The method for manufacturing a blade (32) of a turbomachine (100) comprises: —manufacturing a part (4) comprising an air flow path zone (10) and a layer (20) covering the zone, the layer having cavities (21) forming a periodic pattern, the manufacture taking place by injecting a mixture comprising a binder and a power; —removing a larger part of the binder from the part; —sintering the part, and —removing the layer (20) from the part to obtain the blade (32).
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 3/24 - After-treatment of workpieces or articles
2.
METHOD FOR MAINTAINING A BLADED WHEEL OF A HIGH-PRESSURE TURBINE OF A TURBOMACHINE
A method for maintaining a bladed wheel of a high-pressure turbine of a turbomachine having a disc with cavities, and blades extending radially from a radially internal root mounted in a cavity of the disc, the root being supported on the disc by means of surfaces of the root and of the disc forming bearing surfaces. At least one foil is removably mounted between the root of at least one blade and the disc, at the corresponding bearing surfaces. The method includes measuring a radial clearance between the tip of at least one blade, meaning the radially external end of the blade, and a shroud of the turbomachine situated facing the bladed wheel, and if the radial clearance is greater than a determined value, removing the foil.
The invention relates to a module for a turbine engine comprising a shaft extending in an axial direction and an electric machine (110) comprising: - a rotor (120) rotatably coupled with the shaft and comprising a disc (122) and magnetic elements (124) arranged evenly at the circumference of the disc; and - a stator (130) rigidly connected to a casing of the module comprising a ring (132) and coils distributed in an annular manner inside the ring of the stator, the coils being arranged outside the magnetic elements of the rotor in a radial direction, wherein the rotor has a balancing defect represented by a mechanical force (F1) when the magnetic elements are in a first position, and at least one magnetic element (124) is movable from the first position to a second position suitable for applying to the rotor a magnetic force (F2) opposing the mechanical force when the electric machine is in operation.
The invention relates to a turbine engine compressor casing (300) which comprises an inner annular wall (310) and an outer annular wall (320) defining therebetween an annular cavity (330), the inner annular wall of the casing being provided with first slots (315). The casing comprises a movable ring (200) comprising the outer annular wall (320) and a plurality of sealing bars (230) arranged opposite the outer face (312) of the inner annular wall (310), the bars defining therebetween second slots (240). The movable ring (200) slides between an open position in which the second slots (240) are aligned with the first slots (315) and a closed position in which the sealing bars (230) close the first slots (315).
F01D 11/10 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
F01D 11/22 - Actively adjusting tip-clearance by mechanically actuating the stator or rotor components, e.g. moving shroud sections relative to the rotor
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
F04D 29/42 - CasingsConnections for working fluid for radial or helico-centrifugal pumps
F04D 29/52 - CasingsConnections for working fluid for axial pumps
F04D 29/68 - Combating cavitation, whirls, noise, vibration, or the likeBalancing by influencing boundary layers
5.
METHOD FOR DETECTING DAMAGE TO A PART MADE OF CONDUCTIVE MATERIAL
The invention relates to a method (300) for detecting damage (201) to a part (200) made of conductive material, the damage (201) being a burn or a crack, the part (200) being heated (310) by induction by injecting an energy pulse. The method comprises receiving (330) a temporal sequence of thermal images of the part after heating the part (200), and the temporal signal of the intensity of each pixel of the thermal image is used to construct (370) a phase image of the part by means of a discrete Fourier transformation of the temporal signal.
The invention relates to a fluid passage system (100) for a partition of an aircraft turbine engine, comprising a plate (110) configured to be attached to the partition and characterised in that: - the plate (110) comprises a plurality of fluid passage channels, the channels being joined together at least two by two; - the peripheral wall of each channel extends on either side of the plate to form a first pipe (130) and a second pipe (140); and - the plate (110) and the first and second pipes (130, 140) form a single integral part.
A method for locating a mark on a casing of a turbomachine of an aircraft, wherein the method includes a) a step of visually inspecting the arms, the outer surface of the internal shroud and the inner surface of the external shroud, and, in the event that at least one mark is detected at step a), b) a step of manually mounting a visual inspection device around the inner end or outer end of one or more of the arms, and c) a step of determining the location of the at least one mark in a zone of the casing.
G01B 5/14 - Measuring arrangements characterised by the use of mechanical techniques for measuring distance or clearance between spaced objects or spaced apertures
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
An aeronautical thruster of longitudinal axis has a hub, an annular row of unducted upstream rotor blades and an annular row of unducted downstream stator blades. Each downstream stator blade is of variable pitch, and at least one of the downstream stator blades is in a closed-pitch configuration relative to another of the downstream stator blades in that it has a pitch angle smaller than the pitch angle of the other downstream stator blade.
