B29B 11/14 - Making preforms characterised by structure or composition
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
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
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B29C 70/48 - 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 70/86 - Incorporating in coherent impregnated reinforcing layers
B29K 105/08 - Condition, form or state of moulded material containing reinforcements, fillers or inserts of continuous length, e.g. cords, rovings, mats, fabrics, strands or yarns
B29L 31/30 - Vehicles, e.g. ships or aircraft, or body parts thereof
B32B 5/26 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by the presence of two or more layers which comprise fibres, filaments, granules, or powder, or are foamed or specifically porous one layer being a fibrous or filamentary layer another layer also being fibrous or filamentary
D03D 15/37 - Woven fabrics characterised by the material, structure or properties of the fibres, filaments, yarns, threads or other warp or weft elements used characterised by the structure of the fibres or filaments with specific cross-section or surface shape
F16C 7/02 - Constructions of connecting-rods with constant length
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
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
B29C 70/44 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
B29C 70/48 - 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 70/86 - Incorporating in coherent impregnated reinforcing layers
B29K 105/08 - Condition, form or state of moulded material containing reinforcements, fillers or inserts of continuous length, e.g. cords, rovings, mats, fabrics, strands or yarns
B29L 31/30 - Vehicles, e.g. ships or aircraft, or body parts thereof
B32B 5/26 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by the presence of two or more layers which comprise fibres, filaments, granules, or powder, or are foamed or specifically porous one layer being a fibrous or filamentary layer another layer also being fibrous or filamentary
D03D 15/37 - Woven fabrics characterised by the material, structure or properties of the fibres, filaments, yarns, threads or other warp or weft elements used characterised by the structure of the fibres or filaments with specific cross-section or surface shape
F16C 7/02 - Constructions of connecting-rods with constant length
B29B 11/04 - Making preforms by assembling preformed material
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
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 70/48 - 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 70/86 - Incorporating in coherent impregnated reinforcing layers
B29K 105/08 - Condition, form or state of moulded material containing reinforcements, fillers or inserts of continuous length, e.g. cords, rovings, mats, fabrics, strands or yarns
B29L 31/30 - Vehicles, e.g. ships or aircraft, or body parts thereof
B32B 5/26 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by the presence of two or more layers which comprise fibres, filaments, granules, or powder, or are foamed or specifically porous one layer being a fibrous or filamentary layer another layer also being fibrous or filamentary
COMPAGNIE GENERALE DES ETABLISSEMENTS MICHELIN (France)
SAFRAN ELECTRONICS & DEFENSE (France)
Inventor
Neba, Eric Carin
Abstract
The invention relates to a method and a system for determining the temperature in a mounted aircraft tyre. The system is characterized in that it comprises two temperature sensors installed inside the tyre.
B60C 23/04 - Signalling devices actuated by tyre pressure mounted on the wheel or tyre
B64C 25/36 - Arrangements or adaptations of wheels, tyres or axles in general
B64F 5/60 - Testing or inspecting aircraft components or systems
G01K 1/02 - Means for indicating or recording specially adapted for thermometers
G01K 3/06 - Thermometers giving results other than momentary value of temperature giving mean valuesThermometers giving results other than momentary value of temperature giving integrated values in respect of space
G01L 11/00 - Measuring steady or quasi-steady pressure of a fluid or a fluent solid material by means not provided for in group or
G01L 17/00 - Devices or apparatus for measuring tyre pressure or the pressure in other inflated bodies
6.
MACHINERY FOR WINDING A FIBROUS TEXTURED MATERIAL, AND ASSOCIATED WINDING METHOD
The present invention relates to machinery (10) for winding a fibrous textured material (12, 120, 220) about a support (14, 140, 240), allowing the gradual modification of the shape of the support as the winding progresses, and comprising compaction devices (20,329) for the purpose of reducing the wrinkling of the textured material obtained after compaction.
The present disclosure relates to a method for manufacturing discs for turbomachines, the method comprising: obtaining a nickel-based alloy powder; shaping the powder in order to obtain a disc; characterised in that the step of obtaining a powder comprises: atomising a nickel-based alloy by electrode induction gas atomisation (EIGA) from a nickel-based alloy electrode, resulting in a raw powder; and sieving the raw powder under an inert atmosphere or under a vacuum to a particle size of between 150 µm and 50 µm, for example 125 µm or 75 µm, resulting in the nickel-based alloy powder.
Système d'écrouissage (10) d'une surface d'une pièce métallique comprenant un support de la pièce métallique, un bras articulé robotisé (14) portant un dispositif d'écrouissage (16) de la pièce métallique, un dispositif de mesure de la température (20) sur la surface de la pièce métallique, un dispositif de création d'un champ thermique dans la pièce métallique comprenant un dispositif de chauffe localisée à distance de la pièce métallique, le dispositif de création du champ thermique étant configure pour fonctionner en même temps que le dispositif de mesure de la température et le dispositif d'écrouissage.Procédé d'écrouissage d'une surface d'une pièce métallique au moyen du système d'écrouissage (10)
This method for stopping at least one aircraft turbogenerator (1) comprises: - controlling the stopping (E1) of the turbogenerator (1); - passing from the nominal operating speed (Nref) of the power shaft (3, 12) to a first operating speed (N1) lower than the nominal speed (Nref), for a first predetermined duration (t2); - controlling the extinction of the combustion chamber (6) of the gas turbine (2); - maintaining the rotation of the gas turbine at a second speed (N2) for a second predetermined duration (t3), the power shaft (3, 12) being at a second speed (N2) lower than the first operating speed (N1) and, - controlling the stopping of the reversible electric machine (7) in order to no longer drive the power shaft (3, 12), in order to cause a progressive stopping (E9, E10) of the rotation of the gas turbine (2).
The invention relates to a device comprising: - a mount (10), - a frame (20) mounted so as to rotate relative to the mount; - a control arm (40) mounted so as to rotate on the frame (20); a roll spider (80) mounted so as to rotate on the frame (20); - a pitch spider (100) mounted so as to rotate on the control arm (40); - a roll motor (120) which is fixed relative to the frame (20) and which comprises a roll shaft (121) mounted so as to rotate on the roll spider (80); - a pitch motor (140) which is fixed relative to the frame (20) and which comprises a pitch shaft (141) mounted so as to rotate on the pitch spider (100). Drawing_references_to_be_translated
G05G 9/047 - Manually-actuated control mechanisms provided with one single controlling member co-operating with two or more controlled members, e.g. selectively, simultaneously the controlling member being movable in different independent ways, movement in each individual way actuating one controlled member only in which movement in two or more ways can occur simultaneously the controlling member being movable by hand about orthogonal axes, e.g. joysticks
11.
METHOD FOR DETERMINING AN EFFICIENCY FAULT OF AN AIRCRAFT TURBOSHAFT ENGINE MODULE
The invention relates to a method for determining an efficiency fault (R11-R15) of at least one module (11-15) of a turboshaft engine (T) of an aircraft (A), the method for determining comprising: ? A step of determining an estimated real mapping (CARE), ? A step of determining real indicators (IRE) from the estimated real mapping (CARE), ? A step of determining (E3) a plurality of simulated mappings from a simulation of a theoretical model of the turboshaft engine (T) for different efficiency configurations, ? A step of determining (E4) simulated indicators (ISx) for each simulated mapping (CARSx), ? A step of training (E5) a mathematical model (CLASS) by coupling the simulated indicators (ISx) with efficiency configurations (CR), and ? A step of applying (E6) said mathematical model (CLASS) to the real indicators (IRE) so as to deduce therefrom a real efficiency configuration (CR).
Panneau acoustique comprenant deux peaux (2, 3) sensiblement parallèles entre lesquelles sont disposées des cavités formant des résonateurs de Helmholtz, l'une desdites peaux (2, 3) étant percées d'orifices (8a, 8b, 9) débouchant chacun dans une desdites cavités et formant un col des résonateurs, caractérisé en ce qu'une ou plusieurs desdites cavités sont formées par une structure (4) creuse en forme de prisme droit à base tri angulaire.
B32B 3/12 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a discontinuous layer, i.e. apertured or formed of separate pieces of material characterised by a layer of regularly-arranged cells whether integral or formed individually or by conjunction of separate strips, e.g. honeycomb structure
B32B 3/24 - Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shapeLayered products comprising a layer having particular features of form characterised by a discontinuous layer, i.e. apertured or formed of separate pieces of material characterised by an apertured layer, e.g. of expanded metal
B32B 27/08 - Layered products essentially comprising synthetic resin as the main or only constituent of a layer next to another layer of a specific substance of synthetic resin of a different kind
B32B 37/14 - Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding characterised by the properties of the layers
L'invention concerne une composition pour fabrication additive caractérisée en ce qu'elle comprend :- un mélange polymérique comprenant une polyétheréthercétone et un polyétherimide, et- des fibres de carbone.
