second longitudinal sections extending over a width from the first or second lateral edge smaller than the determined width of the fibrous texture along the lateral direction. The first and second longitudinal sections each include warp strands and weft strands constituted by carbon fibers. The fibrous texture further includes a third section present between the first and second sections.
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
D03D 25/00 - Woven fabrics not otherwise provided for
A method for checking the tightening tension of at least one fastening bolt for fastening at least two aeronautical parts, wherein it includes the steps consisting in mounting a centring device on a head of a screw of the bolt, the centring device including an aperture aligned with an axis of the screw; engaging one end of a sensor in the aperture of the device until this end is in contact with the head of the screw; checking at least one physical parameter of the bolt by the sensor, this parameter being representative of the tightening tension of the bolt; then removing the centring device.
G01L 5/24 - Apparatus for, or methods of, measuring force, work, mechanical power, or torque, specially adapted for specific purposes for determining value of torque or twisting moment for tightening a nut or other member which is similarly stressed
B64F 5/60 - Testing or inspecting aircraft components or systems
33.
VANE MADE OF COMPOSITE MATERIAL HAVING A THREE-DIMENSIONAL WOVEN FIBROUS REINFORCEMENT AND TWO-DIMENSIONAL WOVEN SKIN AND METHOD FOR MANUFACTURING SAME
A blade for an aircraft gas turbine engine includes, in a longitudinal direction, a blade root, a shank and an aerofoil body, the aerofoil body extending in the longitudinal direction between the shank and a blade tip and in a transverse direction between a leading edge made of metal material and a trailing edge. The blade includes a blade core made of composite material having a three-dimensional woven fibrous reinforcement forming the blade root, the shank and a part of the aerofoil body. The blade also includes a skin made of composite material having a two-dimensional woven fibrous reinforcement surrounding the aerofoil body part of the blade core, the skin being interposed between the leading edge made of metal material and a front edge of the aerofoil body part of the blade core to define a thinned leading edge portion, the skin including one or more two-dimensional woven plies.
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/34 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression
34.
AIRCRAFT TURBOMACHINE COMPRISING A DEVICE FOR LUBRICATING A BEARING
The invention relates to a turbomachine for an aircraft, comprising:
a first rotor comprising a first shaft,
a second rotor comprising a second shaft,
a mechanical reduction gearing having an epicyclic gear set comprising a sun gear connected to the second shaft, a ring gear connected to the first shaft, and planet gears located between the sun gear and the ring gear and borne by a planet carrier attached to a stator of the turbomachine,
rolling element bearings for guiding said first shaft and second shaft in rotation,
an annular gutter which extends around the ring gear of the reduction gearing and which is configured to recover oil for lubricating the reduction gearing that is sprayed by centrifugal action out from the ring gear during operation, and
an annular bearing support which is attached, with the gutter, to a stator of the turbomachine and which supports at least one of said bearings, characterized in that it also comprises:
at least one device for conveying oil recovered by said gutter, which device is borne by said annular support and extends as far as said at least one bearing in order to lubricate the latter.
A method for producing an aerodynamic element for an aircraft, including a wall which is covered, at least partially, by a printed film including a plurality of ribs and/or grooves, the method including: providing a raw film made of a deformable material which is devoid of grooves and ribs; providing a pressure plate which includes a face provided with ribs and/or grooves complementary to the ribs and/or grooves of the printed film; placing said raw film on the wall of the element; positioning the pressure plate on the raw film, the printed side of the pressure plate facing the raw film; and a forming step during which the raw film is bonded with the wall of the component and during which the raw film is shaped by cooperation with the pressure plate to obtain the printed film including the ribs and/or the grooves.
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
A turbine for a turbine engine extending along an axis includes an annular casing and at least one turbine stage having a nozzle and a rotor impeller wheel surrounded by a sealing ring with an abradable element. The impeller wheel and the sealing ring are located downstream of the nozzle, and the sealing ring has an upstream end held on the casing by locking means. The turbine includes elastic sealing means in contact with the locking means as well as with the nozzle or the casing so as to press the locking means against the sealing ring. The locking means includes a radially outer portion with a C-shaped cross-section and a radial portion extending radially inwards from the outer portion. The sealing means include two elastic seals bearing on the radial portion, respectively on either side of the radial portion.
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 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
The invention relates to an exhaust case (15) for a turbomachine extending along a longitudinal axis (X), comprising: an annular shroud (23) having a wall (24) extending along the longitudinal axis (X) from a first flange (25), a plurality of openings (27) being provided through the wall (24); a plurality of mouths (28) each forming a channel (29) extending upstream to downstream between a respective inlet (30) and one of the openings (27), each mouth (28) having: a docking flange (31) at the inlet (30) having a radially inner edge (32) which is in contact with the first flange (25) of the annular shroud (23), a mouth wall (34) delimiting the channel (29) and comprising a radially inner wall portion (35) which is formed by a thickened section made on the wall (24) of the shroud (23).
F02K 1/28 - 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
F01D 17/12 - Final actuators arranged in stator parts
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
A vane of a turbine engine blade includes a first portion of structural resistance formed by two end portions including the leading and trailing edges and end strips of the lower surface and the upper surface, and of a core joining them. Two other portions of the blade are constructed of light material, composite for example, between the end portions to reconstitute the complete vane. The core has an oblique or diagonal extension between the end portions.
A turbomachine sealing ring has an axis of revolution and includes an annular support , an annular coating made of abradable material which is carried by the support, and an annular thermal protection plate which is carried by the support. The ring is divided into sectors and has a plurality of ring sectors disposed circumferentially next to one another about the axis. Each ring sector having has a support sector, a coating sector, and a plate sector. Each plate sector having includes a flat tab that is pressed against a face of the corresponding support sector and is fixed to this face by brazing. The tab is inserted into a through-slot in the support sector in order to improve its integrity in operation.