Methods, apparatus, systems and articles of manufacture are disclosed. An apparatus for mounting a gas turbine engine to a pylon, the gas turbine including an upstream section and a downstream section, the gas turbine defining a roll axis, a yaw axis, and a pitch axis, the apparatus including: a first mount to couple the upstream section of the gas turbine engine to the pylon; a second mount to couple the upstream section of the gas turbine engine to the pylon, the second mount downstream of the first mount; a thrust linkage to couple the upstream section to the pylon, wherein the downstream section is decouplable from the upstream section without decoupling the first mount, the second mount, and the thrust linkage.
The invention relates to a module comprising: - an oil tank (22), - an oil supply circuit (23) comprising: - a first pump (25) comprising an oil inlet (26), which is connected to the tank (22), and an oil outlet (27), and - a first heat exchanger (28) comprising a first circuit for oil and a second circuit for a cooling fluid, - a cooling circuit (24) comprising: - a second pump (30) comprising an oil inlet (32), which is connected to the tank (22), and an oil outlet (33), - a second heat exchanger (31) comprising a first circuit for oil and a second circuit for a cooling fluid, and - a control member (34) controlling the second pump (30) and configured to regulate the flow rate of the second pump (30).
The invention relates to a blade (6, 7) for an aircraft turbomachine, the blade (6, 7) comprising: - an airfoil (9) having a pressure face (11i) and a suction face (11e) which are connected by a leading edge (11a) and a trailing edge (11b), the airfoil (9) comprising a first metal material, - at least one heating line (13) extending along the airfoil (9), the heating line (13) comprising at least one electrically conductive body (17) comprising a second metal material, characterized in that the heating line (13) is integrated into the airfoil (9) and forms a monolithic assembly with the airfoil (9).
The invention relates to a turbomachine module comprising: - a rotor (4) of longitudinal axis (X), - a plurality of variable-pitch rotor blades (3) mounted on the rotor and extending radially, the rotor blades pivoting between a first position and a second position, - a splitter (9) separating an air flow passing through the rotor blades into a primary flow (F1) and a secondary flow (F2), the splitter comprising an upstream edge (14) delimiting an inlet (17) of a flow path (10) in which the primary flow circulates, and - stator vanes (18) arranged at the inlet (17). According to the invention, the module comprises fins (21) mounted so as to rotate as one with the rotor and extending radially, the fins being arranged downstream of the rotor blades and upstream of the stator vanes, the fins being configured to divert the air flow at the outlet of the rotor blades towards the stator vanes.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F04D 29/32 - Rotors specially adapted for elastic fluids for axial-flow pumps
13.
GAS TURBINE ENGINE BLADING COMPRISING A BLADE AND A PLATFORM WHICH HAS AN INTERNAL FLOW-INTAKE AND FLOW-EJECTION CANAL
The present invention relates to blading (25, 26) for a turbomachine (10), comprising: —a blade (31) having an aerodynamic profile; —a platform (32, 33) comprising a flow-path surface (321) intended to delimit a primary flow path (21A) of the turbomachine (10), which path is intended, when the turbomachine (10) is in operation, to receive a flow that splits, upstream of the blade (31), into a suction-face flow (EE) and a pressure-face flow (EI); and —an internal canal (34) which has an intake opening (35) and an ejection opening (36), these each opening onto the flow-path surface (321) of the platform (32, 33), the ejection opening (36) opening downstream of the intake opening (35) and the intake opening (35) opening toward the pressure-face flow (EI).
An aeronautical thruster of longitudinal axis includes a hub, an annular row of unducted upstream rotor blades, and an annular row of unducted downstream stator blades. The annular row of downstream stator blades includes at least one downstream stator blade of a first type, each of which is located about the longitudinal axis in a first angular sector about the longitudinal axis. Each downstream stator blade of the first type has a fixed pitch. The annular row of downstream blades further includes at least one downstream stator blade of a second type, each of which is located about the longitudinal axis outside said first angular sector, wherein each downstream stator blade of the second type has a variable pitch.
A method for controlling a motor assembly. The motor assembly includes a gas turbine engine and at least one electric machine which is mechanically coupled to a rotating shaft of the gas turbine engine so as to be rotated for generating electricity. In this control method, a mechanical power take-off setpoint by the electric machine is changed when an operating parameter of the gas turbine engine reaches a predetermined limit. A control unit adapted to perform this method, a motor assembly incorporating this control unit, the electric machine and the gas turbine engine, and a computer program for performing this method.
Turbojet engine fan module wheel, including a plurality of blades made of composite material, each blade having a root assembled with a base distinct from the bases of the other blades, each base having a groove extending axially and opening out on the side of the upstream face and on the side of the downstream face, each root cooperating by axial interlocking in a form-fitting manner, for example in the shape of a dovetail, with the groove of the base, whereby the root is retained on the base along the radial and circumferential directions, and each base cooperates with at least one part configured to axially block the root within the groove of the base, whereby the root is retained along the axial direction.
