The invention relates to a bearing for a rotating machine in an axial direction (Z-Z), comprising an inner ring (3) mounted around a shaft (1) and an outer ring (4) mounted in a housing (2), bearing elements (5) being arranged between said inner ring and said outer ring, characterised in that the outer ring (4) comprises a first half-ring and a second half-ring (4A, 4B), the first half-ring (4A) being locked against translation in the axial direction in relation to the housing (2), the second half-ring (4B) being mounted in a sliding manner in a receiving area between the first half-ring (4A) and an abutment (21) of the housing (2), the abutment (21) of the housing (2) comprising a preload member (7) exerting a pushing force on the second half-ring (4B) with the effect of positioning same such that it presses against the first half-ring (4A).
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
2.
METHOD FOR ISOLATING A COMBUSTIBLE FLUID TANK FROM A DOWNSTREAM PORTION OF A TURBOMACHINE SUPPLY SYSTEM IN CASE OF A FIRE, AND SUCH A SUPPLY SYSTEM
The invention concerns a supply system (200) for supplying fluid to a turbomachine (1) comprising a fluid tank (201), a downstream portion (210) located downstream from the tank (201), and a cutoff valve (230) located between the tank (201) and the downstream portion (210). The cutoff valve (230) is configured to be at least partially open when a rotational speed of a shaft of the turbomachine is higher than a first threshold and in the absence of a fire detected in the turbomachine (1). According to the invention, the cutoff valve (230) is closed when the two following conditions are satisfied: the rotational speed of the shaft is higher than the first threshold, and the rotational speed of the shaft was lower than the first threshold when a fire was detected in the turbomachine (1) during the flight.
The invention relates to a device for aiding remote-diagnosis in the course of an endoscopy of an aircraft engine, comprising at least one endoscope (2) used in situ by a first operator (20) to capture images of said engine and to perform measurements, in at least one remote terminal (10) used by at least one second operator (22) to analyze said video images and to transmit a diagnosis to the operator (20). The device according to the invention furthermore comprises a wireless communication interface allowing the operator (20) to exchange with the expert (22), in real time and interactively, the images captured and the results of the measurements performed.
B64F 5/00 - Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided forHandling, transporting, testing or inspecting aircraft components, not otherwise provided for
G02B 23/24 - Instruments for viewing the inside of hollow bodies, e.g. fibrescopes
4.
HEAT TREATMENT OF AN ALLOY BASED ON TITANIUM ALUMINIDE
The invention relates to a method for the treatment of an alloy based on titanium aluminide. The method comprises the following steps during which no hot isostatic pressing is carried out: a semi-finished product (7) produced by centrifugal casting is obtained; and said semi-finished product is then heat-treated in order to obtain a microstructure of the alloy comprising gamma grains and/or lamellar grains (alpha2/gamma).
The invention relates to a Hall-effect plasma thruster (10) comprising: - an annular discharge channel (121) around a principal axis (A) with an open downstream end, having an ionisation zone (128) between an internal wall (116) and an external wall (118), and furthermore comprising an anode (154) and a distributor (151), which are placed upstream of the ionisation zone (128). Characteristically, at least one side wall (152a,152b; 253d,253e) produced by additive manufacturing made of a material different to that of the anode (154, 254) partially separates the distributor (151, 251) from the ionisation zone (128,228).
The invention concerns a brush seal system, for application as an air/oil seal in a turbomachine shaft bearing, having an axis of rotation, the seal comprising a layer of carbon bristles (23), held between a first ring, referred to as the front ring (40), disposed upstream of a flow of air passing through the seal, and a second ring, referred to as the rear ring (60), disposed downstream of said flow of air, the surface of the layer of bristles, which is turned towards the front ring (40), being provided with a non-metal flexible element (70) that stops a portion of said flow of air (100) that flows through the seal in a direction substantially parallel to said axis of rotation.
The invention relates to a device (12) for a non-streamlined propeller (6, 7) with blades (10) having variable pitch of a turbomachine (1), the device comprising a support (15) intended to support a blade (10) and including a cylindrical foot (17), a bearing comprising a radially internal annular section (20) and a radially external annular section (21) capable of pivoting with respect to each other, the foot (17) being mounted in said internal section (20) of the bearing. The device comprises a hub (27) mounted in the foot (17), said hub (27) including a rim (29) comprising a face (30) coming to bear on the radially internal annular section (20) of the bearing, the hub (27) and the foot (17) of the support (15) comprising complementary engagement elements (18, 28) interacting with a locking component (23).
The invention relates to a blade (11A) for a turbine engine propeller, in particular a propfan engine, comprising a protruding part (16) on the leading edge (17) thereof, characterised in that said blade comprises means for controlling the position of the protruding part along the leading edge thereof.
Mandrels (1, 2) for the rolling of a ring (7) comprise impressions (5, 6) to accommodate the ring, which impressions have fillet radii where the various faces meet, so as to avoid the creation of sharp corners on the rolled ring (7) and the accidental formation of cracks.
The invention relates to a nacelle (1) for an aeroplane turbojet, comprising a frustoconical radially outer annular wall (13) from which a pylon (9) extends radially outwards, the pylon (9) being used to attach the nacelle (1) to a stationary part of the aeroplane. The nacelle is characterised in that it comprises at least one projecting area (22) extending radially outwards from the frustoconical outer wall (13), said projecting area (22) being arranged close to a downstream annular edge (14) of said outer wall (13) and close to the pylon (9).
B64D 29/06 - Attaching of nacelles, fairings, or cowlings
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
11.
TURBOMACHINE CASING COMPRISING A CAVITY-FREE SHROUD AND YOKES REINFORCED BY STIFFENERS
The invention relates to a turbomachine casing (1) comprising: a hub (2), an outer shroud (3), and yokes (11) projecting from the outer shroud (3), for attaching the housing (1), characterised in that it comprises at least one stiffener (10) extending between pairs of yokes (11) facing each other, and comprising a central part (18) and side arms (19) projecting from the central part (18).
The invention concerns a turbomachine (1) characterised in that it comprises: - an exhaust housing (7), comprising a plurality of arms (10), the space separating the arms defining openings (13) in which there circulates a primary air flow (29) of the turbomachine (1), - at least one conduit (2), a) configured to collect a compressed air flow at one of the ends (3) of same, b) the other end of the conduit (2) being connected to at least one opening (13) of the exhaust housing (7), so as to insert the collected air flow into said primary air flow (29), said collected air flow having, when inserted into the opening (13), a Mach number less than or equal to 0.5.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
13.
ROTOR DISK HAVING A CENTRIPETAL AIR COLLECTION DEVICE, COMPRESSOR COMPRISING SAID DISC AND TURBOMACHINE WITH SUCH A COMPRESSOR
The rotor disk (3B) for a compressor comprises, relative to the rotational axis of the disk: a radial web (4), blades (8) at the outer periphery of the web, a bore (5) at the inner periphery of the web, and a cylindrical side wall (12) extending the web in the vicinity of the outer periphery of same and having an air supply port (18), and - a centripetal air collection device (15). Advantageously, the device (15) comprises a cylindrical support (23) and at least one air supply tube (16), the inlet of which is turned towards the port (18) and the outlet of which is turned towards the bore (5) in the web, the disk comprising an inner radial flange (40) extending from the cylindrical side wall (12), the cylindrical support (23) of the device (15) being attached to said inner radial flange (40), and a ring (30) extending from the web (4), the cylindrical support (23) being centred on the ring (30).
The invention proposes a rotor (10) of an aircraft turbomachine having a main axis A, which comprises means (14) for modifying the critical speed of the rotor (10) depending on whether the rotational speed of the rotor (10) is lower or higher than a predefined rotational speed, comprising: a component (16) that is capable of occupying a first state or a second state depending on whether the rotational speed of the rotor (10) is lower or higher than the predefined rotational speed, each state of the component (16) corresponding to a critical speed of the rotor (10); and means (18) for driving the component (16) to one or the other of the two states thereof, depending on the rotational speed of the rotor (10), characterised in that the means (14) for modifying the critical speed of the rotor (10) further comprise a component (38) that engages with the drive means (18) and is capable of being deformed elastically between one or the other of two stable forms, each of which corresponds to a state of said component (16).