B29C 64/165 - Processes of additive manufacturing using a combination of solid and fluid materials, e.g. a powder selectively bound by a liquid binder, catalyst, inhibitor or energy absorber
B33Y 70/10 - Composites of different types of material, e.g. mixtures of ceramics and polymers or mixtures of metals and biomaterials
C08K 3/013 - Fillers, pigments or reinforcing additives
L'invention concerne un procédé de formation d'une couche d'alumine (203) à la surface d'un substrat métallique (201) en alliage comprenant de l'aluminium, le procédé comprenant au moins :- le dépôt d'une première couche d'aluminium sur une surface du substrat métallique,- le dépôt d'une deuxième couche par dépôt en phase vapeur sur la première couche, la deuxième couche comprenant de l'aluminium, un halogène et de l'oxygène, et- le traitement thermique du substrat revêtu des première et deuxième couches sous atmosphère oxydante afin de former la couche d'alumine à la surface du substrat métallique.
C23C 28/00 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and
The invention relates to a turbine vane comprising a root carrying a blade terminated by a tip in the form of a squealer tip, the blade having an intrados wall and an extrados wall, as well as a leading edge, a trailing edge, and a tip wall delimiting a bottom of the squealer tip, by which the intrados wall is connected to the extrados wall, said blade also comprising: - a serpentine median circuit (28), including a first radial pipe (41) that collects air at the root and is connected by a first bend (46) to a second radial pipe (42) that is connected by a second bend (47) to a third radial pipe (43); - a cavity (36) under the squealer tip running along the extrados wall (21) and extending from a central region of the tip (S) to the trailing edge (17); - a central radial pipe (34) collecting air at the root and extending between at least two of the three pipes (41, 42, 43) of the median circuit (28) and directly supplying the cavity (36) under the squealer tip.
Une tête d'extrusion (100) pour fabrication additive comprend une enceinte (110) délimitant une chambre d'alimentation (140). L'enceinte comporte au moins un port d'entrée (120) débouchant dans la chambre d'alimentation (140) et configuré pour recevoir un matériau à extruder sous pression et une pluralité de buses d'extrusion (160). Chaque buse est en communication avec la chambre d'alimentation (140) et débouche à l'extérieur de l'enceinte via un orifice de sortie. Les buses de la pluralité de buses sont disposées adjacentes les unes par rapport aux autres avec un espacement déterminé entre chaque orifice de sortie de buse.
The invention relates to a system for monitoring the health of a helicopter, comprising a helicopter and a device for determining a change in state of the engine that is configured to collect data measured by engine sensors and external conditions during a stable flight phase and to process said measured data in the following way: comparing said measured data with a reference model of the engine, determining, at each time interval, an instantaneous discrepancy between each item of data measured and each item of data estimated by the reference model of the engine, determining, throughout the stable flight phase, an overall discrepancy between the measured data and the data estimated by the reference model of the engine, determining, at each time interval of the stable flight phase, an intrinsic residual corresponding to a difference between the instantaneous discrepancy and the overall discrepancy, determining a deviating portion corresponding to the part of the intrinsic residual that does not satisfy a predetermined criterion, each deviating portion containing an item of information relating to a pilot action that is not recorded, and determining a corrected residual corresponding to the instantaneous discrepancy from which the deviating portion has been subtracted, the corrected residual being analysed in order to determine whether the state of the engine has changed.
The invention relates to a turbine engine vane (20) comprising a blade (21) extending along a radial axis and a first cooling circuit (28) arranged inside the blade, the first cooling circuit (28) comprising a first cavity (34) and a second cavity (35) disposed downstream of the first cavity in a direction of circulation of a coolant in the blade, the first and second cavities radially extending inside the blade and being at least partly separated by a first radial partition (36) having a radially internal free end (37), which at least partly demarcates a first coolant passage (40) connecting the first and second cavities. According to the invention, the radially internal free end (37) is enlarged by having a general transverse section substantially in the form of a keyhole.
The invention relates to a turbine engine blade (20) comprising: - an airfoil (21) with a pressure-side wall and a suction-side wall which are connected upstream by a leading edge (26) and downstream by a trailing edge (27), - a cooling circuit (28) which comprises an internal cavity extending inside the airfoil and a plurality of outlet openings each oriented substantially along a longitudinal axis X, each outlet opening communicating with the cavity and being arranged in the vicinity of the trailing edge, and - a calibration device (33) arranged in the cavity and provided with calibration conduits (34) which are arranged substantially opposite the outlet openings. According to the invention, the calibration conduits (34) each comprise an oblong transverse section which is substantially perpendicular to the longitudinal axis.
The invention relates to a propulsion system (100) for an aircraft, comprising at least one rotor (110) and a nacelle cowling (120) extending around the at least one rotor (110), said nacelle cowling (120) being sectored and comprising at least one sector (130a, 130b) which is fixed and sectors (141a, 141b, 142a, 142b) which are retractable in the peripheral direction (F1, F2) relative to an axis of rotation (X) of the rotor (110), characterised in that the retractable sectors (141a, 141b, 142a, 142b) comprise at least a first series of sectors (141a, 142a) which are telescopically retractable in or on at least one fixed sector (130a) and at least a second series of sectors (141b, 142b) which are telescopically retractable in or on at least one fixed sector (130b), the at least one fixed sector (130a, 130b) having an angular extent around the axis (X) which is less than or equal to 90°.
COMPAGNIE GENERALE DES ETABLISSEMENTS MICHELIN (France)
SAFRAN LANDING SYSTEMS (France)
SAFRAN ELECTRONICS & DEFENSE (France)
Inventor
Riou, Jean-Christophe
Bailly, Eric
Abstract
Pressure measurement device (1) comprising a case (20) extending around an electronic card (30) provided with a pressure sensor (40), - the case (20) delimiting with a first face (31) of the electronic card (30) a first leaktight volume (3); - the case (20) also delimiting with a second face (32) of the electronic card (30) opposite to the first face (31) a second leaktight volume (4); - the case (20) comprising at least one first channel (24) placing in fluidic communication the exterior medium (5) outside the case (20) and the first leaktight volume (3); - the electronic card (30) comprising at least one second channel (33) placing in fluidic communication the first volume (3) and the second volume (4); - the link between the case (20) and the electronic card (30) being designed to permit a relative motion of the case (20) and of the electronic card.
The invention relates to a nickel-based superalloy containing, in weight percentages: - 5 to 6.5% of aluminum, - 4.5 to 7% of cobalt, - 14.5 to 16.5% of chrome, - 0 to 0.2% of hafnium, - 0 to 1.5% of molybdenum, - 2 to 3.5% of tantalum, - 0 to 2% of titanium, - 1 to 2.5% of tungsten, - 0 to 0.08% of zirconium, - 0 to 0.03% of boron and - 0 to 0.07% of carbon, the remainder being nickel and inevitable impurities.
The invention relates to a method for producing a metal bladed element for a turbomachine of an aircraft, said bladed element comprising at least one blade having a lower surface and an upper surface extending between a leading edge and a trailing edge of the blade, the trailing edge having to have a thickness X1, said method comprising the steps of: a) producing the bladed element by lost-wax casting, and b) finishing the bladed element, characterised in that step b) comprises the chemical milling at least of the trailing edge of the or each blade so as to obtain said thickness X1 which cannot be directly obtained by step a).
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
The present invention relates to a core for metal casting an aeronautical part such as a turbine blade, the core being designed for arrangement in an inner cavity defined by a mould, the core comprising: a body (13) intended to form the inner shape of the turbine blade; and an impact portion (15), disposed along at least a portion of the periphery of the body (13) in such a way as to break up a fluid jet when the inner receptacle is filled with fluid, the impact portion (15) having a apex (17) and at least one deflecting wall (19) converging towards the apex (17).
OFFICE NATIONAL D'ETUDES ET DE RECHERCHES AEROSPATIALES (France)
SAFRAN (France)
Inventor
Rame, Jeremy
Jaquet, Virginie
Delautre, Joel
Guedou, Jean-Yves
Caron, Pierre
Lavigne, Odile
Locq, Didier
Perrut, Mikael
Abstract
The invention relates to a nickel-based superalloy comprising, in weight percentages, 4.0 to 5.5 % rhenium, 3.5 to 12.5 % cobalt, 0.30 to 1.50 % molybdenum, 3.5 to 5.5 % chromium, 3.5 to 5.5 % tungsten, 4.5 to 6.0 % aluminum, 0.35 to 1.50 % titanium, 8.0 to 10.5 % de tantalum, 0.15 to 0.30 % hafnium, 0.05 to 0.15 % silicon, the remainder being nickel and inevitable impurities. The invention also relates to a single-crystal blade (20A, 20B) comprising such an alloy and to a turbomachine (10) comprising such a blade (20A, 20B).
OFFICE NATIONAL D'ETUDES ET DE RECHERCHES AEROSPATIALES (France)
SAFRAN (France)
Inventor
Rame, Jeremy
Jaquet, Virginie
Delautre, Joel
Guedou, Jean-Yves
Caron, Pierre
Lavigne, Odile
Locq, Didier
Perrut, Mikael
Abstract
The invention relates to a nickel-based superalloy comprising, in weight percentages, 4.0 to 5.5 % rhenium, 1.0 to 3.0 ruthenium, 2.0 to 14.0 % cobalt, 0.3 to 1.0 % molybdenum, 3.0 to 5.0 % chromium, 2.5 to 4.0 % tungsten, 4.5 to 6.5 % aluminium, 0.50 to 1.50 % titanium, 8.0 to 9.0 % de tantalum, 0.15 to 0.30 % hafnium, 0.05 to 0.15 % silicon, the rest being nickel and inevitable impurities. The invention also relates to a single-crystal blade (20A, 20B) comprising such an alloy and to a turbomachine (10) comprising such a blade (20A, 20B).