F02K 7/00 - Plants in which the working-fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fanControl thereof
B64C 11/06 - Blade mountings for variable-pitch blades
F01D 7/00 - Rotors with blades adjustable in operationControl thereof
F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
17.
PROTECTION FOR A CIRCUIT FOR CONTROLLING THE ORIENTATION OF PROPELLER BLADES OF AN AIRCRAFT ENGINE, ALLOWING THE ORIENTATION OF THESE BLADES TO BE LOCKED
The invention relates to a hydraulic control circuit (17) for actuating a double-acting orientation cylinder (21) for orienting blades (3) of a propulsion propeller of an aircraft engine, said double-acting cylinder (21) comprising a first and a second chamber (18, 19), and the circuit (17) comprising: - a high-pressure pipe (23) and a low-pressure pipe (24); - a distributor valve (26) for connecting the first chamber (18) to the high-pressure pipe (23) and the second chamber (19) to the low-pressure pipe (24), or vice versa, or for isolating the two chambers (18, 19) from the high-pressure and low-pressure pipes; - a controlled valve (36) arranged between the first chamber (18) and the distributor valve (26), and controlled by a high-pressure control line (37), so as to be locked in a closed state in the event of a drop in pressure in the high-pressure control line (37).
B64C 11/38 - Blade pitch-changing mechanisms fluid, e.g. hydraulic
F02K 1/66 - Reversing fan flow using reversing fan blades
F01D 1/30 - Non-positive-displacement machines or engines, e.g. steam turbines characterised by having a single rotor operable in either direction of rotation, e.g. by reversing of blades
F01D 7/00 - Rotors with blades adjustable in operationControl thereof
One aspect of the invention relates to a method (100) for monitoring the weaving of a workpiece by an electronic weaving machine, the method comprising: - obtaining (110) configuration and weaving setpoint data delivered to the electronic weaving machine in order to weave the workpiece; - generating (120) a digital model of the electronic weaving machine by aggregating the obtained configuration and setpoint data; - delivering (140) the generated digital model of the electronic weaving machine (120) to the machine learning model in order to obtain a prediction of the tension exerted at the individual motors during the weaving of the workpiece; and - monitoring (150) the weaving of the workpiece by comparing the predicted tension exerted on an individual motor with the tension measured by the tension sensor associated with the individual motor.
The invention relates to a forming core comprising a liquid-soluble body and a polymerised resin skin coating the soluble body. The invention also relates to a method for obtaining the forming core. The invention also relates to an assembly of a forming core and of a hollow part (62) made of organic matrix composite material. The invention also relates to a hollow part (62) made of organic matrix composite material comprising a cavity (80), wherein the cavity (80) is covered with a polymerised resin skin (54). The invention also relates to a method for the manufacture of the hollow part (62).
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 33/52 - Moulds or coresDetails thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
B29C 70/54 - Component parts, details or accessoriesAuxiliary operations
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
20.
HOLLOW PROPELLER BLADE OR HOLLOW VANE MADE OF ORGANIC MATRIX COMPOSITE MATERIAL
The present invention relates to a hollow propeller blade (1) or hollow vane (1) made of organic matrix composite material, comprising (i) a first portion (3) defining an aerodynamic profile and comprising a first fibre reinforcement obtained by three-dimensional weaving, and (ii) a hollow second portion (5), located inside the first portion and rigidly attached to the first portion, wherein the second portion comprises a second fibre reinforcement comprising a fibre stack of at least (a) a first layer of unidirectional fibres oriented between -65° and -25° with respect to a radial direction (DR) of the vane or propeller blade; and (b) a second layer of unidirectional fibres oriented between +25° and +65° with respect to the radial direction. The present invention also relates to associated manufacturing methods.
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
B64F 5/10 - Manufacturing or assembling aircraft, e.g. jigs therefor
B29C 70/24 - Fibrous reinforcements only characterised by the structure of fibrous reinforcements using fibres of substantial or continuous length oriented in at least three directions forming a three dimensional structure
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
The invention relates to an assembly (1) comprising an aircraft turbine engine (2), a pylon (40) for mounting the turbine engine on an element of the aircraft, and a system for suspending the turbine engine from the pylon, the turbine engine having a longitudinal axis (X) and comprising, from upstream to downstream, in the gas flow direction: > a fan casing (11) centred on the longitudinal axis (X) and having a first axial end (12) located upstream of the fan (10); and > a gas generator (16) configured to receive an air flow generated by a fan, the pylon (40) being generally elongate along the longitudinal axis (X) and being able to be attached to the element of the aircraft, characterised in that the suspension system comprises at least one intake force uptake link (30, 30a, 30b), a first end (31, 31a, 31b) of which is connected to an attachment point (11a, 11b) of the fan casing (11) located on the first axial end or on an upstream portion of the fan casing, and a second, opposite end (32, 32a, 32b) of which is connected to the pylon.