The invention relates to an assembly that includes a cryogenic fluid vessel of a space vehicle, and a thermal protection system for a cryogenic fluid vessel of a space vehicle (1), including: a shell (3) suitable for surrounding the cryogenic fluid vessel, the shell (3) being sized such as to accommodate an inner space (21) between the shell (3) and the vessel; and means (35, 36) for injecting a spray of a heat-transfer fluid into said inner space (21), characterised in that said heat-transfer fluid is injected into the inner space (21) in liquid state, at a temperature that is suitable for allowing the heat-transfer fluid to capture the heat flux reaching the cryogenic fluid vessel, causing said heat-transfer fluid to vaporise, the shell (3) including a plurality of openings suitable for allowing the heat-transfer fluid in gaseous form to exit said inner space (21) through the shell (3).
The present invention relates to an epicyclic reduction device (70) for rotating a first set of blades of a turbomachine, comprising a sun gear (74) centred on a longitudinal axis (12) of the turbomachine and connected to a rotor (76) of the turbomachine so as to be rotated; at least one planet gear (78) meshing with the sun gear; a planet gear carrier (80) rotationally bearing the planet gear and connected to a first set of blades (82) to rotate same; and an annulus gear (72) meshing with the planet gear; the sun gear being connected to the rotor by a first ball-type constant velocity joint (84).
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
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
The invention relates to the field of propulsion systems, in particular to expander-type rocket engines, in which at least one propulsion chamber (40) is supplied with at least one first propellant by a first pump (33) coupled with a first turbine (34) which is actuated by a partial expansion of said first propellant upstream from said propulsion chamber (40) after passing, downstream from the first pump (33), through a heat exchanger (23) heated by said propulsion chamber (40). A method for controlling such a propulsion system (10) according to the invention comprises the steps of determining a risk of instability of the propulsion system (10), and shedding a partial flow of said first propellant, in response to the determination of said risk of instability, via a bypass (28) located between the first turbine (34) and the propulsion chamber (40)
Method (S) for protecting against fire a fan casing (1) comprising a roughly cylindrical barrel (10) having a main direction extending along a longitudinal axis (X) and an upstream flange (20) extending radially with respect to the longitudinal axis (X) from an upstream end of the barrel (10), the fan casing (1) being made of a composite comprising a fibrous reinforcement densified by a matrix, said matrix being polymerized, the protection method(S) comprising the following steps: • - laying (SI) widths containing glass fibre pre-impregnated with a resin capable of affording the fan casing with thermal protection against fire on an upstream radial face (22) of the upstream flange (20), and • - polymerizing (S2) the resin in order to obtain a protective layer (2).
The invention pertains to a method for monitoring an aircraft engine retractable doors thrust reverser, the thrust reverser being a reverser having hydraulic actuators and being provided with contactors (3a, 4a, 5a, 3B, 4b, 5b, Sa, Sb) arranged so as to each return an item of information about the position of the doors, the engine comprising a computer (3) configured to carry out measurements (E1) of a parameter representative of the position of the contactors on the basis of the information returned by the contactors, characterized in that it comprises a calculation (E2) of one or more statistical indicators of the parameter measured and an analysis (E3) of the temporal evolution of the statistical indicator or indicators calculated. The invention extends also to a computer program for the implementation of this method.
The invention relates to a system (1) for generating electrical energy, pertaining to an aircraft, comprising: a primary energy source (P), an auxiliary electrical energy source (APU), and an emergency electrical energy source (EPU), characterised in that said auxiliary electrical energy source (APU) and said emergency electrical energy source (EPU) each comprise a fuel cell, respectively an auxiliary cell (20) and an emergency cell (30), and in that said system (1) comprises means (40) for recovering the heat released by the auxiliary cell (20) during the operation thereof so as to heat the emergency cell (30).
The invention relates to a method for fire protection (S) of a part (1) of a gas-turbine engine made of a composite material comprising a main fibrous reinforcement compregnated by a main matrix, the protection method (S) comprising the following steps: preforming (S1) a panel of prepreg (20) such as to grant same a shape corresponding to the shape of a surface (3) of the part (1) to be protected against fire, said panel of prepreg (20) comprising a secondary fibrous reinforcement compregnated by a secondary matrix; applying (S2) the panel of prepreg (20) thus preformed to the part (1); and securing (S3) the panel of prepreg (20) to the surface (3) by thermal treatment of the part (1) provided with said panel of prepreg (20) in order to obtain a fire-protection layer (2).
Assembly comprising, on a turbomachine: - a combustion chamber (30) having a longitudinal axis (30a) and comprising: internal and external longitudinal walls (90) having a first opening (90a) substantially transverse to said axis (30a), an outer casing (110) having a second opening (110a) likewise substantially transverse to said axis (30a), - a turbomachine sparkplug (150), - a device (130) for radially securing the sparkplug (150), which device (130) comprises a sparkplug adapter (160) fixed toward a first end to the outer casing (110), facing the second opening (110a) and through which adapter and casing the sparkplug (150) in question passes, and a sparkplug guide (190) kept in contact with the external longitudinal wall (90), facing the first opening (90a) and through which the sparkplug (150) passes to emerge in the combustion chamber (30), the sparkplug (150) and the sparkplug guide (190) each having a screw thread (103, 105), the screw threads (103, 105) being screwed together so as to stabilize the sparkplug (150) radially to the axis (30a) of the combustion chamber (30) with respect to said external longitudinal wall (90), and - in this condition, the sparkplug (150) and the sparkplug guide (190) are mounted freely with respect to the outer casing (110) in a direction substantially radial to the axis of the second opening (110a).
Aircraft propulsion assembly (10) comprising an engine (16), a nacelle (18) surrounding the engine, and a system for extinguishing a fire that may occur in the engine and/or in the nacelle, this extinguishing system comprising means (34) for supplying an extinguishant to at least one extinguishant distribution pipe (36) which opens into a cavity (32) of the engine and/or a cavity (26) of the nacelle, characterized in that it further comprises means (48) for supplying said at least one pipe with air so as to ventilate the or each cavity.
The invention proposes a test bench comprising at least one accelerometer (26, 28) which is linked by rigid mechanical linking to a support (12), and comprising a vibration chamber (38) able to convert an electrical signal into a mechanical vibration disseminated to the support and to the accelerometer by rigid mechanical linking between an oscillating part (40) of the vibration chamber and the support, in such a way as to use the support (10) as acoustic amplifier and to excite the accelerometer according to a predetermined mechanical wave.
The invention relates to a method for manufacturing a part coated with a protective coating, the method comprising the following step: forming a protective coating on the outer surface of a part by micro-arc oxidation treatment, the part comprising a niobium matrix which contains metal-silicide insertions, the current passing through the part being monitored during the micro-arc oxidation treatment in order to subject the part to a series of current cycles, the ratio (amount of positive charge applied to the part) / (amount of negative charge applied to the part) being, for each current cycle, 0.80 to 1.6.
Blade intended for a turbo machine impeller comprising N blades. The blade has, at one end, a platform (13) formed as an integral part with an airfoil of the blade. Over part of the axial extent of the blade, a section on a plane perpendicular to the axis (X) of the impeller of the flow path of the platform consists mainly of two straight-line segments (48ps, 48ss) arranged respectively on the two sides of the airfoil. These segments form, on each side of the airfoil, an angle of 90° - 180°/N with respect to the radial direction.
The invention relates to a method for producing multiple metal turbine engine parts, comprising steps consisting in: a) casting a metal alloy in a mould in order to produce a blank (3); and b) machining the blank in order to produce the parts, characterised in that the blank obtained by casting is a solid polyhedron with two generally trapezoidal opposing sides (30a, 30b), and the parts are machined in the blank.
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
B22D 13/00 - Centrifugal castingCasting by using centrifugal force
28.
TURBOMACHINE COMPONENT WITH NON-AXISYMMETRIC SURFACE
The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2) has a non-axisymmetric surface (S) bounded by a first and a second end plane (PS, PR) and defined by at least two class C construction curves each one representing the value of a radius of said surface (S) as a function of a position between the pressure face of the first blade (3I) and the suction face of the second blade (3E) in a plane substantially parallel to the end planes (PS, PR), these including at least one upstream curve and one downstream curve; each construction curve being defined by at least one pressure face control end point and one suction face control end point such that: - the tangent to the downstream curve at the suction face control end point 20 is inclined by at most 5°; - any other tangent to a construction curve at a control end point is inclined by at least 5°.
The present invention relates to a turbomachine component (1) or collection of components comprising at least a first and a second blade (3I, 3E) and a platform (2) from which the blades (3I, 3E) extend, characterized in that the platform (2), between the pressure face of the first blade (3I) and the suction face of the second blade (3E) has a non-axisymmetric surface (S) defining a plurality of fins (4) of substantially triangular section extending downstream of a leading edge (BA) of each of the blades (3I, 3E), each fin (4) being associated with a leading position and a trailing position on the surface (S), between which positions the fin (4) extends, such that: the leading position is situated at between 5% and 35% length relative to a chord of the blade (3I, 3E) extending from a leading edge (BA) to a trailing edge (BF) of the blade (3I, 3E); - the further a fin (4) is from the suction face of the second blade (3E), the further the leading position of said fin (4) is axially from the leading edge (BA) of the blades (3I, 3E).