COMPAGNIE GENERALE DES ETABLISSEMENTS MICHELIN (France)
SAFRAN LANDING SYSTEMS (France)
SAFRAN ELECTRONICS & DEFENSE (France)
Inventor
Fagot-Revurat, Lionel
Destraves, Julien
Abstract
The invention relates to a parameter measurement system for a mounted assembly comprising an electronic device for measuring the parameters of the mounted assembly and a connecting interface made of elastomeric material including the electronic device. The electronic measurement device comprises: - a UHF radiofrequency antenna; - an electronic card with an electronic chip coupled to the UHF radiofrequency antenna, a sensor for measuring the parameter of the mounted assembly, a microcontroller and an electrical circuit. The measurement system is characterised in that it comprises a ground plane connected to the electronic card, in that the electronic card comprises an energy manager and a capacitive element, in that the coupling between the electronic chip and the UHF radiofrequency antenna is an electrical coupling and in that the electronic chip, the microcontroller and the measurement sensor are components with low energy consumption.
The invention relates to a part made from a composite material having a ceramic matrix and comprising a protection device, the protection comprising a coating layer having a gradual composition. The aforementioned gradual-composition coating layer comprises at least one silicon phase and one aluminium phase, the proportions of which change according to the height in the layer, with a first height in the gradual-composition coating layer corresponding to a silicon-free composition, and a second height corresponding to an aluminium-free composition. The invention also relates to a method for producing such a protection, and to a device for carrying out the method.
C23C 28/04 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of inorganic non-metallic material
29.
METHOD FOR DETERMINING A RECOMMENDED INFLATION PRESSURE FOR AN AIRCRAFT TYRE, AND ASSOCIATED MAINTENANCE METHODS
COMPAGNIE GENERALE DES ETABLISSEMENTS MICHELIN (France)
SAFRAN (France)
SAFRAN ELECTRONICS & DEFENSE (France)
Inventor
Arnoux, Michael
Neba, Eric Carin
Abstract
The invention relates to a method for determining an expected inflation pressure in an aircraft tyre that is equipped with a pressure and temperature sensor. The invention also relates to an assisted maintenance method for an aircraft tyre, as well as to a portable electronic device designed to assist with maintenance.
The invention relates to a turbine component comprising a substrate made from monocrystalline nickel-based superalloy comprising rhenium, which has a ?-?' Ni phase, and an average weight fraction of chromium of less than 0.08, a sublayer made from nickel-based metal superalloy covering the substrate, characterized in that the sublayer made from metal superalloy comprises at least aluminium, nickel, chromium, silicon, hafnium and has, predominantly by volume, a ?'-Ni 3 Al phase. Figure 4
The invention concerns a component (10) comprising: a substrate (20) of which at least one part adjacent to a surface (S) of the substrate is made from a material comprising silicon; a bonding sublayer (30) situated on the surface (S) of the substrate and comprising silicon, an environmental barrier (40) that comprises an outer layer (42) made from ceramic covering the bonding sublayer (30), characterised in that said environmental barrier (40) additionally comprises a self-healing inner layer (41) situated between the bonding sublayer (30) and the outer layer (42), said inner layer (41) comprising a matrix in which silica-forming particles are dispersed, these particles being capable of generating a crack-healing phase in the matrix in the presence of oxygen.
C23C 28/04 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of inorganic non-metallic material
32.
DEVICE FOR HOLDING ONE OR MORE ELECTRODES FOR ELECTRICAL DISCHARGE MACHINING, AND METHOD OF OBTAINING SAME
The invention relates to creating an electrode holder for electrical discharge machining, comprising a body presenting a rectilinear part in which is fitted at least a first electrode passage duct. The body also presents a curved monobloc part in which is fitted (at least) a second curved dielectric fluid supply duct where a curved extension of said at least first duct goes through. The curved extension and the second curved duct are in ceramic with an arithmetic average roughness of: Ra < 2 .MU.m.
The invention relates to a turbine part, such as a turbine blade or a distributor fin, for example, comprising a substrate made of superalloy based on monocrystalline nickel, comprising rhenium and/or ruthenium, and having a ?'-NisAI phase that is predominant by volume and a ?-Ni phase, the part also comprising a sublayer made of metal superalloy based on nickel covering the substrate, characterized in that the sublayer has a ?'-NisAI phase that is predominant by volume and in that the sublayer has an average atomic fraction of aluminium of between 0.15 and 0.25, of chromium of between 0.03 and 0.08, of platinum of between 0.01 and 0.05, of hafnium of less than 0.01 and of silicon of less than 0.01. A process for manufacturing a turbine part comprising a step of vacuum deposition of a sublayer made of a superalloy based on nickel having predominantly by volume a ?'-NisAI phase, on a substrate made of superalloy based on nickel comprising rhenium and/or ruthenium, the sublayer having an average atomic fraction of aluminium of between 0.15 and 0.25, of chromium of between 0.03 and 0.08, of platinum of between 0.01 and 0.05, of hafnium of less than 0.01 and of silicon of less than 0.01.
C22C 19/05 - Alloys based on nickel or cobalt based on nickel with chromium
C23C 4/08 - Metallic material containing only metal elements
C23C 14/16 - Metallic material, boron or silicon on metallic substrates or on substrates of boron or silicon
C23C 28/00 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
The invention relates to a casting slurry for manufacturing shell moulds, comprising powder particles and a binder, and characterised in that it contains a surfactant. The invention also relates to the use of a casting slurry of this type for manufacturing a shell mould.
B22C 1/02 - Compositions of refractory mould or core materialsGrain structures thereofChemical or physical features in the formation or manufacture of moulds characterised by additives for special purposes, e.g. indicators, breakdown additives
B22C 1/18 - Compositions of refractory mould or core materialsGrain structures thereofChemical or physical features in the formation or manufacture of moulds characterised by the use of binding agentsMixtures of binding agents of inorganic agents
B22C 1/20 - Compositions of refractory mould or core materialsGrain structures thereofChemical or physical features in the formation or manufacture of moulds characterised by the use of binding agentsMixtures of binding agents of organic agents
A coated turbomachine part (20) comprises a substrate (21) and a layer for protection against the calcium and magnesium aluminosilicates CMAS (22) present on the substrate (21). The protective layer (22) comprises a first phase (220) of a material for protection against the calcium and magnesium aluminosilicates CMAS, capable of forming an apatite- or anorthite-type phase in the presence of calcium and magnesium aluminosilicates CMAS, and a second phase (221) comprising particles of at least one rare earth silicate REa dispersed in the first phase.
A coated turbomachine part comprises a substrate (21) and a layer for protection against the calcium and magnesium aluminosilicates CMAS (24) present on the substrate (21). The layer (24) comprises a first phase (240) of a material for protection against the calcium and magnesium aluminosilicates CMAS and a second phase (241) comprising particles of a non-wetting material dispersed in the first phase.
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
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
37.
PROCEDE DE FABRICATION AMELIORE D'UNE PIECE A MICROSTRUCTURE DUALE
A method for welding together at least two parts of green material, referred to as green parts, by means of co-sintering, comprising the following steps: - assembling the at least two green parts at a junction zone of said parts so as to form a green one-piece assembly, - de-binding the green one-piece assembly, and - sintering the one-piece assembly so as to obtain a dense one-piece assembly forming a final part, characterised in that the two green parts (10, 12) each have a composition of different powder, so as to produce a final part (1) having at least two parts with different grain sizes.
B22F 3/22 - Manufacture of workpieces or articles from metallic powder characterised by the manner of compacting or sinteringApparatus specially adapted therefor for producing castings from a slip
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
38.
HYBRID PROPULSION ARCHITECTURE FOR AN AIRCRAFT COMPRISING A MOTOR WITH TWO REVERSIBLE ELECTRIC MACHINES MOUNTED ON TWO SHAFTS
The aircraft motor architecture comprises two reversible electric machines (3, 4), the rotors (10) of which are linked both to the low pressure shaft (1) and to the high pressure shaft (2) by transmissions (11, 12, 13, 14) alternately disengaged depending on the direction of rotation of the rotor (10), the transmissions comprising passive one-way clutches (15, 16, 17, 18), the engagement directions of which are opposed. Independent modes of operation of the machines, as a starter or as an electric generator of each of the shafts, are thus provided.
The invention concerns a method for manufacturing a turbomachine flow straightener made from composite material having a blade extended by at least one attachment end (3, 4), comprising the operations of: forming a fibrous blank (13) from a three-dimensional fabric comprising warp threads (12) oriented in a direction (AR) corresponding to the longitudinal direction of the blade to be manufactured; dividing at least one end of the fibrous blank (13) by separating in order to constitute at least two sets of integrated layers (21, 22, 23); shifting at least one set of integrated layers (22, 23) in a planar configuration such that the warp threads (12) are inclined in this set of integrated layers (22, 23) relative to the orientation of the warp threads in the fibrous blank portion of the blade.