The invention relates to an assembly (1) including an aircraft turbine engine (2), a pylon (40) for mounting the turbine engine on an element of the aircraft, and a system for suspending the turbine engine from the pylon, the turbine engine having a longitudinal axis (X) and including, from upstream to downstream, in the gas flow direction: > a fan casing (11) extending about the longitudinal axis; and > a gas generator (16) configured to receive an air flow generated by a fan, the pylon (40) being generally elongate along the longitudinal axis and capable of being attached to the element of the aircraft, characterised in that connecting elements for the suspension system are all located upstream of the combustion chamber (19).
The invention relates to a method (100) for manufacturing a bladed part (1) for an aircraft turbine engine, this bladed part being made of an organic matrix composite (OMC) material, the method comprising the following steps: - (110) providing a moulding core (10) made of a material having a melting point denoted T1, the moulding core having a general elongate shape along an axis of elongation (X2) and comprising a first positioning protrusion (20) located at least at one of its ends (11, 12); - (120) producing a first fibrous preform (50) by three-dimensional weaving, this first preform comprising a body (54) having a general elongate and tubular shape and comprising a cavity (57) which extends along an axis of elongation (X1) of the first preform and which opens out at the opposite axial ends (51, 52) of the body (54); - (130) inserting the moulding core into the cavity so that the axes of elongation of moulding core and cavity coincide, the first positioning protrusion (20) extending out of the cavity.
B29C 33/52 - Moulds or coresDetails thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
B29C 70/24 - 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 70/46 - 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
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
F01D 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
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
The invention relates to a forming core (50) comprising a liquid-soluble body (52) and a skin (54) comprising at least one layer of a mould release agent coating the soluble body (52). The invention also relates to a method for obtaining the forming core. The invention also relates to a hollow part (62) made of organic matrix composite material comprising a cavity (80), wherein the cavity (80) is covered with a skin (54) comprising at least one layer of a mould release agent coating the soluble body. The invention also relates to a method for the manufacture of the hollow part (62).
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
B29C 33/60 - Releasing, lubricating or separating agents
B29C 70/54 - Component parts, details or accessoriesAuxiliary operations
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 33/52 - Moulds or coresDetails thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
B29L 31/08 - Blades for rotors, stators, fans, turbines or the like, e.g. screw propellers
25.
ENGINE, IN PARTICULAR FOR A TURBOMACHINE, COMPRISING A SEALING DEVICE COMPRISING A DYNAMIC SEALING TRACK SUPPORT WITH AN INTEGRATED CHANNEL
An engine (10) comprising a rotor shaft, an air chamber, a lubricant chamber comprising lubricant and at least one sealing device (20) arranged between said chambers and comprising an annular radial seal (21) associated with a seal track (31) borne by a dynamic sealing track support (30) secured to the rotor shaft (12). The dynamic sealing track support (30) comprises at least one lubricant distribution channel (37) configured to recover lubricant contained in the lubricant chamber and spray it by the effect of centrifugal force onto an inner heat exchange surface (35a) in the region of a spray zone located axially to the outside of the seal track (31) with respect to the lubricant chamber (15).
26.
METHOD FOR MANUFACTURING A HOLLOW PART MADE OF COMPOSITE MATERIALS AND COMPRISING INTERNAL STIFFENERS
The invention relates to a method for manufacturing a composite part (10, 20), comprising the steps of: producing a fibre preform (36) of the composite part; manufacturing mandrels (31-34) from a moulded soluble paste; forming a fibre reinforcement (35) around each mandrel; inserting the assembled mandrels into the preform at a location for a cavity (3) to be formed in the composite part; inserting the preform containing the mandrels into a mould; injecting a resin into the mould to obtain the composite part, the resin diffusing into the preform and the fibre reinforcements; demoulding the composite part; forming a channel (38) in the composite part in order to reach the mandrels; and injecting a solvent into the channel in order to dissolve the mandrels, thereby forming the cavity in the composite part, wherein the fibre reinforcements between the mandrels form stiffeners in the cavity.
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 9/04 - NozzlesNozzle boxesStator bladesGuide conduits forming ring or sector
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
B29D 99/00 - Subject matter not provided for in other groups of this subclass
B29C 33/52 - Moulds or coresDetails thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
B29C 70/54 - Component parts, details or accessoriesAuxiliary operations
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
The invention relates to a turbine engine (2) for an aircraft, this turbine engine (2) comprising: - an annular flow path (V) for a gas flow (F3); - a motor (M); - at least one piece of equipment (E); - a first ACOC-type heat exchanger (210), which is capable of being swept by the gas flow (F3) and which comprises a first oil circuit (C1) connected to a cooling system (S1) of the motor; - a second ACOC-type heat exchanger (220), which is capable of being swept by the gas flow (F3) and which comprises a second oil circuit (C2) connected to a system (S2) for cooling the at least one piece of equipment (E); and - at least one third ACOC-type heat exchanger (230; 230') which is capable of being swept by the gas flow (F3) and which comprises at least one third oil circuit (C3, C3').