The invention relates to a turbine engine compressor, in particular of an aeroplane turboprop or turbofan, including a stator comprising an annular casing and at least one annular row of variable-pitch vanes, each vane comprising a radially external end including a pivot mounted in an opening of the casing and connected by a linking member to a control ring (38) capable of pivoting axially relative to the casing, the linking member comprising a first end attached to the pivot of the vane and a second end comprising a pin inserted in a hole (52, 58) of the control ring (38), characterised in that at least one (58) of the holes (52, 58) of the control ring (38), which is used for inserting the pins of the linking members, has an oblong shape and extends in the circumferential direction such as to enable the pin to move into said oblong hole (58), during the rotation of the control ring (38).
The invention relates to a device (100) for cleaning, and in particular for degritting or desanding, a turbomachine module (110), characterised by comprising: (i) means (102, 104) for isolating bearings of the module, by containment in a closed enclosure (106); (ii) means (112) for overpressurising said enclosure; (iii) means (114) for stripping material deposited in the walls of annular recesses of the module, for example by spraying compressed air onto said walls; and (iv) means (116) for sucking up the material thus stripped.
Blade for a turbo machine impeller comprising a root, an air foil and a tip. The root and the tip comprise platforms having surfaces (15) on the side of the air foil, the surfaces respectively being referred to as the root and tip flow path. Each of these flow paths is made up of a pressure-face part and of a suction-face part which are situated respectively on the side of the pressure face and of the suction face and are separated by a crest curve (45, 65). Blade manufacture is made easier notably by virtue of the fact that any point on a first surface out of the pressure face and the suction face and any point on the root and tip flow path parts situated on the side of the first surface has a normal that makes an acute or right angle with respect to a direction referred to as the first direction of manufacture. Method for modelling the blade.
The invention relates to a device for supporting and centring a fuel injector in a turbine engine combustion chamber, which includes means for centring a fuel injector along an axis, which are movable in a plane that is radial to the centring axis (52) in supporting means intended for being attached to the bottom of an annular chamber (18). According to the invention, the centring means include at least two radially external tabs (54, 56) each inserted respectively in a circumferential recess (60, 64) of the supporting means, the device including circumferential abutment means (78, 80, 74, 76, 82, 86, 84, 88) of the radial tabs (54, 56) of the centring means in the circumferential recesses (60, 64), the circumferential abutment means being configured such as to enable a greater angular displacement of a first (54, 154) one of the radial tabs in a first circumferential recess (60) relative to a second (56, 156) one of the radial tabs in a second circumferential recess (64).
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE NORMALE SUPERIEURE DE CACHAN (France)
Inventor
Schneider, Julien
Hild, François
Leclerc, Hugo
Roux, Stéphane
Abstract
The invention relates to a method for characterising a part (10), including a step of obtaining an X-ray tomography image of the part, followed by a step (200) of correlating said image with a reference (20), characterised in that the correlation step (200) includes searching, in a predefined set (30) of transformations of X-ray tomography images, for a transformation (40) that minimises the difference (50) between the image and the reference in order to characterise (300, 350) the inside of said part (10).
G01N 23/04 - 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
G06K 9/64 - Methods or arrangements for recognition using electronic means using simultaneous comparisons or correlations of the image signals with a plurality of references, e.g. resistor matrix
The invention relates to a method for producing multiple metal turbine engine parts, comprising steps consisting in: a) casting a metal alloy in a mould in order to produce a blank (3); and b) machining the blank in order to produce the parts, characterised in that the blank obtained by casting is a solid polyhedron with two generally trapezoidal opposing sides (30a, 30b), and the parts are machined in the blank.
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
B22D 13/00 - Centrifugal castingCasting by using centrifugal force
36.
METHOD FOR PRODUCING A TURBINE ENGINE PART, AND RESULTING MOULD AND INTERMEDIATE BLANK
The invention relates to a method for producing multiple metal turbine engine parts, comprising steps consisting in: a) casting a metal alloy in a mould in order to produce a blank (3); and b) machining the blank in order to produce the parts, characterised in that the blank obtained by casting is a solid polyhedron having first and second sides (30a, 30b) and third and fourth sides (30c, 30d), in which the third and fourth sides extend between the first and second sides, flaring apart from the first side towards the second side, first at a first angle and subsequently at a second larger angle, and said at least one part is machined in the blank.
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
B22D 13/00 - Centrifugal castingCasting by using centrifugal force
37.
AIRCRAFT TURBOMACHINE COMPRISING A HEAT EXCHANGER OF THE PRECOOLER TYPE
The invention relates to an aircraft turbomachine (10) comprising a nacelle and an engine (12) comprising at least one outflowing jet of air, characterised in that a heat exchanger (20) of the precooler type for supplying air to the aircraft is mounted in the nacelle, said exchanger comprising a primary circuit, the inlet of which is connected to means for taking compressed air from the engine and the outlet of which is connected to means for supplying air to the aircraft, and a secondary circuit supplied with air taken from said air flow.
A gas turbine housing (100) made from an organic-matrix composite material comprising a reinforcement densified by an organic matrix delimits an inner volume. The housing comprises, on the inner face (101) of same, a structural portion (120) having a first face (120a) facing the inner face of the housing and a second opposing face (120b) defining a flow channel portion (102). Recesses (130) opening into the inner volume of the housing are present between the inner face (101) of the housing and the first face (120a) of the structural portion (120) facing said inner face of the housing. The recesses (130) allow the gases produced by the degradation of the resin of the housing in case of a fire to be discharged from the flow channel side.
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
The invention relates to a compressor shroud for an aircraft turbomachine, said shroud (20) being arranged between two rotating bladed wheels (16) and radially perpendicularly to a stator (12), and comprising a sealing device (30) comprising at least one sealing element (32), one of which is a downstream sealing element (32a) whereon a downstream-projecting structure (44) for driving and deviating air is provided, said structure being designed to axially straighten the discharge air coming from the sealing element (32a).
The invention relates to an annular combustion chamber in a turbomachine, comprising two coaxial walls of revolution, one an internal wall (12) and the other an external wall (14), between which emerge fuel injectors (19) each engaged in centring means (36) which are moveable in the radial direction in support means (38), the chamber also comprising means (50, 68) of axial retention in the upstream direction of the centring means. According to the invention, the means of axial retention in the upstream direction are fixed removably to at least one of the internal (12) or external (14) walls of revolution.
The invention proposes a guide arm for guiding at least one element having an elongated shape (20), corresponding to a set of cables and/or pipes. The arm comprises an inner cavity (62) opening on the outside of the arm at each of the ends thereof, and in which the elements having an elongated shape can extend. According to the invention, this structure more particularly comprises - a frame (8) comprising a beam (18) linked to means (30, 36, 44) for holding the elements having an elongated shape on the outside and along the beam, and - a cover (54) of which the walls (56, 58) cover the holding means of the frame, and are engaged with the beam, in such a way as to form the inner cavity in which the elements having an elongated shape extend, shock-absorbing means (68) being arranged between the means (30, 36, 44) for holding the elements having an elongated shape (20), and the longitudinal walls (56, 58), in such a way as to reduce and damp the movements of the means for holding in the cavity (62).
F01D 9/06 - Fluid supply conduits to nozzles or the like
F16L 3/10 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets substantially surrounding the pipe, cable or protective tubing divided, i.e. with two members engaging the pipe, cable or protective tubing
F16L 3/233 - Supports for pipes, cables or protective tubing, e.g. hangers, holders, clamps, cleats, clips, brackets specially adapted for supporting a number of parallel pipes at intervals for a bundle of pipes or a plurality of pipes placed side by side in contact with each other by means of a flexible band
42.
CUTTING TABLE FOR CUTTING A FIBROUS PREFORM OBTAINED BY THREE-DIMENSIONAL WEAVING AND CUTTING METHOD USING SUCH A TABLE
The invention concerns a cutting table (100) for cutting a fibrous preform obtained by three-dimensional weaving and comprising two portions that are linked together by at least one separating area and that have contours of different shapes, the cutting table comprising a plate (104) provided with a cavity (108) intended to receive, flat, one of the portions of the preform to be cut, sacrificial plates (110) intended to be interposed between the portions of the preform to be cut and to be secured to the plate, at least one cutting template (114) intended to be applied to the portion of the fibrous preform that is not positioned in the cavity, and means (118) for applying a compacting pressure to the cutting template. The invention also concerns a method for cutting a fibrous preform using such a cutting table.