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
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 70/54 - Component parts, details or accessoriesAuxiliary operations
B29D 99/00 - Subject matter not provided for in other groups of this subclass
The invention relates to a turbine blade comprising a blade root and a blade (13) extending radially outwards from the blade root (12), said blade (13) comprising a first inner cooling circuit comprising a lower surface cavity (33, 36) extending radially along the lower surface wall (16) and along a first inner wall (47, 45) arranged between the lower surface wall (16) and the upper surface wall (18), an upper surface cavity (34, 37) extending radially along the upper surface wall (18) and along a second inner wall (47, 43) arranged between the lower surface wall (16) and the upper surface wall (18). The first cooling circuit comprises an inner through-cavity (35, 38) defined between two through-walls (59, 57, 55, 53), each extending between the lower surface wall (16) and the upper surface wall (18). The lower surface cavity (33, 36), the upper surface cavity (34, 37) and the inner through-cavity (35, 38) are connected fluidically in series.
The invention relates to an aviation turbine blade (10), characterised in that it comprises at least a first lower surface cavity (C2) and a first upper surface cavity (C3), each adjacent to a first through-cavity (C1) and a second through-cavity (C4), the first upper surface cavity (C3) being adjacent to the upper surface wall (24), the first lower wing surface cavity (C2) being adjacent to the lower surface wall (22), each of said first and second through-cavities (C1, C4) extending from the lower surface wall (22) as far as the upper surface wall (24), the second through-cavity (C4) comprising a first inner wall (P1) extending from the upper surface wall (24) as far as the first through-cavity (C1), and a second inner wall (P2) extending from the lower surface wall (22) as far as the first through-cavity (C1). The first (P1) and second (P2) inner walls are not connected.
A method for producing a consolidated fibrous preform intended for the manufacture of a part made of composite material, comprising the steps of: -shaping a fibrous texture (10) in a heated metal mould (20), the texture being pre-impregnated with a transient or transitory material (30), or shaping a fibrous texture in a metal mould and injecting a transient or transitory material into the fibrous texture held in shape in the metal mould, - cooling the mould, - removing the set fibrous preform from the mould, - coating the fibrous preform with a slip containing a powder of ceramic or carbon particles, - heat-treating the coated fibrous preform so as to form a porous shell around the fibrous preform by consolidation of the slip and so as to remove the transient or transitory material present in the fibrous preform, - consolidating the fibrous preform by gas-phase chemical infiltration.
A process for manufacturing a power electronic module (20) by additive manufacturing, the electronic module (20) comprising a substrate (21) including an electrically insulating sheet (24) having first and second opposite faces (24a, 24b), and a first metal layer (25a) that is placed directly on the first face (24a) of the insulating sheet (24) and a second metal layer (25b) that is placed directly on the second face (25b) of the insulating sheet (24). At least one of the metal layers (25a) is produced via a step (100) of depositing a thin copper layer and a step (110) of annealing the metal layer (25a, 25b), and the process furthermore comprises a step (120) of forming at least one thermomechanical transition layer (271 to 273, 274 to 276) on at least one of the first and second metal layers (25a, 25b), said at least one thermomechanical transition layer (271 to 273, 274 to 276) including a material having a thermal expansion coefficient lower than that of the metal of the metal layer (25a, 25b).
H01L 21/48 - Manufacture or treatment of parts, e.g. containers, prior to assembly of the devices, using processes not provided for in a single one of the groups or
H01L 23/373 - Cooling facilitated by selection of materials for the device
"The main object of the invention relates to a cluster-shaped model and a carapace (1) for the lost-wax molding of a plurality of turbomachine elements, the carapace (1) comprising a central descendant (3) in fluid communication with a casting cup (2) for receiving the molten metal, and a plurality of carapace elements (4). The carapace (1) is characterized in that it further comprises a plurality of source supply ducts (5) of the carapace elements (4) and a handling accessory carapace (6), independent of the plurality of carapace elements (4) and their metal supply circuit so as to be without fluid communication with the carapace elements (4), the handling accessory carapace (6) being in fluid communication with the central descendant (3) to allow a fall casting of the handling accessory carapace (6)."
A device (100) for non-destructive characterisation of a material (2), the device (100) comprising transmission/reception cells (10), each cell (10) being configured, in a transmission mode, for transmitting ultrasonic waves toward the material (2) to be characterised and, in a receiving mode, for receiving ultrasonic waves transmitted through said material (2), the non-destructive characterisation device (100) comprising a ring consisting of a plurality of adjacent angular sectors, each angular sector comprising a stack of transmission/reception cells (10) in a radial direction (DR) of the ring.
B06B 1/06 - Processes or apparatus for generating mechanical vibrations of infrasonic, sonic or ultrasonic frequency making use of electrical energy operating with piezoelectric effect or with electrostriction
G01N 29/22 - 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 Details
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
SAFRAN (France)
SAFRAN AIRCRAFT ENGINES (France)
"ASSOCIATION POUR LA RECHERCHE ET LE DEVELOPPEMENT DES METHODES ET PROCESSUS INDUSTRIELS" - A.R.M.I.N.E.S. (France)
Inventor
Remacha, Clement
Romero, Edward
Arnaud, Alexiane
Proudhon, Henry
Herbland, Thibault
Abstract
The invention relates to a method for testing the crystallographic orientation of at least one grain of a turbomachine part, comprising the steps of: a) emitting a beam of electromagnetic radiation through an elementary volume of the part and recording diffraction information on the electromagnetic radiation through the part; b) repeating step a) on a given area of the part, c) determining the crystalline spatial orientation of each of said elementary volumes and deducing therefrom the presence of at least one first crystallographic grain for which the elementary volumes are oriented according to the same crystallographic orientation; d) calculating the angular distance between the crystalline spatial orientation of said first grain and a predetermined direction taken from the part and comparing it to a first predetermined threshold value; e) determining a state of use of the part.
G01N 23/20 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by using diffraction of the radiation by the materials, e.g. for investigating crystal structureInvestigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by using scattering of the radiation by the materials, e.g. for investigating non-crystalline materialsInvestigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by using reflection of the radiation by the materials
47.
TURBOMACHINE PART COATED WITH A THERMAL BARRIER AND PROCESS FOR OBTAINING SAME
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
SAFRAN (France)
SAFRAN AIRCRAFT ENGINES (France)
Inventor
Joulia, Aurelien
Malie, Andre Hubert Louis
Ansart, Florence
Delon, Elodie Marie
Duluard, Sandrine
Abstract
The invention relates to a turbomachine part (30) coated with at least one first thermal barrier layer (33) comprising a ceramic material and first ceramic fibers (34) dispersed in said first layer. The first layer may comprise a gradient of chemical composition between a thermal barrier material and a material for protection against calcium and magnesium aluminosilicates the content of which is greatest on an outer zone of the first layer, and/or the first layer may be porous and have a porosity gradient such that an outer portion of the first layer has a reduced porosity. The invention also targets a process for manufacturing such a part comprising the formation of the first layer on the part via a wet process.
Exchanger (1) of heat between a first fluid flowing in a longitudinal direction (X) and a second fluid, the said exchanger (1) comprising: two parallel plates (6) distant from one another; at least a first and second row (8a, 8b) of fins (9) arranged perpendicularly between the said plates (6), each fin (9) being delimited longitudinally by a first edge (10) and a second edge (11), the said first edge (10) comprising at each of its ends a region of connection with the corresponding plate (6); characterized in that the said connection regions of the said first edge (10) are respectively inclined with respect to a normal to the plates (6) in a plane (P) perpendicular to the said plates (6) and parallel to the direction (X), the said first edge (10) and the said second edge (11) of each of the fins (9) having identical profiles in the said plane (P).
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
B33Y 30/00 - Apparatus for additive manufacturingDetails thereof or accessories therefor
F02C 7/08 - Heating air supply before combustion, e.g. by exhaust gases
F28D 9/00 - Heat-exchange apparatus having stationary plate-like or laminated conduit assemblies for both heat-exchange media, the media being in contact with different sides of a conduit wall
F28D 21/00 - Heat-exchange apparatus not covered by any of the groups
F28F 3/04 - Elements or assemblies thereof with means for increasing heat-transfer area, e.g. with fins, with recesses, with corrugations the means being integral with the element
F28F 3/06 - Elements or assemblies thereof with means for increasing heat-transfer area, e.g. with fins, with recesses, with corrugations the means being attachable to the element
F28F 21/08 - Constructions of heat-exchange apparatus characterised by the selection of particular materials of metal
49.
SUPERALLOY BASED ON NICKEL, MONOCRYSTALLINE BLADE AND TURBOMACHINE
OFFICE NATIONAL D'ETUDES ET DE RECHERCHES AEROSPATIALES (France)
SAFRAN (France)
SAFRAN AIRCRAFT ENGINES (France)
SAFRAN HELICOPTER ENGINES (France)
Inventor
Rame, Jeremy
Belaygue, Philippe
Caron, Pierre
Delautre, Joel
Jaquet, Virginie
Lavigne, Odile
Abstract
The invention relates to a superalloy based on nickel, comprising, in mass percentages, 4.0 to 6.0 % of chromium, 0.4 to 0.8 % of molybdenum, 2.5 to 3.5 % of rhenium, 6.2 to 6.6 % of tungsten, 5.2 to 5.7 % of aluminium, 0.0 to 1.6 % of titanium, 6.0 to 9.9 % of tantalum, 0.0 to 0.7 % of hafnium, and 0.0 to 0.3 % of silicon, the rest consisting of nickel and potential impurities. The invention also relates to a monocrystalline blade (20A, 20B) comprising such an alloy and to a turbomachine (10) comprising such a blade (20A, 20B).