F02C 7/14 - Cooling of plants of fluids in the plant
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F02K 3/077 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type the plant being of the multiple flow type, i.e. having three or more flows
F02K 3/115 - Heating the by-pass flow by means of indirect heat exchange
A system including: a blade including, at the level of the blade root, two portions spaced apart from each other in the transverse direction X so as to provide therebetween a cavity extending downwards in the direction Y of the height of the blade, from a cavity bottom to a cavity opening on the outside, located at the level of the lower end of the blade root, an attachment element of the blade partially engaged inside the cavity, the engaged part having a shape, taken in a plane defined by the directions X and Y, which extends towards the bottom while flaring out in the direction X, the two spaced apart portions of the blade root which are in contact with the flared shape of the attachment element having corresponding flared shapes.
The invention relates to a nickel-based alloy powder, that comprises in weight percentages, 14.00 to 15.25% of chromium, 14.25 to 15.75% of cobalt, 4.00 to 4.60% of aluminium, 0 to 0.50% of iron, 0 to 0.15% of manganese, 3.00 to 3.70% of titanium, 3.90 to 4.50% of molybdenum, 0 to 0.015% of sulphur, 0 to 0.06% of zirconium, 0.012 to 0.020% of boron, 0 to 0.20% of silicon, 0 to 0.10% of copper, 0 to 150 ppm of carbon, 0 to 0.5 ppm of bismuth, 0 to 5 ppm of lead, 0 to 1000 ppm of platinum, 0 to 1000 ppm of palladium, 0 to 50 ppm of hydrogen, 0 to 5 ppm of silver, 0 to 120 ppm of nitrogen, 0 to 1000 ppm of rhenium, 0 to 410 ppm of oxygen and 0 to 500 ppm of inevitable impurities, the rest being made up of nickel, and has a particle size D10 between 3 and 10 μm, a particle size D90 between 20 and 40 μm and a particle size D50 between 10 and 20 μm, the values of the particle sizes D10, D50 and D90 having been measured by laser diffraction according to standard ISO 13322-2. The invention also relates to a method for manufacturing a part using said powder and a resulting part.
B22F 1/10 - Metallic powder containing lubricating or binding agentsMetallic powder containing organic material
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 3/24 - After-treatment of workpieces or articles
The invention relates to a method for manufacturing a propeller blade or vane (3), the method comprising: - producing a fibre blank (100) comprising a gap (103) separating a first skin (110) from a second skin (120), wherein the gap (103) extends at least partially from an edge (100a) intended to form the leading edge; - inserting an insert (40) into the internal cavity (103a) of the fibre blank (100) so as to form a fibre preform (10); - stitching the first skin (110) to the second skin (120); - densifying the preform (10) in order to obtain a composite material part (1) having the shape of the blade or vane (3) to be obtained; - adding a leading edge reinforcement element (2) at least partially covering the one or more stitched regions (61, 62).
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 70/86 - Incorporating in coherent impregnated reinforcing layers
B29D 99/00 - Subject matter not provided for in other groups of this subclass
The present disclosure relates to a fan section having a performance coefficient of between 1.05 and 1.3, the performance coefficient being defined as follows: (Cp), where: Cp is the performance coefficient; FPR is the pressure ratio of the fan section (2); Cs is the strength of the fan rotor (9); and U is the peripheral speed at the blade tip of the fan rotor (9); the pressure ratio and the peripheral speed being measured at cruising speed.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
The invention relates to a propulsion assembly (1) comprising a turbomachine having a longitudinal axis, a nacelle, a first element (E1), a second element (E2) movable in translation with respect to the first element in a direction parallel to the longitudinal axis between a first and a second position and a fluidic circuit (70), the nacelle comprising at least one element from among the first and second elements, the propulsion assembly being characterized in that: the fluidic circuit (70) comprises a first duct (71) and a second duct (72) secured respectively to the first and second elements and respectively comprising first and second interlocking connectors (73, 74) configured to cooperate with one another so as to allow the fluid to flow between the first and second ducts when the second element is in the first position and to prevent the flow of the fluid when the second element is in the second position.