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
40.
METHOD FOR CHEMICALLY PICKLING A CAST METAL PART WITH POROUS CERAMIC CORE(S)
A method for chemically pickling a cast metal part, including a metal envelope which delimits an inner space in which at least one porous ceramic core is housed, and an outer space, the ceramic core being in fluid communication with the outer space, which method including: filling the pores of the ceramic core with a liquid; and then chemically pickling the cast metal part. This chemical pickling method may be implemented in a method for manufacturing a metal part by investment casting. This method is applicable at least to manufacture of turbine blades for turbomachines and, especially, for aircraft turbojet engines.
The invention relates to a method for verifying the positioning of a fibrous preform in a blade, the blade having been obtained by injecting a resin into a mould having the shape of a blade and in which a preform has been placed, the blade extending in an orthonormal blade frame of reference X, Y, Z, the blade comprising a blade root extending longitudinally along an axis X, a vane extending from the blade root along an axis Z, the blade having a thickness defined along an axis Y, the preform comprising glass tracers positioned at the surface of the preform, the centre of the tracers defining a neutral axis located at a height along the axis Z in the direction defined by the axis X, the method comprising the following steps: the acquisition (E31) of tomographic 2D projections of the blade using an imaging system comprising an X-ray source, each projection being acquired at a given orientation of the X-ray source with respect to the blade; the combining (E32, E32a, E32b) of the 2D projections in the direction of the axis Y so as to obtain a cumulative 2D image in the directions X and Z; the determining (E33), for each pixel column defined in the direction of the axis Z, of a greyscale profile; the processing (E34) of each of the profiles obtained so as to locate the position, in Z, of the neutral axis in the direction of the axis X.
G01N 23/044 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material using laminography or tomosynthesis
G01N 23/083 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and measuring the absorption the radiation being X-rays
B29C 70/54 - Component parts, details or accessories; Auxiliary operations
42.
ASSEMBLY FOR AIRCRAFT TURBINE ENGINE COMPRISING AN IMPROVED SYSTEM FOR LUBRICATING A FAN DRIVE REDUCTION GEAR
An assembly for an aircraft turbine engine includes a fan drive reduction gear and a lubrication system including: a reduction gear housing; a lubricant tank; a lubricant supply circuit including a feed pump; and a lubricant recovery circuit including a pump for recovering lubricant from the reduction gear housing. The recovery circuit includes a lubricant distributor, including: a lubricant inlet communicating with a lubricant outlet of the housing; an air inlet; and a distributor outlet, the distributor being able to adopt a lubricant recovery configuration and a configuration for retaining the lubricant in the housing.
A manufacturing method for a component made from composite material, in particular of a turbomachine, includes the steps of producing a preform with a fibrous reinforcement comprising a first fibrous portion and a second fibrous portion, and injecting a pressurized matrix into an injection chamber of an injection mold, in which the preform is arranged. The method further includes the steps of polymerizing the preform and positioning a flexible pocket that encloses a fluid and that is arranged between the first fibrous portion and the second fibrous portion before the injection step. The fluid is configured to apply an additional pressure to the preform of the fluid during the polymerization step.
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
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
F01D 5/18 - Hollow blades; Heating, heat-insulating, or cooling means on blades
44.
METHOD AND SYSTEM FOR CONTROLLING A THRUST REVERSER OF AN AIRCRAFT
A method for controlling a turboj et engine thrust reverser during an aborted aircraft takeoff, the thrust reverser including doors movable between a stowed position, an overstowed position and a deployed position; door actuators to move the doors between the stowed, overstowed and deployed positions; a device for locking the doors in the stowed position, moveable between a locking position and an unlocking position; and a lock actuator to move the locking device between the locking and unlocking positions. The method includes decreasing the engine speed of the turbojet engine by following a setpoint value below a first engine speed threshold value at which the aerodynamic forces being exerted on the doors are equal to the forces developed by the door actuators; controlling the door actuators to bring the doors into the overstowed position; controlling the lock actuator to bring the locking device into the unlocking position.
A control valve composed of sectors of stationary blades, which are externally connected to a stator, includes an inner platform divided into an outer portion formed of sectors and associated, by radial slide connections, with an inner portion in the form of a ring which is continuous over a circumference, which carries an abradable seal. Springs ensure the radial position of the crown, and the concentricity thereof to the rest of the machine. The diameter of the crown is essentially determined by the temperature of the gases which pass in front of it. The independence of the radial thermal deformations is ensured by slides composed of sliding sleeves.
A turbine rotor includes a disc with cavities, a plurality of blades, each with a root received in one of the cavities, and an axial retention system including a first series and a second series of strips circumferentially distributed around an axis of the rotor, the first and the second series being axially superimposed and arranged such that at least two strips circumferentially adjacent to the first series are axially superimposed on a strip of the second series and each strip of the first and second series is arranged opposite a cavity of the disc so as to axially block the root of a blade.
A turbomachine turbine casing that extend around an axis and includes an annular wall and a cooling device. The annular wall is provided with a casing hook which extends in radial protrusion from an inside of the annular wall. The casing hook allows a mounting, on the turbomachine turbine casing, of ring segments disposed circumferentially end to end around the axis. The cooling device includes a collector duct intended to convey cooling air, the collector duct extending circumferentially around the annular wall. The collector duct has a cooling air inlet and a cooling air outlet. The collector duct and the annular wall have a common portion, which delimits the collector duct and from which the corresponding casing hook extends.