The invention relates to a method for monitoring the soundness of helicopters comprising the determination of the severity of a plurality of flight missions of a plurality of helicopters, comprising a step for acquiring and storing flight data from helicopter flight missions, and a step for acquiring and storing maintenance data from the plurality of helicopters. The method is characterised in that said determination comprises a mission-type construction step, comprising a sub-step for constructing descriptors, a sub-step for partitioning the descriptors and a sub-step for allocating a mission type to each flight by associating the descriptor of said flight and a sub-set, in which this descriptor is found, and a step for interpreting the severity of the mission types, comprising a sub-step for estimating the severity models, and a sub-step for associating a severity model with each mission type determined in the mission type construction step.
B64D 45/00 - Aircraft indicators or protectors not otherwise provided for
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
G07C 5/00 - Registering or indicating the working of vehicles
51.
METHOD FOR NONDESTRUCTIVE INSPECTION BY ULTRASOUND OF A BONDED ASSEMBLY
The invention relates to a method for nondestructive inspection by ultrasound of a bonded assembly (1). The method comprises two steps, consisting of measuring a thickness of an adhesive joint (7) of the bonded assembly (1) by means of an ultrasound transducer (9) arranged on the bonded assembly (1) in a determined position, and measuring the degree of adhesion of parts of the bonded assembly (1) by means of the same ultrasound transducer (9) maintained in said determined position, the degree of adhesion being measured by means of ZGV Lamb waves (13).
COMMISSARIAT A L'ENERGIE ATOMIQUE ET AUX ENERGIES ALTERNATIVES (France)
SAFRAN (France)
Inventor
Bernard, Benjamin
Quet, Aurelie
Herve, Emmanuel
Bianchi, Luc
Joulia, Aurelien
Malie, Andre
Abstract
The invention relates to a method for coating at least one surface of a solid substrate with at least one layer comprising at least one ceramic compound by a suspension plasma spraying (SPS) technique, in which at least one suspension of solid particles of at least one ceramic compound is injected into a plasma jet, and then the thermal jet that contains the solid particle suspension is sprayed onto the surface of the substrate, by means of which the layer comprising at least one ceramic compound is formed on the surface of the substrate; method characterised in that, in the suspension, at least 90 vol% of the solid particles have a larger dimension (referred to as d90), such as a diameter, smaller than 15 µm, preferably smaller than 10 µm, and at least 50 vol% of the solid particles have a larger dimension, such as a diameter (referred to as d50), no smaller than 1 µm. The invention also relates to a substrate coated with at least one layer that can be obtained by said method. The invention also relates to a part comprising said coated substrate. The invention further relates to the use of said layer in order to protect a solid substrate against degradations caused by contaminants such as CMAS.
C23C 4/12 - Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the method of spraying
C23C 28/00 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and
C23C 28/04 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and only coatings of inorganic non-metallic material
Method for producing a thermal barrier system on a metal substrate (1) of a turbo engine part, such as a high-pressure turbine blade, the thermal barrier system comprising at least one columnar ceramic layer (31,..., 3i,..., 3n), characterised in that the method comprises a step of compressing at least one of said at least one columnar ceramic layer (31,...3i,..., 3n).
C23C 28/00 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and
The invention relates to a method of assembling a first core and a second core for the creation of a non-permanent model configured to form, by moulding using the lost-wax method, a component comprising a first cavity and a second cavity corresponding respectively to the first core and the second core. The invention is characterized in that the first and second cores are assembled using a first spacer piece, the first spacer piece being positioned between the first and second cores.
The invention relates to a process for manufacturing a shell mold having several layers, including at least one contact layer from a model of a part to be manufactured made of wax or other similar material, the process comprising a step of dipping the model in a contact slip forming the contact layer and comprising an inorganic or organic binder and a powder, wherein the powder is a mullite-zirconia composite.
B22C 1/08 - Compositions of refractory mould or core materialsGrain structures thereofChemical or physical features in the formation or manufacture of moulds characterised by additives for special purposes, e.g. indicators, breakdown additives for decreasing shrinkage of the mould, e.g. for investment casting
B22C 1/16 - Compositions of refractory mould or core materialsGrain structures thereofChemical or physical features in the formation or manufacture of moulds characterised by the use of binding agentsMixtures of binding agents
Directional solidification cooling furnace (20) for a metal casting, comprising a cylindrical internal chamber (26) with a vertical central axis (X) and a mold support (28) positioned in the internal chamber (26), the internal chamber (26) comprising a casting zone (A) and a cooling zone (B), the casting zone (A) and the cooling zone (B) being superimposed on one another, the casting and cooling zones being thermally insulated from one another, when the mold support is positioned in the casting zone (A), by a fixed first heat shield (31) and a second heat shield (32) borne by the mold support (28), the casting zone (A) comprising at least a first heating device and the cooling zone (B) comprising a second heating device (60), the first and second heating devices being configured so that the temperature of the casting zone (A) is higher than the temperature of the cooling zone (B), the cooling zone (B) comprising an upper portion (B') and a lower portion (B") superposed on one another and thermally insulated from one another by a third heat shield (33), the upper portion (?') of the cooling zone (B) comprising the second heating device (60).
The invention relates to a method for protecting a hafnium-free, nickel-based monocrystalline superalloy part (1) against corrosion and oxidation. Said method is characterized in that it involves at least the steps of: - manufacturing a hafnium-free, nickel-based monocrystalline superalloy part (1); - depositing a first hafnium-containing layer (2), an undercoat made of an alloy containing at least 10 atom % aluminum, and a second hafnium-containing layer on the part (1) such that a mixed layer (3) is formed, and depositing a third hafnium-containing layer (4) on the part (1); - performing a diffusion treatment in order for the first (2) and third (4) layers to diffuse such that a first interdiffusion region (21) is formed on the top portion of the part (1) and a second interdiffusion region (41) is formed on the surface of the mixed layer (3); - subjecting the second interdiffusion region (41) to an oxidation treatment so as to obtain a hafnium-doped alumina layer (42).
C23C 10/28 - Solid state diffusion of only metal elements or silicon into metallic material surfaces using solids, e.g. powders, pastes
C23C 12/00 - Solid state diffusion of at least one non-metal element other than silicon and at least one metal element or silicon into metallic material surfaces
C23C 28/00 - Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of main groups , or by combinations of methods provided for in subclasses and
58.
PROCESS FOR MANUFACTURING A PART MADE OF NICKEL-BASED SUPERALLOY CONTAINING HAFNIUM
The invention relates to a process for manufacturing a part made of nickel-based monocrystalline superalloy containing hafnium. This process is noteworthy in that it comprises the following successive steps consisting in: - manufacturing a nickel-based monocrystalline superalloy that is not doped with hafnium, - manufacturing a part from this superalloy, - directly depositing on said part a layer of hafnium having a thickness of between 50 nm and 800 nm, - carrying out a diffusion treatment of the hafnium so as to form an interdiffusion layer at the surface of said part and to thus obtain a part made of nickel-based monocrystalline superalloy containing hafnium.
The invention relates to a mould (3) made of a ceramic material intended to be used to mould a turbomachine blade from a molten metal, the blade comprising a base, a platform, a vane, and a root; the mould comprising: a cavity (4) having the shape of the blade, and an auxiliary grain duct (5) comprising a first portion (51) and a second portion (52) extending the first portion, said first portion leading, at one end (511), into a first part (40) of the cavity forming the base of the blade and, at another end (512), into a second part (412) of the cavity forming a spoiler of the blade platform, said second portion (52) leading, at one end (521), into said second part (412) of the cavity and, at another end, (522) into a third part (432) of the cavity forming a spoiler of the blade root. The invention also aims for an installation and a method of manufacture implementing such a mould.
The invention concerns a supply system (5) for conveying a molten metal intended to mould a part, the system comprising a supply channel (51) made from ceramic material configured to allow the molten metal to flow by gravity inside said supply channel, said supply channel comprising a first portion (51a) extending in a first direction (A), and at least one second portion (51b) extending in a second direction (B) different to the first direction, said second portion being arranged downstream from the first portion and being linked to the first portion by a joint (52). The system further comprises a damping channel (54) comprising a first end (54a) opening in the joint and a second closed end (54b), said damping channel extending the first portion (51a) of the supply channel. The invention also concerns a facility comprising such a system, and a manufacturing method implementing same.
A method for producing a part made of composite material comprises placing a fibrous texture (10) in a mould (110) comprising in its lower portion a part made of porous material (130) on which rests a first face (10b) of the texture (10), injection under pressure of a liquid (150) containing a powder of refractory ceramic particles in the fibrous texture (10), and drainage through the part made of porous material (120) of the liquid having passed through the fibrous texture (10), and retention of the powder of refractory ceramic particles inside the texture by the part made of porous material (130). An added rigid element (140) is interposed between the bottom (111) of the mould (110) and the part made of porous material (130).