F02C 7/14 - Cooling of plants of fluids in the plant
F02C 7/18 - Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
F02K 1/72 - Reversing fan flow using thrust reverser flaps or doors mounted on the fan housing the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow
The invention relates to a part for an aircraft turbine engine (100), comprising: - a substrate (10) made of a nickel superalloy consisting of a γ phase and a γ' phase; - a coating (20) covering the substrate made of an alloy comprising, for more than 95% by volume, a γ' phase and having a chromium atomic content of between 2% and 5% and an aluminium atomic content of between 18% and 25%, wherein nickel accounts for the majority of the alloy; and - a thermal barrier layer (30); wherein the coating is arranged between the substrate and a thermal barrier layer.
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 14/00 - Coating by vacuum evaporation, by sputtering or by ion implantation of the coating forming material
The invention relates to a part for an aircraft turbine engine (100) comprising: - a substrate (10) made of a nickel-based superalloy consisting of a γ phase and a γ' phase, wherein the substrate havs a chromium content by weight of less than or equal to 7.0%; and - a coating (20) covering the substrate made of an alloy consisting of a γ phase and a γ' phase and comprising a chromium content by weight of between 6.0% and 14% and an aluminium content by weight of between 6.0% and 12%, and a platinum content by weight of between 0% and 17%, wherein nickel accounts for the majority of the alloy.
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 30/00 - Coating with metallic material characterised only by the composition of the metallic material, i.e. not characterised by the coating process
The invention relates to a turbine engine (2) for an aircraft, the turbine engine comprising: - an annular flow path (V) for a gas flow (F3); - a motor (M); - at least one piece of equipment (E); - a first ACOC-type heat exchanger (210), which is capable of being swept by the gas flow (F3) and which comprises a first oil circuit (C1) connected to a cooling system (S1) of the motor; - a second ACOC-type heat exchanger (220), which is capable of being swept by the gas flow (F3) and which comprises a second oil circuit (C2) connected to a system (S2) for cooling the at least one piece of equipment (E); and - a third, OCOC-type heat exchanger (230), which comprises third and fourth oil circuits (C3, C4) that are capable of exchanging heat energy with one another.
The invention relates to a positioning and/or holding system (10) comprising: - a body (11) provided with at least one through-hole (12); and - at least one attachment device (13) comprising: - a pin (14) inserted into the through-hole (12) and comprising at least one threaded intermediate section (18) and a threaded end (23), which is intended to be screwed partially into a prevailing torque nut (24) floatingly mounted on a structural element (25); and - a locknut (28) mounted on the threaded intermediate section (18) of the pin (14), such that tightening the locknut (28) against the body (11) is capable of bringing about an axial movement of the pin (14) in order to apply a force along an axis of the pin (14) so as to press the prevailing torque nut (24) against the structural element (25)
The present invention relates to a tool (4) for mounting a degassing tube in a hollow shaft of an aircraft turbine engine, the tool comprising a tubular body (5) that has an elongate shape along an axis (A), and comprising: - an internal longitudinal passage (50) extending over the entire longitudinal dimension of the body (5) and configured to accommodate the degassing tube; and - a longitudinal slot (51) that extends over the entire longitudinal dimension of the body (5) and which opens into the passage (50), wherein this slot (51) is configured to allow the degassing tube to be inserted into the passage (50) by translationally moving the degassing tube through the slot (51) in the transverse direction, and wherein the body (5) is formed by assembling a first axial segment (6) made of a first material and a second axial segment (7) made of a second material different from the first material.
The present invention relates to a tool (4) for mounting a degassing tube in a hollow shaft of an aircraft turbine engine, the tool comprising a tubular body (5) that has an elongate shape along an axis (A), and comprising: - an internal longitudinal passage (50) extending over the entire longitudinal dimension of the body (5) and configured to accommodate the degassing tube; and - a longitudinal slot (51) extending over the entire longitudinal dimension of the body (5) and opening into the passage (50), wherein this slot (51) is configured to allow the degassing tube to be inserted into the passage (50) by transationally moving the degassing tube through the slot (51) in the transverse direction, wherein the body (5) comprises a first axial segment (6) and a second axial segment (7) having a different shape from the first axial segment, and wherein the first and second axial segments are formed as a single piece.
The invention relates to a method for manufacturing a turbomachine component (2), comprising a wall (3) and a cellular structure (4) made of composite material on at least one portion of a face (30) of the wall (3), the cellular structure (4) comprising at least one first network of cells (42) and at least one first skin (44) covering this first network of cells (42), the method comprising the following steps of: (a) providing the wall (3); and (b) producing the cellular structure (4) on the face, step (b) of producing the cellular structure (4) comprising the following sub-steps of: (b1) depositing first plies of composite material by draping to form the first skin (44); and (b2) forming the first network of cells (42) by additive manufacturing, sub-step (b1) being carried out before or after sub-step (b2).