ARRANGEMENT FOR AN AIRCRAFT TURBINE ENGINE HAVING IMPROVED LUBRICATION, THE ARRANGEMENT COMPRISING A SHAFT ROTATABLY COUPLED TO A FOLLOWING MEMBER BY MEANS OF SPLINES
An arrangement for an aircraft turbine engine, including a shaft and a follower element rotatably coupled to the shaft by a spline connection, the arrangement including upstream and downstream connections for radially centring the follower element relative to the shaft; means for spraying a lubricant into a collection cavity; a passage for receiving lubricant, which passage opens into the collection cavity and into a cavity for lubricating the splines which is partially defined by the upstream and downstream radial centring connections; and a passage for discharging lubricant, which passage opens into the cavity for lubricating the splines and outside the arrangement.
A turbofan aircraft propulsion assembly includes a retractable heat exchanger. The heat exchanger can be deployed in a secondary flow path of the propulsion assembly so as to cool a fluid circulating in a circuit of a transfer module of the heat exchanger using a secondary flow circulating in the secondary flow path.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE CLAUDE BERNARD LYON I (France)
Inventor
Laffay, Paul Clément Guillaume
Aguilera, Fernando Gea
Regnard, Josselin David Florian
Jacob, Marc Cornelius
Moreau, Stéphane
Abstract
An acoustic treatment grid intended to be mounted inside or at the outlet of a duct of a bleed valve of a turbomachine of an aircraft intended to convey a gas flow, the grid comprising a perforated plate and circular orifices traversing the perforated plate along a first direction, the orifices having a diameter and a geometrical center.
An acoustic treatment grid intended to be mounted inside or at the outlet of a duct of a bleed valve of a turbomachine of an aircraft intended to convey a gas flow, the grid comprising a perforated plate and circular orifices traversing the perforated plate along a first direction, the orifices having a diameter and a geometrical center.
Each orifice is separated from an adjacent orifice by a space, the length of which is equal to the product of the diameter of said orifice and a spacing coefficient of a value between 1.1 and 6.
F16K 47/08 - Means in valves for absorbing fluid energy for decreasing pressure and having a throttling member separate from the closure member
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
51.
ANALYSIS DEVICE FOR DETECTING SOLID PARTICLES IN A LUBRICANT
Analysis device for detecting solid particles in suspension in a lubricant, the analysis device comprising one or more ferromagnetic solid particle sensors, one or more other sensors able to detect non-ferromagnetic solid particles, and one or more magnets. The ferromagnetic solid particle sensors are offset in a direction perpendicular to a main direction of flow of the lubricant in relation to the other sensors, and the magnets are arranged so as to attract ferromagnetic solid particles towards the sensors of ferromagnetic solid particles by drawing them away from the other sensors.
The invention relates to a combustion chamber for a gas turbine engine, in which two openings extend through a metal wall (58) and a wall made of a refractory material (16). A device (76) guides a spark plug (48) in said two orifices, the guiding device comprising a metal flange (80) and a floating ring (82), the flange being shrunk onto or welded to the metal wall (58).
The present invention relates to a fan rotor with variable pitch blades, comprising a rotor disc, equipped at its periphery with a plurality of rotary fasteners (16), each fastener (16) comprising a cell (17) for receiving the root (150) of a blade (15). This rotor is characterised in that an elongated wedge (2) and a prestressing rod (3) with at least one cam (33) are also arranged in each cell (17), the bottom of the cell (17) comprising as many retraction cavities (173) as the rod (3) comprises cams (33), in that the prestressing rod (3) is interposed between the wedge (2) and the bottom (171) of the cell (17), so that said cam (33) is facing a corresponding retraction cavity (173) and in that the prestressing rod (3) can rotate about its longitudinal axis (X2-X′2), between a rest position, in which the cam (33) is housed in the retraction cavity (173), and an armed position, in which the cam (33) exerts a radial pressure on the central region (22) of the wedge (2) so as to move the wedge (2) towards the blade root (150).
A turbomachine blade is made of composite material formed of woven fibers and embedded in a polymerized resin. The blade includes a root connected by a stilt to a vane which comprises a pressure side and a suction side. At least one anti-wear strip of fabric is located on a surface of the root and/or the stilt. At least one anti-wear strip is secured to the root surface in a unitary manner by the resin.
B29C 70/68 - Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts by incorporating or moulding on preformed parts, e.g. inserts or layers
55.
TURBINE ASSEMBLY, AND GAS TURBINE ENGINE PROVIDED WITH SUCH AN ASSEMBLY
A turbine assembly (1) comprising: —a plurality of turbine ring sectors (20) made of ceramic-matrix composite material, —a ring support structure (3), comprising an annular shroud (6), and in addition −a plurality of angular spacer sectors (70) together forming an annular spacer (7), said annular spacer (7) being, on the one hand, fixed to the turbine ring (2) and, on the other hand, fixed to said annular shroud (6), characterized in that said turbine assembly (1) comprises at least one air diffuser (8), which is configured to diffuse cooling air onto the radially outer face (212) of at least one of said turbine ring sectors (20), and in that said at least one air diffuser (8) is mounted by being nested on one of said angular spacer sectors (70), in a nested position.
An actuation system includes: a generating rotating electrical machine including a first stator including at least one output stator winding; a driving rotating electrical machine including a second rotor, intended to actuate a lubrication pump, and a second stator including at least one input stator winding; an array of switches electrically connected to the at least one output stator winding and to the at least one input stator winding, for electrically connecting them together or disconnecting them from each other depending on the state of the array of switches; and a control device configured to control the array of switches, depending on the state of a control signal, in such a way as to electrically connect or not connect the at least one output stator winding and the at least one input stator winding together.