B01D 29/05 - Filters with filtering elements stationary during filtration, e.g. pressure or suction filters, not covered by groups Filtering elements therefor with flat filtering elements supported
B28B 7/34 - Moulds, cores, or mandrels of special material, e.g. destructible materials
B28B 7/46 - MouldsCoresMandrels characterised by means for modifying the properties of the moulding material for humidifying or dehumidifying
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 70/44 - 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
C04B 35/14 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on oxides based on silica
C04B 35/18 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on oxides based on silicates other than clay rich in aluminium oxide
C04B 35/447 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on oxides based on phosphates
C04B 35/56 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxides based on carbides
C04B 35/565 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxides based on carbides based on silicon carbide
C04B 35/58 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxides based on borides, nitrides or silicides
C04B 35/584 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on non-oxides based on borides, nitrides or silicides based on silicon nitride
C04B 35/626 - Preparing or treating the powders individually or as batches
C04B 35/80 - Fibres, filaments, whiskers, platelets, or the like
D03D 25/00 - Woven fabrics not otherwise provided for
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
The invention relates to a turbomachine turbine vane comprising a root having a blade (12) comprising a leading edge (16) and a trailing edge (17) and a lower surface wall (14) and an upper surface wall (13), as well as: a manifold for cooling (18) the leading edge (16); a supply line (19) for collecting air from the root and supplying the manifold (18); a side cavity (23) extending along the upper surface wall (13) and supplied with air from the root, to form a heat shield facing the supply line (19); and a circuit forming a trombone comprising a middle portion (24) extending along the supply line (19) and extending laterally to the upper surface wall (13), said middle portion (24) being supplied by the supply line at the tip of the blade (12).
The invention relates to a core (16) for casting a blade of a turbomachine, said core (16) comprising a first part (17) for defining a first cavity and a second part (18) of which at least one portion (19) defines a second cavity located between the first cavity and the tip of the blade, wherein: - said portion (19) of the second part (18) includes a through-hole that ends opposite the first part (17) so as to define, in the cast blade, an outer face of a duct for removing dust from the first cavity; - an alumina pin (27) which is secured to the end face (28) of the first part (17) extends into the through-hole to define the inner face of the duct; - centering means (29) are placed between the pin (27) and the through-hole so as to center the pin in relation to the through-hole, said centering means (29) being made of a material that is intended to dissolve before the blade is cast around the core (16).
Method for manufacturing at least one portion of a part using successive deposition of layers, said method involving the steps of: a) depositing a first layer (110) of a molten metal on a substrate (80) such that a first metal strip is formed on the substrate; b) depositing a second layer of a molten metal on the first strip such that a second metal strip is formed on the first strip; and c) repeat steps a) and b) for each new metal layer to be deposited on a preceding strip until the at least one portion of the part has been formed; characterized in that after n deposition steps, where n is greater than or equal to 1, the method includes a compression step in which the obtained strip is compressed, and in that the compression takes place in the hot state, i.e. before the strip has entirely cooled down.
The invention relates to a process for treating a metal alloy part, characterized in that it comprises the following steps: - producing a stock formulation by mixing, in equal molar parts of silicon, an alcoholic solution of hydrolysed epoxysilane and an alcoholic solution of hydrolysed aminosilane, - mixing the stock formulation with a suspension comprising conductive nanowires in an amount by weight of between 0.1% and 10% relative to the total weight of the stock formulation in order to obtain a dilute formulation, and - depositing the dilute formulation on the part in order to obtain the coating.
C23C 18/12 - Chemical coating by decomposition of either liquid compounds or solutions of the coating forming compounds, without leaving reaction products of surface material in the coatingContact plating by thermal decomposition characterised by the deposition of inorganic material other than metallic material
66.
METHOD FOR FORMING DUST-REMOVAL HOLES FOR A TURBINE BLADE AND ASSOCIATED CERAMIC CORE
The invention relates to a ceramic core (10) used for the production of a hollow turbine blade for a turbomachine, using the lost-wax foundry technique, said blade comprising calibrated dust-removal holes extending from an apex of at least one cavity (22) and opening into a channel of the blade, each of said calibrated dust-removal holes being formed in a part of the core with a sufficient pre-determined height to ensure mechanical strength, said part of the core comprising a through-opening having an axis perpendicular to a longitudinal axis of the calibrated dust-removal hole and defining, on each side of the through-opening, a core cylinder having a pre-determined diameter corresponding to the dust-removal hole to be formed and a remaining core volume that is intended to be refilled after casting, such that the calibrated dust-removal hole is obtained without being pierced and without the use of connecting rods.
A refractory core (12) for manufacturing a hollow turbomachine blade (10) according to the lost-wax process comprises a main body (14) and at least one shell (16) connected to the main body (14) and defining a cavity (18) between the main body and the shell, the shell (16) being configured such that it contacts the blade (10) during manufacture.
The invention relates to an aircraft engine part (1) comprising at least one metal substrate (2) and an erosion protective coating (3) present on the substrate, the coating comprising at least one phase (4) comprising at least chromium in an atomic content greater than or equal to 45% and carbon in an atomic content between 5% and 20%, said phase comprising chromium carbides Cr7C3 and Cr23C6. The invention also relates to a process for manufacturing such a part in which a coating composition is deposited on the part by electrodeposition and the part is heat treated at a temperature between 250°C and 700°C.
B32B 15/04 - Layered products essentially comprising metal comprising metal as the main or only constituent of a layer, next to another layer of a specific substance
B32B 15/16 - Layered products essentially comprising metal next to a particulate layer
B32B 33/00 - Layered products characterised by particular properties or particular surface features, e.g. particular surface coatingsLayered products designed for particular purposes not covered by another single class
C25D 3/06 - ElectroplatingBaths therefor from solutions of chromium from solutions of trivalent chromium
C25D 5/50 - After-treatment of electroplated surfaces by heat-treatment
69.
METHOD FOR MANUFACTURING A COMPONENT MADE OF COMPOSITE MATERIAL COMPRISING A BODY INTEGRAL WITH ONE OR MORE PLATFORMS
A method for manufacturing a component made of composite material having a body integral with at least one platform present at one end of the said body, the method comprising: - forming, from a fibrous blank, a preform of the component that is to be produced, by unfolding, on either side of a first part, segments (104a; 114a) of a second part and of a third part (106a; 116a) not connected to the first part, by shaping the segment (104a; 114a) unfolded from the second part and the segment (106a; 116a) unfolded from the third part to form preform parts of a platform of the component that is to be manufactured, and by folding the segment (102a; 112a) of the first part over on itself to fill a slot (108; 118) present at the separation between the second and third parts, - densifying the preform with a matrix in order to obtain a component made of composite material having at least one inbuilt platform.
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 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29D 99/00 - Subject matter not provided for in other groups of this subclass
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
70.
DEVICE FOR GENERATING A STRUCTURAL-GRADIENT MICROSTRUCTURE ON AN AXISYMMETRIC PART
The invention relates to a device (1) for generating a structural-gradient microstructure on an axisymmetric mechanical part (P) hollowed out in the center thereof and initially having a fine-grained uniform structure. The device (1) includes a first heating means (2) that defines a first chamber for receiving the mechanical part (P) and is capable of heating the outer periphery (E) of said mechanical part (P) to a first temperature (T1) of greater than the solvus temperature. The device (1) includes a second heating means (3) that defines a second chamber arranged inside the first chamber and is capable of heating the inner periphery (I) of said mechanical part (P) to a second temperature (T2) of less than the solvus temperature. The space between the first chamber and the second chamber defines a recess (L) capable of accommodating the axisymmetric metal part (P) hollowed out in the center thereof.
C21D 1/00 - General methods or devices for heat treatment, e.g. annealing, hardening, quenching or tempering
C22F 1/10 - Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of nickel or cobalt or alloys based thereon
71.
TURBINE ENGINE PART COVERED WITH A PROTECTIVE CERAMIC COATING, METHOD FOR MANUFACTURING AND FOR USING SUCH A PART
The invention relates to a turbine engine part (10) comprising at least one substrate (11) and one ceramic coating (12) for protection against calcium and magnesium aluminum silicates present on the substrate, the ceramic coating including Al2O3 with a molar content of between 33% and 49%, Y3Al5O12 with a molar content of between 21% and 53%, and yttria-stabilized zirconia with a molar content of between 13% and 31%. The invention also relates to a method for manufacturing such a part and to a method for using such a part.
C23C 30/00 - Coating with metallic material characterised only by the composition of the metallic material, i.e. not characterised by the coating process
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
72.
PROCESS FOR MANUFACTURING AN IMPREGNATED FIBROUS ASSEMBLY
The present invention relates in particular to a process for manufacturing an impregnated fibrous assembly, the process comprising at least the following steps: - introducing a first suspension (10) comprising a first powder of solid particles into an internal volume (2) delimited by an inner face (1a) of a first fibrous texture (1) of hollow shape placed in a mould, an outer face (1b) of the first fibrous texture (1) being present facing a wall (3) of the mould, - impregnating, by action of the centrifugal force, the first fibrous texture (1) with the first suspension (10) by rotating the mould about itself, - positioning, after impregnation of the first texture (1) with the first suspension, a second fibrous texture on the inner face (1a) of the first fibrous texture (1) in order to obtain a fibrous assembly, - introducing a second suspension comprising a second powder of solid particles into the internal volume (2) after positioning the second fibrous texture, and - impregnating, by action of the centrifugal force, the second fibrous texture with the second suspension by rotating the mould about itself in order to obtain the impregnated fibrous assembly.