The invention concerns a method and system for forecasting maintenance operations to be applied to an aircraft engine comprising a plurality of elements monitored by damage counters, each damage counter being limited by a corresponding damage ceiling, characterised in that it comprises: - processing means (7) suitable for simulating a consumption of said damage counters (C1-Cm) by iteratively pulling a series of simulation missions from a learning database (9) containing test missions, - processing means (7) suitable for determining, at each iteration, an accumulation of consumption of each of said damage counters until at least one counter counting damage related to a current simulation mission reaches the damage ceiling associated with same, - processing means (7) suitable for applying a maintenance strategy to said current simulation mission to determine maintenance indicators representative of the maintenance operations to be planned on the aircraft engine.
A system for regulating the flow of a propellant fluid for an electric thruster of a space vehicle, comprising a propellant fluid reservoir and a flow regulator mounted at the outlet of said reservoir, the flow regulator comprising a heating element controlled by a computer and designed to heat the propellant fluid and modify the physical properties thereof in order to vary the flow of propellant fluid exiting the reservoir, said system being characterised in that the computer further comprises a plurality of empirically determined empirical calibration curves defining the flow of propellant fluid depending on the intensity of heating and environmental parameters, such that said computer also performs a function of determining the flow of propellant fluid.
Aircraft turbomachine comprising at least one heat exchanger (40) and a gearbox (10) in a V-shaped overall configuration and comprising two lateral arms (20) joined together by a central joining piece, the heat exchanger being mounted between the arms of the gearbox.
Drained fluid evacuation stub (16) for a propulsion assembly (10), comprising a drained fluid storage cavity and at least one orifice (32) for evacuation of the fluids contained in said cavity, characterized in that it comprises means (36, 38) for detecting a pressure difference with the exterior of the stub and a component for purging the cavity which is movable between a first closed position of the evacuation orifice and a second release position of the orifice, the component being configured to move from the first to the second position when the pressure difference is greater than or equal to a predetermined value.
The invention relates to a device (110) for retaining drained fluids for a propulsive assembly, comprising a cavity for storing the drained fluids and two walls (118, 120) mounted at the opening of said cavity, the cavity having a fluid storage volume V1 when the device is in a substantially vertical position, and each wall being configured such as to define a fluid storage volume (V2 and V3 respectively) in the cavity when the device is in a substantially horizontal position, each of the volumes V2 and V3 being at least equal to the volume V1. The invention also relates to a propulsive assembly comprising a device for retaining drained fluids.
The present invention relates to a torque-measurement device for a turbine engine shaft (1) including a proof body (7) capable of being mounted on the shaft, characterised in that the proof body (7) forms a mounting for at least one acoustic-wave strain gauge (20) and is magnetised such as to allow the torque to be measured by magnetostrictive effect. The invention also relates to a method for calibrating the torque-measurement system including a first step of calibrating said device, the device being mounted on a shaft but outside of the engine, by applying reference torques to the shaft with the device and by establishing a rule regarding the relationship between the strain measured by said strain gauge and the actual torque applied, a step of mounting the shaft with the device inside the turbine engine together with placement of the magnetostrictive measurement system, the calibration of the first step being optionally reset with the engine stopped, and a step of establishing a calibration rule regarding the relationship between the torque measured by magnetostrictive effect and the reference torque provided by the strain gauges.
G01L 25/00 - Testing or calibrating of apparatus for measuring force, torque, work, mechanical power, or mechanical efficiency
G01L 1/16 - Measuring force or stress, in general using properties of piezoelectric devices
G01L 3/10 - Rotary-transmission dynamometers wherein the torque-transmitting element comprises a torsionally-flexible shaft involving electric or magnetic means for indicating
49.
PROPULSION ASSEMBLY COMPRISING A BOX FOR RETAINING DRAINED FLUIDS
The invention relates to a propulsion assembly comprising an engine surrounded by a nacelle (14), means for draining the fluids from the engine and means (18) for guiding said fluids to a retention box (17) arranged outside the engine and comprising a cavity (36) for storing drained fluids, characterised in that said retention box is fixed to a cowling of the nacelle and supported thereby. The invention also relates to a fluid retention box for a corresponding propulsion assembly.
A method for impregnation of a fibrous preform (10) by an impregnation composition (20), the method comprising the following step: • a) application of a liquid (30) onto a structure, the structure comprising: • - a chamber (2) in which a fibrous preform (10) to be impregnated is present, the chamber (2) being defined between a rigid support (3) on which the fibrous preform (10) is placed and a wall (4), the wall (4) comprising a face (4a) located facing the fibrous preform (10), and • - an impregnation composition (20), intended to impregnate the fibrous preform (10), the impregnation composition being present in the chamber (2), the liquid (30) being applied on the wall (4) of the opposite side of the chamber (2), the wall (4) being configured so that the face (4a) located facing the fibrous preform (10) retains its shape during application of the liquid (30), the applied liquid (30) enabling creation of a sufficient pressure to displace the wall (4) towards the rigid support (3) and impregnating the fibrous preform (10) with the impregnation composition (20).
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B30B 1/00 - Presses, using a press ram, characterised by the features of the drive therefor, pressure being transmitted directly, or through simple thrust or tension members only, to the press ram or platen
B29C 43/10 - Isostatic pressing, i.e. using non-rigid pressure-exerting members against rigid parts or dies
51.
METHOD FOR ADHERING PARTS AND DEVICE FOR IMPLEMENTING SAID METHOD
The invention relates to a method, and a device for implementing the method, for adhering at least two parts (10; 11; 11'; 11"; 11"') comprising the following steps: a) assembling at least two parts (10; 11; 11'; 11"; 11"') and at least one layer of adhesive (30; 30'; 30"; 30"'), the adhesive being configured such that the adhesive power thereof increases by heating due to the polymerisation thereof, at least one of the parts being made of a composite material or a metal material, the layer of adhesive (30; 30'; 30"; 30"') being present, once the assembly is carried out, between the parts (10; 11; 11'; 11"; 11"'), the assembled parts (10; 11; 11'; 11"; 11"') and adhesive being present in a chamber (2) defined by a wall (20; 20'), the wall (20; 20') having an inner surface (20a; 20'a) located opposite a first part (11; 11'; 11"; 11"') present between the layer of adhesive (30; 30'; 30"; 30"') and the wall (20; 20'); and b) applying a liquid (40; 40') to the wall (20; 20') on the side opposite the chamber (2), the applied liquid (40; 40') imposing pressure on the parts (10; 11; 11'; 11"; 11"'), the wall (20; 20') being configured such that, at least after applying the liquid (40; 40'), the inner surface thereof (20a; 20'a) has the same shape as the first part (11; 11'; 11"; 11"'); the adhering of the parts (10; 11; 11'; 11"; 11"') via the layer of adhesive (30; 30'; 30"; 30"') being obtained after implementing steps a) and b).
A method for producing a part by selective melting of powder, which involves: depositing a first layer (12) of a first powder (2) having, as the main element, a first element; depositing, on the first layer, a second layer (15) of a second powder (22) having, as the main element, a second element, different from the first element; and moving a first energy beam (11), for example a laser beam or an electron beam, over the second layer (15), the energy delivered by the first beam making it possible to initiate an exothermic reaction between the first element and the second element, the energy released by this exothermic reaction making it possible to melt the first and second layers (12, 15) together locally.
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
B22F 7/06 - Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting of composite workpieces or articles from parts, e.g. to form tipped tools
B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
B22F 5/04 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product of turbine blades
C22C 1/04 - Making non-ferrous alloys by powder metallurgy
C22C 1/05 - Mixtures of metal powder with non-metallic powder
B29C 67/00 - Shaping techniques not covered by groups , or
C22C 32/00 - Non-ferrous alloys containing at least 5% by weight but less than 50% by weight of oxides, carbides, borides, nitrides, silicides or other metal compounds, e.g. oxynitrides, sulfides, whether added as such or formed in situ
53.