A bypass turbine engine includes a fixed casing, a first shaft (low-pressure shaft), a second shaft (high-pressure shaft), at least one accessory to be driven by a motor powered with electrical energy, a first intermediate shaft tapping mechanical power off the low-pressure shaft, a second intermediate shaft tapping mechanical power off the high-pressure shaft, and an electrical energy generator assembly coupled to the first and second intermediate shafts so as to receive mechanical power from the first and second intermediate shafts. The generator assembly converts the mechanical power received from the first and second intermediate shafts into electrical energy to power the motor or motors, which comes simultaneously from the mechanical power tapped off the low-pressure shaft and the mechanical power tapped off the high-pressure shaft. The generator assembly is housed in an arm in the lower part of the turbine engine and extending vertically into a bypass flow duct.
Aircraft turbomachine comprising a casing, a fan, a compressor and a turbine and an epicyclic gear train comprising an input driven in rotation by the turbine, a first output stage configured to drive in rotation the compressor and a second output stage coupled to the first output stage and configured to drive in rotation the fan, the compressor being driven in rotation by the ring gear of the first output stage.
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
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
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
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
Device for distributing oil from a rolling bearing (8) for an aircraft turbine engine, comprising: - a rolling bearing (8), an oil distributor ring (5) and - an annular track (26) of a dynamic seal (22), characterised in that it further comprises a nut (16) screwed on to a thread (5d) of the distributor ring and bearing axially against an axial end of the inner ring so as to clamp it axially, and in that the annular track is configured to bear axially against the distributor ring and comprises rotating locking elements (27) engaging with the additional elements (28)of the nut.
A system for managing the deceleration of an aircraft enabling the control in real time of the position of the aircraft on a braking axis, includes a braking system; a calculator configured to: calculate, from aircraft data and from external data, a sequence of use of the braking system intended to brake the aircraft over a predetermined braking distance which associates a predetermined position on the braking axis with each braking instant; update in real time the sequence of use as a function of the difference between the position of the aircraft and the predetermined position; and a controller configured to control the braking system as a function of the sequence of use.
A mold for manufacturing a turbomachine fan casing of composite material, includes a mandrel around which a fibrous preform of the fan casing is to be wound; counter-mold angular sectors assembled on the external contour of the mandrel which are intended to close the mold and to compact the fibrous preform wound on the mandrel; wherein each angular sector includes a first lateral flange positioned at the first end of the angular sectors and a second lateral flange positioned at a second end of the angular sectors, the first and second lateral flanges cooperating respectively with the second and first lateral flanges of the adjacent angular sectors. An angular sector includes a first groove formed in the first of second lateral flange, a first seal being positioned in the first groove, the first seal to be compressed between the first and second lateral flanges of two adjacent angular sectors.
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
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 70/32 - 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
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
62.
AIRCRAFT TURBINE ENGINE WITH AN OFF-AXIS PROPELLER
An aircraft turbine engine includes a turbine shaft having a first axis of rotation, a propulsion propeller having a second axis of rotation parallel to and spaced from the first axis, and a mechanical reduction gear coupled to the turbine shaft and rotating the propeller. The reduction gear has a sun gear connected to the turbine shaft, a ring gear, and at least two planet gears, each including a first external toothing that is meshed with an external toothing of the sun gear. A second external toothing is located within the ring gear and is meshed with an internal toothing of the ring gear.
Disclosed is an aircraft turbine engine (10), comprising a gas generator (12) and a fan (14) arranged upstream from the gas generator (12) and configured to generate a gas inlet stream (F), part of which flows into a duct of the gas generator to form a primary stream (36), the turbine engine (10) comprising an electrical machine that is mounted coaxially downstream from the fan (14) and that comprises a rotor (62a) surrounded by a stator (62b) carried by an annular shroud (64), this shroud (64) being surrounded by a casing (40) of the gas generator that defines, with this shroud (64), a section of the flow duct for the primary stream (36), stationary vanes (42, 68) for straightening this primary stream (36) extending into this path.
The present invention relates to a turbomachine rotary-fan blade (2), comprising a body (20) made of a composite material, a metal reinforcement part (3) comprising a metal upstream nose (31), characterised in that the metal upstream nose (31) comprises, at least on the metal part (27b) of the blade tip, a recess (4) of longitudinally tapering thickness (AX), delimiting on the metal part (27b) over a height (H) at least one metal projection (5) with prescribed wear, which has a longitudinally tapering thickness and which is configured to detach at least partially in the presence of tangential friction in the second thickness direction (EP) against the metal part (27b), the recess (4) and the metal projection (5) with prescribed wear extending the first metal fin (32) and/or the second metal fin (33) and/or the upstream edge (22) of the body (20) made of composite material.
The invention relates to a planetary gear reducer (21) for an aircraft turbine engine (1), the reducer (21) comprising a sun gear (30) and a ring gear (33) which are centred on a longitudinal axis X and a planet carrier (32) which carries at least one planet gear (31) rotatably mounted about a planet axis A parallel to the longitudinal axis X, the sun gear (30) being rotatable about the longitudinal axis, the planet gear (31) meshing both with the sun gear (30) and the ring gear (33), the planet carrier (32) being movable about the longitudinal axis and the ring gear (33) being rotationally fixed. According to the invention, the reducer (21) comprises an electric machine (50) integrated therewith and which comprises a rotor (51) mounted on the planet carrier (32) in such a way as to be rotated about the longitudinal axis X and a stator (52) mounted on the ring gear (33).
A rotary assembly for a turbomachine including a rotor including at least two consecutive rotor stages provided with a plurality of blades and an annular rotor shroud connecting the two consecutive rotor stages, a stator including: at least one stator stage provided between the two consecutive rotor stages including a plurality of vanes, a turbomachine stator vane root, an annular clamping part and an annular support of abradable material, the root extending radially and being clamped axially between the annular support of abradable material and the annular clamping part, A space separates a radially internal end of the root and the annular support of abradable material. A turbojet engine including a rotary assembly as previously.