B22F 7/04 - Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting of composite layers with one or more layers not made from powder, e.g. made from solid metal
B28B 1/20 - Producing shaped articles from the material by centrifugal or rotational casting
B28B 1/28 - Producing shaped articles from the material by slip-casting, i.e. by casting a suspension or dispersion of the material in a liquid-absorbent or porous mould, the liquid being allowed to soak into or pass through the walls of the mouldMoulds therefor involving rotation of the mould
B28B 1/52 - Producing shaped articles from the material specially adapted for producing articles from mixtures containing fibres
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 70/32 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or coreShaping by spray-up, i.e. spraying of fibres on a mould, former or core on a rotating mould, former or core
The invention relates in particular to a method for impregnating a hollow fibrous texture, the method comprising at least the following steps: injecting a first suspension (10) comprising a first powder of solid particles made of ceramic or carbon material into an inner space (2) defined by an inner surface (1a) of a hollow fibrous texture (1) placed in a mould, an outer surface (1b) of the fibrous texture (1) being presented opposite a wall (3) of the mould; and impregnating the fibrous texture (1) with the first suspension (10) by centrifugal force, by rotating the mould on itself, varying the speed of rotation of the mould during the impregnation of the texture (1) with the first suspension (10).
B28B 23/20 - Arrangements specially adapted for the production of shaped articles with elements wholly or partly embedded in the moulding material wherein the elements are reinforcing members the shaping being effected by centrifugal or rotational moulding
B29C 41/04 - Rotational or centrifugal casting, i.e. coating the inside of a mould by rotating the mould
The invention relates to an intermediate casing hub of an aircraft turbojet engine, which includes: an outer shroud (14) intended for defining a secondary flow space of a stream of secondary gas on the inside and an inter-flow area on the outside, the outer shroud (14) being provided with a secondary opening (29), and a bleed valve comprising an outlet pipe (30) made of composite material, located in the inter-flow area, wherein the outlet pipe (30) is attached to the outer shroud (14) at the secondary opening (29), at least one gasket (33) for sealing against air and fire being arranged between the outlet pipe (30) and the outer shroud (14), and the outlet pipe (30) made of composite material includes a draped composite wall (30a, 30b), made up of a plurality of folds impregnated with resin.
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
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
75.
FIBROUS STRUCTURE FOR REINFORCING COMPONENTS MADE OF COMPOSITE MATERIAL WITH A GREAT VARIATION IN THICKNESS
A fibrous structure (200) comprises a plurality of weft layers (tl-t34) and of warp layers (C1-C12) connected to one another in a three-dimensional or multi-layer weave, the fibrous structure (200) comprising at least first and second parts (203, 204), which are adjacent in the warp direction, the first part (203) having, in a direction perpendicular to the warp and weft directions, a thickness that is greater than the thickness of the second part (204). Weft layers (tl3-t22) situated in the heart (2031) of the first part (203) of the fibrous structure (200) comprise braids (10). The weft layers (tl-tl2; t23-t34) extending on each side of the weft layers (tl3-t22) comprising the braids (10) and up to the skin (2032; 2033) of the first part (203) comprised threads or strands (20; 21; 22), the braids (10) having a cross section greater than the cross section of the threads or strands (20; 21; 22).
D03D 11/00 - Double or multi-ply fabrics not otherwise provided for
D03D 15/43 - Woven fabrics characterised by the material, structure or properties of the fibres, filaments, yarns, threads or other warp or weft elements used characterised by the structure of the yarns or threads with differing diameters
D03D 25/00 - Woven fabrics not otherwise provided for
76.
CONSTANT-VOLUME COMBUSTION MODULE FOR A TURBINE ENGINE, COMPRISING COMMUNICATION-BASED IGNITION
The invention relates to a turbine engine combustion module (10), in particular for an aircraft turbine engine, designed to carry out constant-volume combustion, comprising: at least two combustion chambers (12A, 12B) arranged about an axis, each chamber (12A, 12B, 12C) comprising a compressed gas intake port (16) and a burnt gas exhaust port (18); and an ignition means that triggers combustion in the combustion chambers (12A, 12B, 12C). The module (10) comprises at least one duct (80) which establishes a communication between a first combustion chamber (12A) and at least one second combustion chamber (12B) in order to inject burnt gases from the first combustion chamber (12A) into the second combustion chamber (12B) so as to trigger combustion in the second combustion chamber (12B).
F02C 5/12 - Gas-turbine plants characterised by the working fluid being generated by intermittent combustion the combustion chambers having inlet or outlet valves, e.g. Holzwarth gas-turbine plants
F23R 3/02 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
F23R 3/48 - Flame tube interconnectors, e.g. cross-over tubes
F23R 7/00 - Intermittent or explosive combustion chambers
77.
CONSTANT-VOLUME COMBUSTION MODULE FOR A TURBINE ENGINE
The invention relates to a turbine engine combustion module (10), in particular for an aircraft turbine engine, designed to carry out constant-volume combustion, comprising: a plurality of combustion chambers (12) distributed angularly in a regular manner about an axis (A), each chamber (12) comprising a compressed gas intake port (16) and a burnt gas exhaust port (18), each intake (16)/exhaust (18) port being designed to be opened or closed by a corresponding common intake (28)/exhaust (30) rotary valve coaxial with the axis (A). The invention also relates to a turbine engine comprising a combustion module (10) of the type described.
F02C 5/12 - Gas-turbine plants characterised by the working fluid being generated by intermittent combustion the combustion chambers having inlet or outlet valves, e.g. Holzwarth gas-turbine plants
F23R 3/02 - Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
F23R 3/48 - Flame tube interconnectors, e.g. cross-over tubes
F23R 7/00 - Intermittent or explosive combustion chambers
78.
PHANTOM INTENDED FOR USE IN QUALITY CONTROL OF TOMOGRAPHIC IMAGES.
The invention relates to a phantom (2) intended for use in quality control of tomographic images, comprising a cylindrical plate (4) produced from a homogeneous material having a density d1, two cylinders (6) inserted in the plate, the cylinders being produced from homogeneous materials having different densities d2 and d3, the density of one of the cylinders being greater than the density d1 of the plate and the density of the other cylinder being less than the density d1 of the plate, and a first series of pairs of holes (10i to 10n) of different diameters drilled in the plate, an axis of the holes of the first series being orientated axially with respect to an axis of rotation of the plate and the holes of a given pair being spaced apart by a distance equal to the diameter of said holes.
G01N 23/046 - 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 tomography, e.g. computed tomography [CT]
79.
METHOD FOR MANUFACTURING A REFRACTORY PART MADE OF COMPOSITE MATERIAL
The invention relates to a method for manufacturing a part made of composite material, which includes the following steps: forming a fibrous texture (10) from refractory fibres; placing the fibrous texture (10) in a mould (110) including an impregnation chamber (101) comprising in the lower portion thereof a part made of porous material (120), wherein the impregnation chamber (101) is closed at the top thereof by a deformable impervious membrane (140) separating the impregnation chamber (101) from a compaction chamber (102); injecting a slurry (150) containing a powder of refractory particles into the impregnation chamber (101); injecting a compression fluid (160) into the compaction chamber (102) so as to force the slurry (150) to pass through the fibrous texture (10); draining the liquid from the slurry via the part made of porous material (120) and maintaining the powder of refractory particles inside said texture so as to obtain a fibrous preform (20) filled with refractory particles; drying the fibrous preform (20); removing the fibrous preform (20) from the mould; and sintering the refractory particles contained in the fibrous preform so as to form a refractory matrix in said preform.
The invention relates to a method for manufacturing a fibrous preform filled with refractive ceramic particles (20), comprising the following steps: a) placing a fibrous structure comprising refractive ceramic particles into a mould cavity (2) defined by a mould (3) and a counter mould (4; 4'), b) injecting a slip, comprising a powder that has refractive ceramic particles (20) and is present in a liquid medium, into the porosity of the fibrous structure (1) in the mould cavity (2), the injection being carried out at least through a first surface (1a) or a first edge (1a') of said fibrous structure (1), and c) draining, through a component (5) made of porous material, the liquid medium from the slip that has penetrated the fibrous structure (1), the component made of porous material having a thickness of greater than or equal to 0.1 mm, the drainage being carried out at least through a second surface (1b) or a second edge of the fibrous structure (1) that is different from the first surface (1a) or the first edge (1a'), said component (5) made of porous material further allowing the powder having refractive particles (20) to be retained in the porosity of the fibrous structure (1) in order to obtain a fibrous preform filled with refractive particles (20), the component (5) made of porous material being between the mould (3) and the fibrous structure (1) or between the counter mould (4; 4') and the fibrous structure (1), and the component (5) made of porous material being between at least one outlet vent (16) and the fibrous structure (1), pumping being carried out in the region of the outlet vent (16) in order to drain the liquid medium through said outlet vent (16).