The invention relates to a mechanical part (100) comprising a core (110) having a fibrous reinforcement densified by a matrix, comprising at least one opening adjacent to the core (110) through which a shaft is intended to pass in order to establish a connection with another part, and comprising a belt (120) having a fibrous reinforcement densified by the matrix surrounding the core (110) and the at least one opening. According to the invention, the part (100) is characterized in that at least one insert (140) extending from the opening is positioned between the core (110) and the belt (120).
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 70/34 - 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 and shaping or impregnating by compression
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
41.
METHOD AND DEVICE FOR TWO-WAY COMMUNICATION OVER A SINGLE OPTICAL FIBRE WITH A SINGLE LIGHT SOURCE
The invention relates to a device for two-way communication between a first terminal (4) and at least one second terminal (6), the device including: - a single optical fibre (2) between the first terminal (4) and the at least one second terminal (6); - the first terminal (4) including a radiation source (8) of radiation having at least a first wavelength (λ1), a photodetector (10) and means (12) for directing a signal originating from the second terminal (6) via the optical fibre (2) towards the photodetector (10) and for directing a signal originating from the radiation source (8) towards the optical fibre (2); - the second terminal (6) including at least one modulator (24), a photodetector (22), and power supply means (9).
The invention relates to an optical spectral modulator device (1) comprising: • - at least one optical fibre (2) comprising optical functionalisation means (4) having a reflection spectrum of a portion of light radiation travelling through the fibre; • - a deformable element (10) having a length L; • - a first mechanical element (8) and a second mechanical element (14) that are attached to the deformable element (10) and capable of undergoing a relative deformation or a relative displacement as a result of this deformable element (10), wherein the optical fibre is attached to the mechanical elements at two attachment points (16, 18) arranged on either side of the optical functionalisation means (4), wherein the distance d between the first attachment point and the second attachment point is shorter than the length L of the deformable element, and wherein the ratio L/d is at least equal to 5.
The present invention relates to an aeronautical propulsion system (1), characterized in that (Eq1), (Eq2), (Eq3), (Eq4), where L is a distance between the inlet of the compressor (5) and the outlet of the turbine (8), D is a diameter of the fan rotor (9), n is the total number of stages between inlet and outlet, XN is the actual maximum rotational speed of the shaft (11); BPR is a bypass ratio of the system, when it is in the steady-state under takeoff conditions in a standard atmosphere at sea level, Te is the temperature at the inlet of the turbine (8); Tref = 273 K; GAMMA is equal to 1.4, co is an actual maximum rotational speed of the fan rotor (9), GR is a reduction gear ratio of the mechanism (19) for stepping down from the fan shaft (20) to the shaft (11).
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
44.
METHOD FOR REWORKING OR REPAIRING A BLADE COMPRISING A LEADING EDGE
The invention relates to a method for manufacturing a blade, the method comprising: - a shaping step, during which a blade structure (10) is formed; and - a positioning step, during which a leading edge (20) is positioned on the front edge (136) of the blade structure (10). The leading edge (200) is attached to the front edge (136) by means of a heat-sensitive adhesive material.
B29C 65/48 - Joining of preformed partsApparatus therefor using adhesives
B29C 65/76 - Making non-permanent or releasable joints
B29C 73/04 - Repairing of articles made from plastics or substances in a plastic state, e.g. of articles shaped or produced by using techniques covered by this subclass or subclass using preformed elements
A method for carrying out tomographic analysis of a composite part comprising a matrix and fibers, the method including: acquiring at least one two-dimensional image of the part by a tomographic device, generating a plurality of shapes characteristic of the fibers and of convolution masks having a plurality of copies of the characteristic shape, computing, for each of the characteristic shapes, a product of convolution of the two-dimensional image with the corresponding convolution mask, and obtaining a convolution image, detecting a position corresponding to an overall maximum in all the obtained convolution images and attributing a fiber center to the position, marking as processed a region of the image placed around the fiber center, and iterating the steps of computing, detecting and marking on the unprocessed regions of the image.
The invention relates to a sealing system (21) for a turbomachine bearing enclosure (E1), comprising a stator (22) carrying a rotor (23), this system (21) comprising a seal (25) carried by the stator (22) and surrounding the rotor (23), an external peripheral groove (26) carried by the stator (22) in the vicinity of the seal (25) for collecting oil (H) which exits the enclosure (E1) running along the rotor (23) through the seal (25) and which is centrifuged by the rotor (23) towards this groove (26), and a drain (36) at the bottom for discharging the oil (H) from the groove (26). This groove (26) is open towards the inside and is delimited by a base (33) extended by two side walls (34, 35), the side wall (35) opposite the seal (25) is extended by a skirt (38) oriented obliquely towards the base (33) to delimit a gutter (43) opening towards the base (33).