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 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
F01D 11/02 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
67.
Blade pivot with adjustable orientation and protected integrity for a turbomachine fan hub
The invention relates to a blade pivot with adjustable orientation for a turbomachine fan hub, comprising: a stud having a fastener configured to retain a blade root and coupling means for the transmission of a twist torque; a ball bearing for taking centrifugal forces having an inner ring and an outer ring; a first clamping nut intended to be screwed on an inner thread of the hub to ensure clamping of the outer ring of the ball bearing; a rolling-element bearing for taking transverse forces having an inner ring and a smooth outer ring; a locking ring mounted between these two inner rings to provide for them a respective transverse support; an anti-rupture sleeve carrying the inner ring of the rolling-element bearing for taking transverse forces and the terminal end of which is extended transversely beyond the outer ring of this bearing.
An enclosure for a turbomachine includes a turbomachine drive shaft rotating about a longitudinal axis (X) by means of two roller bearings, an upstream bearing and a downstream bearing, each having an inner ring carried by the drive shaft. The two bearings share a single integral outer ring that has an upstream end and a downstream end connected to one another by a section of studs. The single outer ring is carried by an upstream base plate and a downstream base plate of a bearing support configured to be attached to a stationary structure of the turbomachine.
F16C 19/06 - Bearings with rolling contact, for exclusively rotary movement with bearing balls essentially of the same size in one or more circular rows for radial load mainly with a single row of balls
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
69.
Assembly of a turbomachine turbine ring and stator
The invention relates to a turbine assembly comprising: a plurality of ring sectors forming a ring; a support structure comprising a shroud from which there project radial flanges which serve to hold latching tabs of each ring sector; a stator positioned downstream of said ring and comprising a blade provided with a radially outer platform and extending axially opposite said ring; an annular space being defined between the ring, the support structure and the platform, this annular space having passing through it a leakage air current, the assembly comprising a hollow ring which occupies said annular space and which is shaped so as to collect, channel and expel said leakage air current, more specifically to channel and expel it into the radially inner region of said annular space.
A method for manufacturing a metal component includes the following steps: a shell made of a titanium-based material is provided, the shell having a first surface and a second surface remote from the first surface; a covering layer made of a titanium fire-resistant material is produced by additive manufacturing on the shell such that the covering layer at least partially covers the first surface and/or the second surface; and, after the additive manufacturing step, the metal component is heat treated at a temperature of between 200° C. and 1000° C.
C22F 1/18 - High-melting or refractory metals or alloys based thereon
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
B22F 10/25 - Direct deposition of metal particles, e.g. direct metal deposition [DMD] or laser engineered net shaping [LENS]
C23C 24/10 - Coating starting from inorganic powder by application of heat or pressure and heat with intermediate formation of a liquid phase in the layer
71.
Profiled structure for an aircraft or turbomachine for an aircraft
OFFICE NATIONAL D'ETUDES ET DE RECHERCHES AEROSPATIALES (France)
Inventor
Gea Aguilera, Fernando
Barrier, Raphaël
Gruber, Mathieu Simon Paul
Polacsek, Cyril
Posson, Hélène Dominique Jeanne
Abstract
A turbomachine includes a rotor and a stator, the stator having a plurality of profiled structures, each profiled structure being elongated in a direction of elongation in which the profiled structure has a length exposed to an airflow, and having a leading edge and/or a trailing edge, at least one of which is profiled and has, in said direction of elongation, serrations defined by a succession of peaks and troughs and having a geometric pattern transformed, over at least a part of said length exposed to the airflow, by successive scaling, via multiplicative factors, in the direction of elongation and/or transverse to the direction of elongation. The geometric pattern, as defined with reference to a radial distribution of the integral scale of the turbulence, evolves in a non-periodic manner.
A turbine includes a rotor surrounded by a stator, a first movable stage including a series of rotating vanes, a second movable stage including a second series of rotating vanes, a distributor including a series of stationary vanes, the vanes including platforms jointly delimiting a separation between a main space in which a hot flow circulates through the vanes, and a secondary space surrounding a hub of the rotor. The rotor includes an aperture through which air is blown towards the secondary space, this air being discharged towards the main space. Straightening fins are carried by the distributor in the secondary space for straightening the air flow in order to change its gyration.
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 11/24 - Actively adjusting tip-clearance by selectively cooling or heating stator or rotor components
A blade of a turbomachine turbine, including a platform and a vane, including a cooling area extending along the trailing edge, with three distinct types of orifices, wherein the blade includes, at the root of the blade, a first type of orifices formed in the thickness of the trailing edge fillet, at the top of the blade, a third type of orifice formed in the thickness of at least one of the lower surface and upper surface walls, and in the middle of the blade, a second type of orifice formed in the thickness of at least one of the lower surface and upper surface walls. The lower surface wall, the upper surface wall and the trailing edge fillet define a dihedral with a radially changing radius and spacing.
The invention relates to a field coil (18, 20), in particular for a satellite hall-effect plasma thruster, said field coil (18, 20) comprising a core (22) on which a conductor (24) is wound, characterized in that the conductor comprises an inorganic insulation cable (26) impregnated with a high-temperature-resistant silicone coating (32).
A fibrous texture forming the fibrous reinforcement of a turbomachine blade made of composite material which has a three-dimensional weaving between warp yarns or strands made of first fibers and weft yarns or strands made of first fibers, the texture including a blade root portion, a blade airfoil portion and a free end, the texture extending along the transverse direction between a first edge corresponding to a leading edge of the blade and a second edge corresponding to a trailing edge of the blade. The blade airfoil portion has a reinforced area extending along the longitudinal direction from the free end over a first length, and along the transverse direction from the second edge over a second length less, the reinforced area including weft yarns or strands made of second fibers different from the first fibers, the second fibers having an elongation at break greater than the first fibers.