B28B 13/02 - Feeding the unshaped material to moulds or apparatus for producing shaped articles
B28B 23/00 - Arrangements specially adapted for the production of shaped articles with elements wholly or partly embedded in the moulding material
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
C04B 35/622 - Forming processesProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products
C04B 35/80 - Fibres, filaments, whiskers, platelets, or the like
The invention relates to a method for manufacturing a part made of composite material, said method including the following steps: a step of forming a fibrous texture (10) from refractory fibres; a first impregnation step of impregnating the fibrous texture (10) with a first slurry (150) containing first refractory particles (151); a step of removing the liquid phase (152) from the first slurry (151) so as to allow only the first refractory particles (151) to remain inside said texture; a second impregnation step of impregnating the fibrous texture (20) with a second slurry (160) containing second refractory particles (161); a step of removing the liquid phase (162) from the second slurry (160) such as to allow only the second refractory particles (161) to remain inside said texture and to obtain a fibrous preform (30) filled with the first and second refractory particles (151, 161); and a step of sintering the first and second refractory particles (151, 161) in the fibrous preform (30) so as to form a refractory matrix inside said preform.
The invention relates to a diffuser vane (2) for a gas turbine engine, comprising a vane assembly (4) made of a composite material having a fibrous reinforcement densified by a matrix, the fibrous reinforcement being obtained from long fibres pre-impregnated and agglomerated in the form of a mat, the vane assembly being provided on at least one leading edge with a reinforcement strip (10-1), said reinforcement strip being made of a single strip of a unidirectional fabric or a textile, or of a stack of a plurality of pre-impregnated plies of a unidirectional fabric or a textile made of carbon fibres or glass fibres, and at least one platform (6, 8) positioned at a radial end of the vane assembly, the platform being made of a composite material having a fibrous reinforcement densified by a matrix, the fibrous reinforcement being obtained from pre-impregnated long fibres. The invention also relates to a method for manufacturing such a vane.
B29C 70/20 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in a single direction, e.g. roving or other parallel fibres
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 70/78 - Moulding material on one side only of the preformed part
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
The invention relates to a pivot link assembly (1) including a shaft (60) and first and second linking portions (12A, 40), each receiving the shaft (60) so as to enable the pivoting of said linking portions relative to one another. The first linking portion (12A) is axially arranged between a first end (62A) of the shaft (60) and a radial plane (PA1), whereas the second linking portion (40) is axially arranged between said radial plane (PAI) and the other end (62B) of the shaft (60). The shaft (60) is hollow and comprises an inner cross-section that has, in at least one first axial plane (Q) on a first section (T1) of the shaft (60) that is axially defined between a first end (62A) and the radial plane (PAI), a variable radial dimension (DA) that decreases as it axially approaches the radial plane (PAI). The assembly (1) includes a ring (20A, 20A') that is radially inserted between the shaft (60) and a bore of the first linking portion (12A) and has a variable radial thickness (E, E1, E2).
METHOD FOR MANUFACTURING A PART MADE OF A COMPOSITE MATERIAL COMPRISING AT LEAST ONE PORTION FORMING A FORCE-INSERTION PORTION OR LOCAL THICKENED PORTION
Method for manufacturing a part made of a composite material, comprising precompaction of a first preform (220) of staple fibres having a length of between 8 mm and 100 mm. The first preform is positioned with a second preform (210) comprising a fibrous texture of staple fibres 15 impregnated with a second thermosetting resin and the polymerization of the first and second preforms. This makes it possible to make complex geometric parts out of composite materials that have no areas of weakness at the bonding interface between the subassemblies of the part.
The invention relates to a method and device for digital reconstruction of an elementary volume representing a microstructure of composite material. The method includes: defining (F10) an elementary volume; filling the volume with digital elements modelling elements of fibres of the composite material, and extending longitudinally along a main axis, which includes: associating (F30) with each element a position in a plane of the space and an orientation of the main axis thereof; and consecutively positioning each element in the volume, in accordance with the position and the orientation which have been associated with same, said positioning including placing the element in contact (F40) with at least one wall of the elementary volume and/or at least one previously positioned element, and geometrically adapting (F50) the element to said at least one wall and/or to said at least one previously positioned digital element with which same is placed in contact, at least one portion of an element used to fill the volume undergoing, during the geometric adaptation, a deformation other than a tilting of the longitudinal axis thereof relative to the main axis of said element.
The method for estimation of a quality index of a 3-D image encoded in grey level of a piece of composite material, comprises: a step for obtaining (E20) a histogram (HIST) from the image, said histogram representing for each class of a plurality of classes each comprising at least one grey level, a number of voxels of the image having a grey level belonging to said class; an extraction step (E40) of a predetermined number of Gaussians present in the histogram; an estimation step (E60) for estimating at least one quality index of the image from parameters characterizing the Gaussians extracted from the histogram.
G01N 23/046 - 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 tomography, e.g. computed tomography [CT]
87.
FIRE PROTECTION OF A PART MADE OF A THREE-DIMENSIONAL WOVEN COMPOSITE MATERIAL
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).
The invention concerns a method and a system for non-intrusively measuring the volume density of a specific phase in a part, comprising: - processing means (9) for producing a three-dimensional image (17) of said part (11), said image being formed by a three-dimensional grid of voxels of which the values indicate the arrangement of said specific phase in said part, - processing means (9) for associating a binary coefficient with each voxel of said three-dimensional image, thus constructing an initial three-dimensional matrix representation of binary coefficients, said binary coefficients being representative of a presence or absence of said specific phase in areas of said part corresponding to the voxels, - processing means (9) for convolving said initial matrix representation with a matrix convolution kernel corresponding to a predefined reference volume, said convolution being performed by implementing a composition of three (successive) one-dimensional convolutions in three separate directions, thus forming a resulting matrix representation of which each resulting coefficient represents a volume level (density) of said specific phase in said reference volume.
G01N 23/046 - 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 tomography, e.g. computed tomography [CT]
89.
METHOD FOR IMPREGNATION OF A FIBROUS PREFORM AND DEVICE FOR IMPLEMENTATION OF THE SAID METHOD
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 43/10 - Isostatic pressing, i.e. using non-rigid pressure-exerting members against rigid parts or dies
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
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.
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
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
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).
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
D03D 25/00 - Woven fabrics not otherwise provided for
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
92.
EQUIPMENT FOR ATTACHING A METAL REINFORCEMENT ONTO THE LEADING EDGE OF A TURBINE ENGINE BLADE, AND METHOD UTILIZING SUCH EQUIPMENT
The invention relates to equipment for attaching a metal reinforcement (300) onto the leading edge (6) of a turbine engine blade (2), which includes a blade mounting (100) for receiving a blade while leaving surfaces of the leading edge of the blade exposed, and a mounting (200) for reinforcing the leading edge, onto which the blade mounting (100) is to be mounted, said mounting for reinforcing the leading edge including two side shims (208), between which the metal reinforcement (300) of the leading edge of the blade is positioned. The shims (208) can be moved closer and farther from one another and are each provided with a downdraft grate for grasping the metal reinforcement (300). The mounting (200) reinforcing the leading edge also includes heating elements for enabling polymerization of an adhesive film applied onto surfaces of the leading edge of the blade. The invention also relates to a method for attaching a metal reinforcement onto the leading edge (6) of a turbine engine blade (2) by means of such equipment.
The invention relates to an injection mould (4) for manufacturing a rotary part made of a composite material having external flanges, including a mandrel (2) on which a fibrous reinforcement (9) is intended to be supported and comprising a central wall, (6) the profile of which matches that of the part to be manufactured, two side plates (8), moulding wedges (12) for engaging with a non-covered surface of the fibrous reinforcement, and two external bells (14) for covering the moulding wedges and the side plates of the mandrel, sealing O-rings (24) being inserted between the bells and the side plates of the mandrel, the bells each being provided with an attachment flange (20) for engaging with one another, at least one sealing O-ring (26) being inserted therebetween, and the attachment flanges of the bells being clamped against one another.
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]
Fibrous preform for a turbo machine blade, which preform is obtained by one-piece three-dimensional weaving. According to the invention, the preform comprises a first longitudinal portion (41) able to form a blade root, a second longitudinal portion (42) extending the first longitudinal portion and able to form a blade section, and a first transverse portion (51) extending transversely from the junction between the first and second longitudinal portions (41, 42) and able to form a first platform.
The method for producing a propeller blade comprises the following steps -forming a spar (1) comprising a blade core (2); forming a blade preform (6) comprising a fibre fabric, said fabric having a separating area (7) where it is divided into two superposed skins (26, 27), the separating area comprising a location intended to receive the blade core; installing the blade preform in a tool (10) comprising an imprint (11) of said preform; opening the separating area (7) by moving the skins (26, 27) apart; installing the spar (1) by inserting the blade core (2) into the separating area (7); reclosing the separating area by moving the skins back together; injecting resin into the blade preform; polymerising the resin and machining the blade preform; a step for creating tracers (14), which are marks at predetermined positions on the blade preform (6), while forming said blade preform; a step for shaping the blade preform (6) in a predetermined position in the imprint after the step of installing the blade preform, matching the tracers (14) with target positions defined relative to the tooling, and a step of clamping the blade preform (6) in the imprint outside the separating area (7) before opening the separating area and until the spar (1) is installed.
B29B 11/16 - Making preforms characterised by structure or composition comprising fillers or reinforcements
B29C 31/08 - Feeding, e.g. into a mould cavity of preforms
B29C 33/00 - Moulds or coresDetails thereof or accessories therefor
B29C 70/08 - Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, with or without non-reinforced layers
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 70/54 - Component parts, details or accessoriesAuxiliary operations
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
B29D 99/00 - Subject matter not provided for in other groups of this subclass
patents