METHOD FOR FIXING A TUBE TO A CONNECTOR, AND CONNECTING KIT
Method for fixing a tube (50) to a connector (101, 102), comprising the following steps: a) a connecting kit is provided, having a sleeve and a connector for connecting at least one pair of tubes, having an overall tubular shape; b) the connector is put in a first position in which the first end of the connector protrudes from the outside of the sleeve by way of a first end (10A) thereof; c) the first tube is then fixed to the first orifice in a sealed manner; d) the connector is moved with respect to the sleeve into a second position in which the first orifice is located inside the sleeve; and e) with the connector in the second position, the sleeve (10) is fixed at a sufficient distance from the connector (20). Connecting kit for implementing the method.
F16L 27/107 - Adjustable jointsJoints allowing movement comprising a flexible connection only the ends of the pipe being interconnected by a flexible sleeve
F28D 7/02 - Heat-exchange apparatus having stationary tubular conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall the conduits being helically coiled
54.
DEVICE FOR CENTRING AND GUIDING THE ROTATION OF A TURBINE ENGINE SHAFT INCLUDING IMPROVED MEANS FOR RETAINING THE EXTERNAL BEARING RING
The invention relates to a device (10) for centring and guiding the rotation of a turbine engine shaft, in which the outer ring (18) of a bearing is retained axially upstream and downstream by retaining means (52, 72) that engage with a bearing mounting (20) and with coupling means (29) including resiliently deformable means (32) connecting the outer ring to the bearing mounting, said retaining means being separate from a binding band (28) of the device. The invention also provides a method for assembling such a device in which the retaining means (52, 72) are pre-assembled with the coupling means (29) prior to the final assembly of the coupling means with the bearing mounting (20). The device and the method have the combined advantages of axially retaining the outer ring in two opposing directions and having a particularly straightforward assembly.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
The invention relates to a system for humidifying the air in a given space, characterised in that sais system includes a steam-generation unit including a stack (1) of elementary cells of a fuel cell, and terminal plates (2, 3) arranged on either side of said stack (1), in which system said terminal plates (2, 3) include a heat exchanger (21, 31) suitable for performing a heat transfer between said stack (1) and water flowing in said terminal plates (2, 3) such as to vaporise all or part of the water circulating in said terminal plates (2, 3), said system also including pipes suitable for conveying the generated steam into the given space such as to humidify the air in said space.
H01M 8/04 - Auxiliary arrangements, e.g. for control of pressure or for circulation of fluids
F22B 1/02 - Methods of steam generation characterised by form of heating method by exploitation of the heat content of hot heat carriers
F22B 1/18 - Methods of steam generation characterised by form of heating method by exploitation of the heat content of hot heat carriers the heat carrier being a hot gas, e.g. waste gas such as exhaust gas of internal-combustion engines
56.
DEVICE FOR GUIDING SYNCHRONIZING RING VANES WITH VARIABLE PITCH ANGLE OF A TURBINE ENGINE AND METHOD FOR ASSEMBLING SUCH A DEVICE
The invention relates to a device for guiding synchronizing ring vanes with variable pitch angle of a turbine engine, including a plurality of angular inner ring sectors placed end-to-end to form an inner ring (26), each inner ring sector including shafts (24) passing radially from one side of the inner ring sector to the other, a plurality of cylindrical bushes (22) which are each mounted in a shaft of the inner ring from the inside and which are each intended for receiving a guiding pivot (12) of a synchronizing ring vane (4), a plurality of angular reconstitution ring sectors which are placed end-to-end to form a reconstitution ring (36) and which are mounted radially from the inside on the inner ring, and a plurality of locking elements passing axially through the inner and reconstitution rings such as to assemble said rings together. The invention also relates to a method for assembling such a device.
The invention relates to a balanced turbine engine portion. Said portion comprises at least one angular section (21) arranged such as to form a balancing ring (20) centred on a ring axis (C). Said angular section (21) comprises a plurality of attachment elements (30), a bearing surface (5) with a shape that matches the balancing ring (20), the angular section (21) abutting with said bearing surface (5). Said portion also comprises a plurality of balance weights (40), each attached to the corresponding attachment element (30) of the angular section (21), at least one of said balance weights also being useful as an attachment means for attaching the angular section (21) to the bearing surface (5). The invention also relates to a turbine engine comprising such a balanced portion.
The invention relates to a fan, in particular for a small turbine engine such as a jet engine, having a hub ratio corresponding to the ratio of the diameter of the inner limit of the air intake section (26) at the radially internal ends of the leading edges of the fan blades (10), divided by the diameter of the circle through which the outer ends of the fan blades pass, which has a value of 0.25 to 0.27.
F01D 5/32 - Locking, e.g. by final locking-blades or keys
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream (26) at the radially inner ends of the leading edges of the fan blades (10), divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.
The invention relates to a method for determining the wear of a cutting tool flank, in which said wear is determined by means of a general function which calculates a typical length of said wear depending on at least one time-dependent variable. The invention also relates to a related device.
The invention relates to a lubrication device for a turbine engine, comprising an oil intake pipe (23) provided with a pump (24) for supplying oil and control means (30) located downstream from the supply pump (24), a supply pipe (26) intended for supplying oil to a member to be lubricated and a recirculation pipe (27) connected upstream from the supply pump (24), the control means (30) making it possible to direct all or part of the flow of oil from the intake pipe (23) towards the supply pipe (26) and/or towards the recirculation pipe (27).
A YSZ-type ceramic layer is deposited on a tie sublayer by thermal spraying using a plasma arc torch, said tie sublayer being itself deposited on the part to be protected. A sintering post treatment is carried out by means of a sweep of the ceramic layer by the beam of the plasma arc torch, the temperature at the point of impact of the beam at the surface of the ceramic layer (C) being, during this sweep, between 1300°C and 1700°C.
C23C 4/08 - Metallic material containing only metal elements
C23C 4/10 - Oxides, borides, carbides, nitrides or silicidesMixtures thereof
C23C 4/12 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
The invention relates to a turbomachine (30) comprising: a compressor stage and a turbine stage, each stage comprising at least one disk (42); and a tubular shaft (31) sleeve (33) extending along the axis (32) of the turbomachine, wherein the sleeve (33) comprises at least one tab (40) extending from an outer radial surface (41) of the sleeve and facing the disk (42), the tab (40) being designed to come into contact with the disk (42) when the sleeve (33) is in rotation about the axis (32) of the turbomachine.
A computer processor implemented process for identifying a template associated with a queried customer, such that a type and at least one subject of the template are selected based on (i) an influence score of the queried customer, (ii) an interest score of the queried customer, and (iii) a sensitivity profile of the queried customer. The process includes identifying sensitivity variables, influence variables, and interest variables, collecting data for these variables, processing sensitivity data to position the queried customer on a first map, processing an influence score and an interest score of the queried customer to position the queried customer on a second map, and identifying a template type for the queried customer, based on the second map and at least one template subject based on the first map. The process outputs a template with the identified template type and the identified at least one subject for the queried customer.
G06Q 10/04 - Forecasting or optimisation specially adapted for administrative or management purposes, e.g. linear programming or "cutting stock problem"
G06Q 10/06 - Resources, workflows, human or project managementEnterprise or organisation planningEnterprise or organisation modelling
G06Q 30/02 - MarketingPrice estimation or determinationFundraising
65.
MODULAR ENGINE, SUCH AS A JET ENGINE, WITH A SPEED REDUCTION GEAR
The present invention relates to an engine (1) with a modular structure comprising a plurality of coaxial modules (A, B, C) with, at one end, a first module (A) comprising a power transmission shaft (3) and a speed reduction gear (7), said power transmission shaft being driven via the speed reduction gear (7) by a turbine shaft (2) secured to one (C) of said coaxial modules that is separate from the first module, the speed reduction gear comprising a drive means (8 and 9) fixed to the turbine shaft (2) and to a journal (13) of a shaft of a low-pressure compressor rotor (1 a), characterized in that it comprises a first nut (16) for fastening the drive means to the journal and a second nut (14) for fastening the drive means to the turbine shaft.
The invention relates to a turbofan engine comprising a fan (S) driven, via a fan shaft (3) supported by at least two first bearings (11, 12), by a turbine shaft (4) supported by at least one second bearing (10) comprising a stationary ring (25) and a movable ring (26), said turbine shaft driving said fan shaft (3) through a device for reducing the speed of rotation (7), said device for reducing the speed of rotation and said first and second bearings being housed in a lubrication enclosure (E1) in which the shell comprises stationary portions and movable portions connected to one another by sealing means (29, 30, 31), said reducing device comprising an inducer (27) shaped so as to receive the torque transmitted by said turbine shaft via driving means (8, 9) connected to said movable ring, characterised in that the lubrication enclosure forms a coaxial ring with the turbine shaft and said driving means (8, 9) comprise a girth gear which is part of the movable sealing walls of the shell of the lubrication enclosure (E1).