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
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
76.
MOLD FOR MANUFACTURING A TURBINE ENGINE FAN CASING FROM A COMPOSITE MATERIAL
A mold for manufacturing a turbomachine fan casing made of composite material, includes a main axis mandrel around which a fibrous preform of a fan casing is intended to be wound; a plurality of counter-mold angular sectors assembled on the outer contour of the mandrel which are intended to close the mold and to compact the fibrous preform wound on the mandrel; wherein a flat seal with a main elongation axis directed along the main axis is arranged between each angular sector, the flat seal being compressed between two adjacent angular sectors, a first angular sector including a sealing portion passing below a lower face of the flat seal while a second angular sector including a sealing portion passing above an upper face of the flat seal.
B29C 70/34 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
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
77.
Acoustically optimized discharge line grid with channels
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
Regnard, Josselin David Florian
Laffay, Paul Clément Guillaume
Jacob, Marc Cornelius
Moreau, Stéphane
Abstract
Discharge grate intended to be mounted inside or at the outlet of a conduit of a discharge valve of a turbine engine of an aircraft, the discharge grate comprising an upstream face intended to receive a gas flow, a downstream face parallel to the upstream face and intended to deliver the gas flow received on the upstream face, and orifices passing through the perforated plate from the upstream face to the downstream face and intended to convey the gas flow through the perforated plate.
The discharge grate comprises for each orifice of the perforated plate a tubular channel, coaxial with the orifice with which it is associated, and projecting from the downstream face of the perforated plate.
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F01D 17/12 - Final actuators arranged in stator parts
F02K 3/075 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type controlling flow ratio between flows
A sealing assembly (1) for a turbine engine comprising a first element (2) and a second element (3), the first and second elements (2, 3) being concentric and in relative rotational movement with respect to each other about an axis of rotation (X), the sealing assembly (1) comprising at least one first wiper (4a) and an abradable member (5), the first wiper (4a) being annular in shape and carried by the first element (2), the first wiper (4a) extending radially towards the abradable member (5) and continuously around the axis of rotation (X), the abradable member (5) being annular in shape and carried by the second element (3), the abradable member (5) extending tangentially opposite the first wiper (4a), the first wiper (4a) comprising primary angular portions (11), each extending tangentially along a primary angular sector (11′), the primary angular portions (11) each having, in cross-section, a first constant profile, characterised in that the first wiper (4a) comprises secondary angular portions (13) each extending tangentially along a secondary angular sector (13′), the secondary angular portions (13) each having, in cross-section, a second profile different from the first profile, the number of secondary angular portions (13) being equal to the number of primary angular portions (11), the secondary angular portions (13) being interposed between the primary angular portions (11).
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
79.
Fibrous texture for turbine engine blade made of composite material
The invention relates to a fibrous texture intended to form the fibrous reinforcement of a turbine engine blade made of composite material, the texture being in a single piece and having a three-dimensional weave between a plurality of first fiber warp yarns or strands extending in a radial direction and a plurality of first fiber weft yarns or strands extending in a chord direction, the texture comprising a blade root portion and a blade airfoil portion extending between the blade root portion and a free end of the fibrous texture. The blade airfoil portion has a reinforced area in the vicinity of the free end of the texture comprising weft yarns or strands made of second fibers different from the first fibers.
Turbine shroud assembly comprising sections (10) made from CMC and forming a shroud (1) and a support structure (3), each section having a base (12) with a radially internal face (12a) and a radially external face (12b), from which there extend in a projecting manner an upstream attachment lug (14) and a downstream attachment lug (16), the support structure comprising a collar (31), from which there radially extend in a projecting manner towards the shroud an upstream radial flange (32) and a downstream radial flange (36), by which the lugs of each section of the shroud are retained, the shroud (1) being retained by axial pins (119, 120) which cooperate, on the one hand, with the upstream radial flange, via first and second annular end plates (33, 34), and directly with the downstream radial flange and, on the other hand, with the upstream and downstream attachment lugs, respectively.
An assembly for a fluid circuit of a turbine engineincludes a main branch, a fluid control valve, and a bypass branch which is arranged in parallel to the main branch. The fluid control valve includes a main outlet which is fluidly connected to the main branch and a bypass outlet which is fluidly connected to the bypass branch. The fluid control valve including a shutter, a spring for biasing the shutter, and control members for electrically controlling the shutter to a main open position or a bypass position. The spring for biasing the shutter is configured to bias the displacement of the shutter to the bypass position when a fluid pressure value is lower than a first threshold value.
A turbofan engine has a primary duct , in which there flows a primary flow at a primary pressure and a secondary duct, which radially surrounds the primary duct and in which there flows a secondary flow at a secondary pressure. The primary duct includes at least one compressor configured to compress the primary flow , a turbine driving the compressor in rotation and a combustion chamber designed to receive, at an inlet, the primary air flow compressed by the compressor The turbomachine further includes a cooling circuit extending between the compressor and the turbine.The cooling circuit has an air flow rate regulating device arranged upstream of the turbine and having at least one valve that is configured to move between an open position and a closed position, located between the compressor and the combustion chamber in the primary duct and the pressure (PS) in the secondary duct.
A composite platform for a fan of an aircraft turbine engine. The platform includes an elongate wall and is configured to extend between two fan blades. The wall includes an aerodynamic outer surface and an inner surface, on which a fastening tab is located, wherein the fastening tab is configured to be attached to a fan disc. The fastening tab is integrally formed with a metal reinforcement which has a plate having an elongate shape and which extends over more than 50% of the longitudinal extent of the wall, the wall being produced by overmolding a resin on the plate so as to be integrated into the wall.
A fan module for an aircraft turbomachine includes a fan with a disc carrying fan blades and a rotor of an electric machine. The rotor has an annular shape and is mounted coaxially downstream of the fan. The module further includes an annular support for the rotor with a downstream end fixed to the rotor and with an upstream end fixed to the fan disc. The support is fixed to the disc by shear bolts configured to break when a torque transmitted by the disc to the support exceeds a predetermined threshold.
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F01D 5/06 - Blade-carrying members, e.g. rotors - Details thereof, e.g. shafts, shaft connections
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
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
85.
METHOD FOR MANUFACTURING A COMPOSITE PLATFORM FOR AN AIRCRAFT TURBINE ENGINE FAN
A composite platform for an aircraft turbine engine fan includes a wall of elongate shape that is configured to extend between two fan blades. The wall has an aerodynamic external face and an internal face on which is disposed a fixing tab configured to be fixed to a fan disc. A method for manufacturing the composite platform includes the steps of: a) producing a preform by three-dimensionally weaving of fibers, b) unbinding some of the fibers of the preform to detach at least one longitudinal layer of fibers from the rest of the preform, c) inserting a metal reinforcement between this layer and the rest of the preform, and d) injecting a resin into the preform so as to form said wall and secure the reinforcement to this wall.
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
B29C 70/72 - Encapsulating inserts having non-encapsulated projections, e.g. extremities or terminal portions of electrical components
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
The invention relates to a method for controlling a pitch angle of the vanes or blades of a propellant body of a turbine engine, comprising generating a pitch command (ifinal) according to a rotational speed of the propeller (XNmes) and a speed setpoint (XNcons), the method comprises a nominal regulating chain (13), wherein the pitch command is further generated according to a value of a pitch angle (βmes) of the vanes or blades of the propellant body, and an off-nominal regulating chain (16), wherein the pitch command is generated independently of a value of a pitch angle of the vanes or blades of the propellant body.
A turbomachine turbine blade, includes a platform, a vane, a cooling cavity supplying a plurality of cooling outlets provided along the trailing edge, two radially adjacent cooling outlets defining therebetween a rib. At least one cooling hole is formed in the thickness of at least one rib and/or in the thickness of a portion of the trailing edge fillet located in the axial extension of at least one rib, so as to ensure fluid communication for a cooling flow between the inside and the outside of the blade for cooling the at least one rib.
A vane for a turbine engine includes a body in the form of an aerodynamic profile formed by a shell produced from a composite material formed from a three-dimensional textile of reinforcement fibres consolidated by a hardened resin. The shell forms a pressure surface and a suction surface of the vane connected to each other while forming on one side a leading edge and on the opposite side a trailing edge of the vane, and a core including a core body surrounded by the shell. The core is produced from non-porous plastics material, and the core body is in the form of an aerodynamic profile and delimits at least one closed cavity of the core.
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 70/24 - 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
The present invention relates to a fan blade (3) having a structure made from a composite material, comprising a fibrous reinforcement (5), which is obtained by three-dimensional weaving of warp strands and weft strands, and a matrix in which the fibrous reinforcement (5) is embedded, wherein—the fibrous reinforcement (5) comprises a first portion (14) forming the trailing edge (9) of the structure made from a composite material and a second portion (15) forming its leading edge (8), and wherein—the warp strands of the fibrous reinforcement (5) comprise first strands (12) having a predetermined stiffness and second strands (13) having a greater stiffness than that of the first strands (12), the first portion (14) comprising all or part of the first strands (12) and being devoid of second strands (13) while the second portion (15) comprises all or part of the second strands (13).
A method for refitting blade shrouds of a rotor wheel in an aircraft turbomachine is described. The rotor wheel has a disc bearing blades that each have an airfoil extending between a root and a shroud, the shroud of each blade having lateral edges comprising including shapes complementary to the lateral edges of the shrouds of the adjacent blades. The lateral edges of the shrouds are interlocked in engagement with one another such that anti-Wear coatings of these edges are in contact With one another in a desired interlocking engagement position, and at least one of the lateral edges of at least one of the shrouds being able to be disengaged from the lateral edge of an adjacent shroud in an undesired disengagement position. The method includes, when an undesired disengagement position is detected, a step of inserting a re-engagement device into the turbomachine.
A nozzle blade for a turbine engine includes a blade body in which at least one through-cavity is provided extending between an inner end of the blade and an outer end of the blade and housing, respectively, a first and a second liner, the first liner extending along a trailing edge of the blade, the second liner extending along a leading edge of the blade, wherein each of the first and second liners includes an intertwining system suitable for linking the first liner to the second liner, independently of the blade body.
The invention relates to an abradable coating for a turbomachine, comprising a mineral compound having a Mohs hardness of 6 or less and a melting temperature of more than 450° C. or even 800° C., and a polymeric compound with a content of between 40% and 70% by volume.
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 4/04 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the coating material
93.
AIR OUTLET FOR A NACELLE FOR AN AIRCRAFT BYPASS TURBOJET ENGINE COMPRISING A GUIDING DEVICE TO FAVOR A REVERSE THRUST PHASE
An air outlet for a nacelle for an aircraft turbofan having an inner wall and an outer wall connected to each other by a trailing edge, at least one radial through aperture extending over an angular portion of the air outlet and having inner and outer open faces and, for each aperture, a guiding device having inner and outer movable members movably mounted between a closed position wherein the inner and outer movable members respectively close off the inner and outer open faces, the aperture defining a closed cavity and an open position wherein the inner and outer movable members are configured to allow an external air flow to circulate in the aperture to support a reverse thrust phase.