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
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 3/067 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages having counter-rotating rotors
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
67.
DEVICE FOR SUPPLING PROPELLANT TO A PROPULSIVE ROCKET ENGINE CHAMBER
A device (20) for supplying propellant to at least one propulsive rocket engine chamber (19) of a propulsive assembly (21) comprises a main propellant reservoir (1, 2), a main supply pipe (3, 4) extending between the main reservoir (1, 2) and the propulsive chamber (19) and on which there is placed a main supply valve (VAH, VAO). The supply device additionally comprises an auxiliary reservoir (11, 12) for said propellant and an auxiliary supply pipe (13, 14) provided with an auxiliary supply valve (VH1, VO1), connecting the auxiliary reservoir (11, 12) to the main supply pipe (3, 4) downstream of the main supply valve (VAH, VAO).
The present invention relates to a bypass engine bearing holder (1) that holds an upstream bearing (6) and defines, with said upstream bearing, an oil chamber (100) and an air chamber (200), comprising a frusto-conical portion (11) defining an upstream bearing chamber (160) and a downstream inner chamber (150), and comprises an outer collar (13) connected, by a weld (135), to a flange (15) that extends outward from the frusto-conical portion (11). The outer collar (13) has a sealable gimlet (131) engaging with the upstream bearing (6) such as to seal the oil chamber (100). The bearing holder (1) comprises a plurality of oil recovery ducts (8) leading to the downstream inner chamber (150) and to the upstream bearing chamber (160). The oil recovery ducts (8) lead to the upstream bearing chamber (160), downstream from the weld (135) of the outer collar (13) on the flange (15), the weld (135) of the outer collar (13) being axisymmetric.
The invention relates to a multi-point fuel injection device (1) for an aircraft engine (M), comprising an inlet line (10), at least two injection lines (11, 12), and a purge line (14), a fuel distributor member (2) connected to each line and comprising a moveable element (22) which comprises an injection passage (223), in which the moveable element (22) additionally comprises a purge passage (226), and is configured to adopt a first range of positions in which the injection passage (223) interconnects the inlet line (10) and the injection lines (11, 12), and a second range of positions in which the injection passage (223) interconnects the inlet line (10) and at least a first injection line (11) while the purge passage (226) interconnects the purge line (14) and at least a second injection line (12), the device being characterized in that it additionally comprises an actuator adapted to move the moveable element into a safety position when a failure of the distribution member is detected, the injection passage (223) interconnecting, in this safety position of the moveable element (22), the inlet line (10) and the first injection line (11) while the purge passage (226) does not interconnect the purge line (14) to any of the injection lines (12).
F23R 3/34 - Feeding into different combustion zones
F23D 11/00 - Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
F23K 5/06 - Liquid fuel from a central source to a plurality of burners
F02C 9/34 - Joint control of separate flows to main and auxiliary burners
F02C 7/228 - Dividing fuel between various burners
F02C 7/232 - Fuel valvesDraining valves or systems
F02M 41/04 - Fuel-injection apparatus with two or more injectors fed from a common pressure-source sequentially by means of a distributor the distributor being spaced from pumping elements the distributor reciprocating
The invention concerns the field of turbomachines, and more specifically an annular element (13) of a turbomachine casing, comprising at least one inner face (14), delimiting a flow channel for the flow of a working fluid of the turbomachine, an outer face (15), and a damper (18), comprising at least one elastic coil (18a, 18b, 18c) clamped around a surface of revolution (15a) of the outer face (15), in such a way as to exert pressure against said surface of revolution (15a). The present invention also concerns the damping of a rotating wave deforming such an annular element (13) of a turbomachine casing, in which said rotating wave is damped by friction between said surface of revolution (15a) and the at least one coil (18a, 18b, 18c) of the damper (18).
The invention concerns a sealing system, in a cavity (C) under a stator (10), of a turbomachine vein (VC, VT), the cavity (C) being located between a stator (10) vane (PS) root (SI) and an additional rotor member (11), the root (SI) comprising two surfaces (21, 24a) each provided with an abradable coating (22, 32), the rotor member (11) being provided with first and second sealing elements (23, 33), disposed respectively facing the first and second surfaces (21, 24a), the first surface (21) and the first sealing element (23) forming a first sealing pair (20) and together delimiting a first leakage section, the second surface (24a) and the second sealing element (33) forming a second sealing pair (30) and together delimiting a second leakage section, one of the two pairs (20, 30) moving to a minimum leakage section when the other (30, 20) moves to a maximum leakage section, and vice versa.
The invention relates to the field of technical trials and specifically to a technical trial method for assessing at least one device operation parameter (X) across a series of operation increments, each one of which corresponds to a stable value of at least one operation rule of the device. The method includes at least: time sampling (S201) Tau values minus an operation parameter (X); filtering (S206) the sampled values in order to obtain a filtered signal (Xfilt) for each operation parameter (X); calculating (S208) the variance of the values sampled for each operation parameter (X) in a sampling window during said operation increment ; calculating (S205) the absolute value of the time derivative of the filtered signal (Xfilt) for each operation parameter (X); and changing (S211) the value of said at least one operation rule when, for each operation parameter (X), said variance of the sampled values and the absolute value of the time derivative of the filtered signal (Xfilt) are less than respective predetermined bottom thresholds.
The invention relates to the field of testing methods, and specifically to a method for technical testing of a device, including at least one operational step corresponding to a stable value of at least one operational setting of the device and/or of a test bench for the device. Said operational step ends within a maximum duration threshold if a criterion associated with a set of physical parameters collected during the operational step is met and a level of confidence associated with said set of physical parameters reaches at least one predetermined threshold.
The invention relates to a method for characterising a signal (U), comprising the following step: A) acquiring a signal (U); and characterised in that said method includes the following steps: B) determining a vibration rate (Formula (I)) of at least one portion of the signal (U) acquired in step A; C) determining a level of variations (Formula (II)) of at least one portion of the signal (U) acquired in step A; D) determining information (R1, R2) that characterises the signal by applying a function using at least said rate of vibrations (Formula (I)) and said level of variations (Formula (II)) as arguments. The invention also relates to a signal characterisation device for implementing the method.
The invention relates to a device (1) for replacing machining inserts (900) on a tool (800) including a body (810) and a head (820) supporting at least one machining insert (900), each insert being maintained on the head (820) by a screw. The device (1) includes: a positioner (50) comprising a supporting element (59) capable of supporting the body (810) of the tool; a screwing station having a screwdriver (60) capable of screwing and unscrewing the screws, the positioner (50) being capable of moving the tool relative to the screwing station; a gripping device (70) capable of gripping and placing an insert (900); a conveyor (500) comprising a plurality of insert containers (510), along which a first station (100) having the gripping device (70), a second station (200) having a mechanism (20) for rotating the inserts on the axes thereof, a third station (300) having a mechanism (30) for unloading the inserts, and a fourth station (400) having a mechanism (40) for supplying inserts are distributed; a transport mechanism (80) capable of moving the gripping device (70) between the positioner (50) and the first station (100) of the conveyor (500); and a control centre (600) capable of automatically controlling one or more of the mechanisms and devices and/or the conveyor of the replacement device (1).
The invention relates to the field of rocket engines, and specifically to a propulsion assembly (10) including a first tank (11), for a first liquid propellant, a second tank (12), for a second liquid propellant, a pressurisation device (60) configured to maintain a substantially higher pressure in the second tank (12) than in the first tank (11), and a single-shaft turbopump (30), including a turbine (33), a first pump (31) and a second pump (32). The turbine (33) is actuated by the expansion of the first propellant after passing through a regenerative heat exchanger (44), and in turn actuates, via the single rotary shaft (34) of the turbopump (30), said first and second pumps (31, 32), pumping said first and second propellants, respectively, in order to supply a thrust chamber (40).
The invention relates to a turbine engine, comprising two structural annular casings (16, 22) connected to one another by means (40, 54) for absorbing stresses from the thrust of the engine, which include connecting rods (54), characterised in that said thrust-absorbing means also include at least one accessory gearbox (40) which is attached to a first one of said casings (16) and which is connected by said connecting rods to the other one of said casings (22).
The invention concerns a method for detecting a failure in a fuel return valve of an aircraft engine fuel circuit, the fuel circuit comprising a fuel tank (10); an engine fuel system (20) connected to the fuel tank, said engine fuel system being capable of delivering a flow of fuel to the engine depending on a speed of said engine; a fuel return pipe (2) connected between the engine fuel system (20) and the fuel tank; a fuel return valve (30) arranged to switch between an open position and a closed position, said valve being capable of blocking the fuel return pipe (2), when in the closed position, and of bringing the fuel return pipe into communication with the fuel tank, when in the open position; the method comprising the following steps implemented in a computer: (EI) measuring a pressure of the flow of fuel from the fuel tank; and if the measured pressure is lower than a predefined threshold, (E2) measuring the engine speed.