F02K 1/66 - Reversing fan flow using reversing fan blades
F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
F02K 3/04 - 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 invention relates to a composite blade (5) for a turbine engine rotor, for example, an unducted propeller, comprising a skin (6) made of woven fibres forming the outer profile of the blade and an attachment (11) with cylindrical geometry in the direction of the span emerging from the blade root (9), which attachment is intended to retain the blade (5) on a hub of the rotor, characterised in that it further comprises a spar (16) having a hollow tube structure made of braided carbon fibres, fixed to the attachment (11) and extending inside the skin (6) over at least part of the span of the blade (5). The invention also relates to a propeller comprising said blade and to a method for manufacturing said blade.
A device for cleaning a vent of a first forging die suitable for receiving, between the first die and a second die, a metal blank to be shaped in order to produce a metal part, the device for cleaning including a pump for injecting an aqueous solution, suitable for injecting the aqueous solution into the vent, an ultrasound generator, suitable for generating ultrasound waves within the vent, an air intake suitable for ensuring a venting of the vent, and a valve system for alternately controlling at least the injection of aqueous solution and ultrasound and the venting of the vent.
B08B 3/12 - Cleaning involving contact with liquid with additional treatment of the liquid or of the object being cleaned, e.g. by heat, by electricity or by vibration by sonic or ultrasonic vibrations
B21C 35/06 - Cleaning dies, ducts, containers or mandrels
B30B 15/00 - PRESSES IN GENERAL; PRESSES NOT OTHERWISE PROVIDED FOR - Details of, or accessories for, presses; Auxiliary measures in connection with pressing
96.
WAVY TILTING OF PLATFORMS AT THE ROTOR-STATOR GAPS IN A TURBINE ENGINE COMPRESSOR
A set of compression stage(s) of a turbomachine, forming an annular fluid passage and comprising at least one annular stator platform and/or at least one annular rotor platform having an outer longitudinal profile inclined (ISi/IRi) inwards and upstream with respect to a nominal profile of the fluid stream, where the inclination (ISi/IRi) of the outer longitudinal profile of the or each of the annular platforms, relative to the nominal profile of the fluid stream, oscillates along the circumference of the annular platform or platforms, between a maximum value in front of the blades of the annular platform and a minimum value between each pair of adjacent blades of the annular platform.
A turbomachine module with a longitudinal axis comprising an unducted propeller rotated about the longitudinal axis and at least one straightener. The module includes a plurality of unducted variable-pitch stator blades extending along a radial axis, perpendicular to the longitudinal axis, from a fixed casing. The module includes a first stator blade pitch-changing system. The pitch-changing system includes at least one first control that includes a first fixed body connected to the fixed casing and a first body which is axially mobile in relation to the first fixed body and at least one first joining mechanism joining each stator blade to the first mobile body of the first control. The first joining mechanism includes: a joining ring centered on the longitudinal axis, joined to the feet of each stator blade and at least one lever joined, on one hand, to the joining ring and, on the other hand, to the first mobile body of the first control.
Device for distributing oil from a rolling bearing (8) for an aircraft turbine engine, comprising: —a rolling bearing (8), —a body (5) for distributing oil, which body is configured to be mounted on a turbine engine shaft (4), said body comprising: i) a first outer cylindrical surface (5a) for mounting the inner ring (12) of the bearing, ii) a ring-shaped scoop (11) for recovering oil, iii) a ring-shaped track (26) of a dynamic seal (22), iv) a circuit (7) for lubricating the bearing and cooling the track, said circuit being formed in the body, characterised in that the ring-shaped scoop is the first scoop (11a) which supplies a first portion (7x) of the circuit with a view to cooling the track, and in that the body comprises a second ring-shaped scoop (11b) for recovering oil, which scoop supplies a second portion (7y) of the circuit with a view to lubricating the bearing.
A casing of an aircraft turbomachine includes an annular shell extending around an axis A and made of a composite material having fibers that which are woven and embedded in a resin. An annular layer made of abradable material extends inside the shell, around axis A and is obtained by spreading and polymerizing a paste. Support panels extend around axis A and are interposed between the shell and the abradable layer.
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
B05C 11/02 - Apparatus for spreading or distributing liquids or other fluent materials already applied to a surface; Control of the thickness of a coating
B05C 17/10 - Hand tools for removing partially or for spreading or redistributing applied liquids or other fluent materials, e.g. colour touchers
B05C 11/10 - Storage, supply or control of liquid or other fluent material; Recovery of excess liquid or other fluent material
B05C 5/02 - Apparatus in which liquid or other fluent material is projected, poured or allowed to flow on to the surface of the work from an outlet device in contact, or almost in contact, with the work
B05C 9/06 - Apparatus or plant for applying liquid or other fluent material to surfaces by means not covered by groups , or in which the means of applying the liquid or other fluent material is not important for applying two different liquids or other fluent materials, or the same liquid or other fluent material twice, to the same side of the work
100.
METHOD FOR MONITORING ELEMENTS IN A TURBINE ENGINE COMBUSTION CHAMBER
Method for monitoring an ignition system of an aircraft turbine engine, the ignition system including a combustion chamber fitted with a spark plug and an ignition casing, the turbine engine including an accelerometer and an acoustic sensor. The method includes steps of: acquiring accelerometric and acoustic data representative of spark plug breakdown noises from signals generated by the accelerometer and the acoustic sensor, detecting breakdown peak times from the accelerometric data and the acoustic data, correlating the breakdown peak times detected from the accelerometric data and the breakdown peak times detected from the acoustic data, and establishing a diagnosis of the health of the ignition casing and a diagnosis for the spark plug on the basis of the correlation results.
patents