The invention concerns a method for detecting a failure in a fuel return valve of an aircraft engine fuel circuit, said fuel circuit comprising: a fuel tank (10); an engine fuel system (20) connected to the fuel tank (10), said fuel system comprising a high-pressure pump (21) delivering a flow rate Q, depending on an engine speed of said engine, to an actuating cylinder (23) capable of actuating variable geometries, a cutoff valve (22) capable of feeding the actuating cylinder and disposed in a feed pipe (28) feeding the engine; a fuel return pipe (2) connected, on the one hand, downstream of the high-pressure pump (21) and upstream of the cutoff valve (22) and, on the other, to the fuel tank (10); a fuel return valve (30) arranged to switch between an open position and a closed position, said fuel return valve (30), in the closed position, being capable of blocking the fuel return pipe (2) and, in the open position, of bringing the fuel return pipe into communication with the fuel tank (10); the method comprising the following steps performed in a computer (28): starting (E1) the engine at an engine speed NO; increasing (E3) the engine speed until flow rate Q reaches a predefined value QO sufficient for opening the cutoff valve; measuring (E4, E5) the position of the actuating cylinder and engine speed N corresponding to the opening of said cutoff valve (22).
A fiber preform for a hollow turbine engine vane, the preform comprising a main fiber structure obtained by three-dimensional weaving and including at least one main part (41), wherein the main part (41) extends from a first link strip (44p), includes a first main longitudinal portion (46) suitable for forming essentially a pressure side wall of an airfoil, then includes an U-turn bend portion (45) suitable for forming essentially a leading edge or a trailing edge of the airfoil, then includes a second main longitudinal portion (47) facing the first main longitudinal portion (46) and suitable for forming essentially a suction side wall of the airfoil, and terminating at a second link strip (44q), wherein the first and second link strips (44p, 44q) are secured to each other and form a link portion (44) of the main fiber structure, and wherein the main longitudinal portions (46, 47) are spaced apart so as to form a gap between said main longitudinal portions (46, 47) suitable for forming a hollow in the airfoil.
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]
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
81.
AUTOGENOUS PRESSURISATION DEVICE FOR A PROPELLANT RESERVOIR
An autogenous pressurisation device (50) for a main propellant reservoir (10, 30), comprises a pressurisation pipe (13, 33) connected to the main reservoir (10, 30) for injecting the propellant into said main reservoir, a pressurisation valve (13a, 33a) disposed on the pressurisation pipe (13, 33), and a heater (17, 37) for heating the propellant upstream from the pressurisation valve (13a, 33a). The pressurisation device (50) comprises a buffer reservoir (15, 35) connected to the pressurisation pipe (13, 33) upstream from the pressurisation valve (13a, 33a).
Fibrous preform for a hollow blade of a turbomachine, such a hollow blade and a method for manufacturing such a hollow blade. The preform comprises a first fibrous structure (41') obtained by three-dimensional weaving and comprising at least one main longitudinal portion (46, 47) able to form essentially an air foil pressure-face wall, a second fibrous structure (42') obtained by three-dimensional weaving and comprising at least one main longitudinal portion able to form essentially an air foil suction-face wall; the first and second fibrous structures (41, 42) furthermore each comprising a first connecting zone (43), extending along the front edge of their respective main longitudinal portion (46), secured to one another and forming a first connecting portion (43) of the preform and a second connecting zone (44), extending along the rear edge of their respective main longitudinal portion (46, 47), secured one to the other and forming a second connecting portion of the preform (44); the main longitudinal portions (46, 47) of the first and second fibrous structures (41', 42') being non-attached so as to leave between said main longitudinal portions (46, 47) a space (D2) able to form an air foil hollow.
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]
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
83.
PROPULSION CHAMBER FOR A ROCKET AND METHOD FOR PRODUCING SUCH A CHAMBER
The invention relates to a propulsion chamber (100) for a rocket and a method for producing such a propulsion chamber (100). The propulsion chamber (100) comprises a combustion chamber (12), one wall of the combustion chamber (12) comprising a cooling circuit (14) in which a first propellant (16) flows. According to the invention, an envelope made from a thermostructural composite material (24) is attached externally to said combustion chamber (12) and comprises a divergent portion (24c) extending beyond the lower end (12ee) of the combustion chamber (12) and at least a part of said envelope made from thermostructural composite material (24) is covered with an external reinforcement envelope (26) with high radial strength to contain deformations of the combustion chamber (12) and of said envelope made from thermostructural composite material (24), the envelope made from thermostructural composite material (24) and the external reinforcement envelope (26) forming a unitary assembly (28).
The invention relates to a method for monitoring a locking system (1) comprising N locks (2a, 2b, 2c, 2d, 2e). Each lock (2a, 2b, 2c, 2d, 2e) is monitored by two locking sensors (3a, 4a, 3b, 4b, 3c, 4c, 3d, 4d, 3e, 4e). Each locking sensor is capable of indicating if the lock that said locking sensor is monitoring is in a locked or unlocked state. Each locking sensor can be in a valid or invalid state. The method comprises the following steps: - determining the state of the locking system (1) on the basis of the state of the locks detected by the locking sensors; and - determining a reliability level associated with the state of the locking system on the basis of the number of valid locking sensors monitoring the locks that are in the same state as the locking system.
The invention relates to a method for measuring pollutants contained in an exhaust stream exiting an engine, comprising the steps consisting of: Positioning a probe such that a sampling opening of said probe is positioned on a sampling surface provided at the outlet of the engine in the exhaust stream, and sampling the exhaust stream with said probe; Activating an analysis unit coupled with the probe in order to acquire characteristic data of the exhaust stream sampled by the probe; Controlling a movement of the probe to impart a continuous movement of the sampling opening along a specific trajectory on the sampling surface with constant surface scanning per unit of time, while continuing the sampling and the acquisition of characteristic data of the exhaust stream sampled by the probe; - Processing the data acquired by the analysis unit to measure the pollutants present in the exhaust stream. The invention also relates to a device for implementing this measuring method.
SOCIETE LORRAINE DE CONSTRUCTION AERONAUTIQUE (France)
Inventor
Poisson, Mathieu, Ange
Orcel, Stéphane
Glemarec, Guillaume
Pacary, Jean-Luc
Abstract
Turboprop air intake A turboprop (110), comprising a rotary propeller (112) upstream from an engine (114) and an air intake (116) that is not coaxial to the propeller, said air intake defining a conduit (119) for supplying air to the engine and further defining a bypass (124) to said conduit, the bypass comprising an outlet (126) oriented substantially axially towards the downstream of the engine, the turboprop further comprising a nacelle (130) surrounding the engine and the air intake, characterised in that the air intake is secured to a housing (123) of the engine and is not rigidly connected to the nacelle, so as to allow, during operation, relative movements between the air intake and the nacelle, said outlet (126) being connected by a flexible link (140) to an intake of an air circuit carried by the nacelle.
The invention relates to a method and system for monitoring an aircraft engine (2), comprising: - acquisition and processing means (11) configured to collect a time signal for a residual temperature margin at the exhaust gas outlet from said aircraft engine (2), - acquisition and processing means (11) configured to smooth said time signal, thus forming a curve representative of said residual temperature margin, - acquisition and processing means (11) configured to identify descending pieces of said first curve, - acquisition and processing means (11) configured to build a second curve by concatenating said descending pieces, said second curve being continuous while being restricted to said descending pieces of said first curve, - acquisition and processing means (11) configured to build a prediction model from said second curve to determine at least one failure prognosis indicator.
B64F 5/00 - Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided forHandling, transporting, testing or inspecting aircraft components, not otherwise provided for
88.
DEVICE FOR TRANSFERRING OIL BETWEEN TWO REPOSITORIES ROTATING RELATIVE TO EACH OTHER, AND PROPELLER TURBOMACHINE FOR AN AIRCRAFT WITH SUCH A DEVICE
The device (20) comprises two outer and inner concentric rings (22, 23), one of which is connected to an oil supply from one of the repositories, the other ring being connected to the other repository, the oil flowing between said rings, and bearings between the rings in order to change repositories between the two rings. According to the invention, the device (20) further comprises a flexible means (31) forming a shock absorber, provided between a first of said rings and an intermediate ring (41) that is separated from a second of said rings by said bearings (25), said flexible means (31) defining a deformable sealed chamber (32) in which oil travels between the two repositories.
The invention relates to a method for the acoustic analysis of a machine (M), comprising the acquisition of at least one acoustic signal supplied by at least one microphone (7) positioned in the machine, characterised in that it comprises the following steps: separation of at least one acoustic signal into a plurality of sound sources, said signal being modelled as a mixture of components, each one corresponding to a sound source; for at least one separate sound source, determination of a characteristic acoustic signature; comparison of at least one characteristic acoustic signature with at least one reference acoustic signature recorded in a reference database (5).
G01M 15/12 - Testing internal-combustion engines by monitoring vibrations
G01N 29/14 - Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic wavesVisualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object using acoustic emission techniques
G01N 29/44 - Processing the detected response signal
G01N 29/46 - Processing the detected response signal by spectral analysis, e.g. Fourier analysis
90.
METHOD FOR DESIGNING A VALVE AND METHOD FOR PRODUCING A VALVE
The invention comprises a method for designing a valve (12) having an opening (36) of which the cross section (S) depends on the position of the valve shutter (30) and can therefore be controlled. The method comprises the following steps: • a) determining a desired law (S *(t)) of variation of the valve passage cross section, which defines a desired variation of this cross section as a function of time; • b) setting a predefined law (P(t)) of movement for the shutter; and • c) defining the shape of the fluid passage opening such that, at any time (t), if the shutter moves according to said predefined law (P(t)) of movement, the passage cross section (S(t)) remains equal to the desired cross section (S (t)). The method can be used to design a valve that is easy to control.
F16K 3/26 - Gate valves or sliding valves, i.e. cut-off apparatus with closing members having a sliding movement along the seat for opening and closing with sealing faces shaped as surfaces of solids of revolution with cylindrical valve members with fluid passages in the valve member
The present invention concerns a turbomachine part (1) comprising at least first and second blades (3, 31, 3E), and a platform (2) from which the blades (3, 31, 3E) extend, characterised in that the platform (2) has a non-axisymmetric surface (S) limited by first and second end planes (PS, PR), and defined by at least three construction curves (PC-A, PC-C, PC- F) of class C1 each representing the value of a radius of said surface (S) on the basis of a position between the lower surface of the first blade (31) and the upper surface of the second blade (3E) according to a plane substantially parallel to the end planes (PS, PR), including: - a first curve (PC-C) that increases in the vicinity of the second blade (3E); - a second curve (PC-F) disposed between the first curve (PC-C) and a trailing edge (BF) of the first and second blades (3, 31, 3E), and that decreases in the vicinity of the second blade (3E); - a third curve (PC-A) disposed between the first curve (PC-C) and a leading edge (BA) of the first and second blades (3, 31, 3E), and having a minimum at the second blade (31).
The invention relates to a combustion chamber for a turbine engine, including an annular bottom wall (18) provided with injection systems (20) each centred on a respective axis (24) and each having an upstream end forming a socket (26') intended for receiving a head of a fuel injector, and an annular fairing (40') covering said bottom wall (18) and including injector-passage openings (42) arranged respectively opposite said injection systems (20), wherein said annular fairing (40) comprises air-intake openings separated from said injector-passage openings (42), and said socket (26') of each injection system passes through the corresponding injector-passage opening (42) and includes, at the upstream end thereof, a flange (62) having a free end (64) separated from said axis (24) of the injection system by a first distance (d1) which is greater than a second distance (d2) separating an edge of said injector-passage opening and said axis.
Jet pump (10) for a device for depressurizing a turbomachine lubricating chamber, comprising a tubular body (12), a first injection duct (14) which opens into the body (12) via a first outlet nozzle (18), a second injection duct (16) which opens into the body (12) via a second outlet nozzle (20) which surrounds the first outlet nozzle (18), the downstream end (20a) of the second outlet nozzle (20) being situated axially level with the downstream end (18a) of the first outlet nozzle (18), in which jet pump the upstream end (18b) of the first nozzle (18) comprises an axial orifice (30) centred on the main axis of the body (12), the jet pump (10) comprising means (34) for either plugging or not plugging said axial orifice (30) of the first outlet nozzle (18).
In order to improve the cooling of an annular wall (13) of a turbomachine combustion chamber provided with microperforations (53) and, in particular, the cooling of a region of the wall facing a wake (52) caused by an ignition plug, deflector means (60, 68) are proposed, these being designed to deflect the air (34') bathing the ignition plug towards a mid plane (P) of the wake (52) and in the direction of the annular wall (13) of the combustion chamber so as to increase the pressure of the air within the wake (52) near the annular wall (13).
The invention relates to a compacting assembly comprising a forming mould (24) defining an upwardly open housing that can receive a previously cut woven preform (10a), and a vertically mobile compacting tool (128), and forming, with the forming mould (24), an assembly for compacting said preform previously placed in the housing. The compacting tool (128) comprises at least one foot portion (128A). The compacting tool comprises at least three separate compacting blocks (1281-1287). The invention is applicable to the production of composite fan blades for turbomachines.
B29C 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 43/36 - Moulds for making articles of definite length, i.e. discrete articles
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
96.
COMPACTING ASSEMBLY AND METHOD FOR MANUFACTURING A TURBOMACHINE COMPOSITE BLADE
Compacting assembly comprising a shaping mould (24) delimiting a housing open at the top able to receive a precut woven preform (10a), and a compacting tool (128) that is able to move vertically and forms, with the shaping mould (24), a compacting assembly for compacting said preform placed beforehand in the housing. The compacting tool (128) comprises at least one root portion (128A). Application to the manufacture of turbomachine composite blades.
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29C 70/48 - 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
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
The invention relates to a fuel injector (10) such as an injector for an annular combustion chamber of a turbomachine, comprising a downstream head (16) having a central outlet (22) and an annular peripheral outlet (24) surrounding the central outlet (22), and an injector arm (12) upstream of the head (16) comprising a coaxial central channel (18) and a coaxial annular channel (20), characterised in that the central channel (18) is in fluid communication with the peripheral outlet (24) and the annular channel (20) is in fluid communication with the central outlet (22).
A fibrous structure (200) comprises a preform portion (210) formed as a single piece by three-dimensional weaving between a first plurality of layers of threads and a second plurality of layers of threads, the preform portion corresponding to all or part of a fibrous reinforcement preform for a component made of composite. The fibrous structure (200) comprises, outside of the preform portion (210), one or more layers of two-dimensional woven fabric (220a, 220b), each layer of two-dimensional woven fabric grouping together the threads (2010a) of one same layer (201a) belonging at least to the first plurality of layers of threads and situated outside of the preform portion (210).
D03D 25/00 - Woven fabrics not otherwise provided for
B29C 70/22 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least two directions forming a two dimensional structure
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
99.
CLEARANCE LIMITING DEVICE FOR A TURBOMACHINE AND METHOD FOR REMOVING A CARTRIDGE
The invention concerns a device (20, 30) for limiting clearance between a rotor (3) and a stator (2) of a turbomachine comprising: • - a cartridge (21) comprising at least one right-hand lateral lug and one left-hand lateral lug (22); • - a mounting support (SI) forming a transverse cross-section of a ring and comprising a right-hand side wall (14) having at least one right-hand lateral groove (11, 12) and a left-hand side wall (15) having at least one left-hand lateral groove (11, 12), the right-hand lateral groove being intended to receive a right-hand lateral lug (22) and the left-hand lateral groove being intended to receive a left-hand lateral lug (22), each lateral groove (11, 12) comprising: • o a first part (111, 114) allowing a radial displacement of a lug (22), the dimension of said first part defining a radial clearance area of the cartridge (21); • o a second part (110, 113) allowing a circumferential displacement of a lug (22); at least one of the two lateral grooves also comprising a third part (112) allowing the radial insertion of a lug (22).
F01D 11/16 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
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 - CasingsCasing parts, e.g. diaphragms, casing fastenings
The invention relates to a propulsion assembly (1, 201) for a rocket, comprising a reservoir (2) configured to contain a propellant, an engine comprising a combustion chamber (9), a propellant supply line (11) that extends between the reservoir (2) and the combustion chamber (9) and on which a sectional valve (24) is positioned, and a heater (15) whereof an inlet is connected to the supply line (11) and an outlet is connected to the reservoir (2). The inlet of the heater comprises an inlet line (13a) connected to the supply line (11) downstream from the sectional valve (24) on the one hand and to a neutral fluid intake (31) on the other hand.