A tool (50) and method for positioning and shaping an adhesive film (28), the tool comprising: - a body (52) comprising a first portion (60) which is made of porous material and which comprises an outer surface (65) of predetermined shape on which an adhesive film (28) is intended to be positioned, and - a support (54) connected to the body (52) and comprising at least one port (56) which is connected to the first portion (60) by at least one internal channel (58, 59), the at least one port (56) being configured to be connected to a vacuum source such that the adhesive film (28) is held in shape and in position on the surface (65) by a suction effect through the porous material of the first portion (60).
The invention relates to an assembly of a turbine engine having a longitudinal axis (X), for an aircraft, which comprises: - a variable-pitch stator vane (2) including a blade and a pivot (5) extending radially from the blade (4); and - a heating element (26) comprising a first portion (26a) mounted in the blade (4) and a second portion (26b) connected to an electrical connection device (29). According to the invention, the pivot comprises a main bore (33), a drilled hole (34) passing through the wall of the pivot so as to open into the main bore (33) and onto an outer surface of the pivot (5), and a radial groove (35) provided in the wall of the pivot which opens onto the outer surface and into an outlet of the drilled hole (34), the second portion (26b) forming an extension of the first portion (26a) and extending into the main bore (33), the drilled hole (34) and then the radial groove (35) towards the electrical connection device.
The invention relates to a propulsive fan (100) comprising two rows of blades (110, 120) and a second row of blades (120) arranged downstream of the first row of blades (110). The leading edge (BA) of one blade (120) of the second row and the trailing edge (BF) of one blade (110) of the first row are at least partially serrated. The radial position of each tooth peak in the leading edge (BA) of the blade (120) of the second row is substantially identical to that of a trough in the trailing edge (BF) of the blade (110) of the first row, and/or the radial position of each trough in the leading edge (BA) of the blade (120) of the second row is substantially identical to that of a corresponding tooth peak in the trailing edge (BF) of the blade (110) of the first row.
The present invention relates to a method for manufacturing an electronic power module, comprising a step of applying a protective layer to a face of a substrate opposite the face carrying electronic power components, arranging the substrate in a mould so that the protective layer is applied to the bottom of the mould, and then injecting resin.
The invention relates to a titanium-based alloy powder which comprises, in percentages by weight, 32.0 to 33.5% aluminium, 4.50 to 5.10% niobium, 2.40 to 2.70% chromium, 0 to 0.1% iron, 0 to 0.025% silicon, 0 to 100 ppm carbon, 0 to 100 ppm nitrogen, 0 to 1000 ppm dioxygen, 0 to 50 ppm dihydrogen and 0 to 500 ppm unavoidable impurities, the balance being titanium, and which has a D10 particle size of between 3 and 10 μm, a D90 particle size of between 20 and 40 μm and a D50 particle size of between 10 and 25 μm, the D10, D50 and D90 particle size values having been measured by laser diffraction in accordance with standard ISO 13322-2. The invention also relates to a method for manufacturing a part using this powder and to a part thus obtained.
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 1/107 - Metallic powder containing lubricating or binding agentsMetallic powder containing organic material containing organic material comprising solvents, e.g. for slip casting
The invention relates to an assembly of a turbine engine having a longitudinal axis (X), for an aircraft, the turbine engine comprising: - a variable-pitch stator vane (2) including a blade (4) and a pivot (5) which extends radially from the blade (4); - a pitch-change system (3) for changing the pitch of the stator vane; and - a heating element (26) comprising a first portion (26a), which is mounted in the blade (4), and a second portion (26b), which is connected to an electrical power supply source (30) of an electrical connection device (29). According to the invention, the electrical connection device (29) comprises a connection housing (31) which connects the second portion to a power supply harness (32a, 32b) which is coupled to the electrical power supply source, and the pivot comprises a main bore (33) in which the connection housing (31), constrained to rotate with the pivot, and the second portion (26b) are housed.
abcc), in the radial direction, from the inclined-tooth point to a second adjacent gullet that is farther from the central axis (X) than the inclined-tooth point.
G06T 3/14 - Transformations for image registration, e.g. adjusting or mapping for alignment of images
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]
The present invention relates to a method for manufacturing (S) a blade (7), comprising the following steps: - S1: three-dimensional weaving of a fibrous reinforcement (20) so as to form a first skin (21) and a second skin (21) delimiting a cavity (23) therebetween; - S2: inserting a core (40) into the cavity (23); - S3: inserting a first plug (31) opening into the cavity (23) so as to come into contact with the core (40); - S4: placing the fibrous reinforcement (20) comprising the core (40) and the first plug (31) into a mould and injecting a matrix; - S5: dissolving the first plug (31) and the core (40), so as to form a through-passage (33) and a housing (41); and - S6: inserting a second plug (32) into the through-passage (33).
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 33/52 - Moulds or coresDetails thereof or accessories therefor with means for, or specially constructed to facilitate, the removal of articles, e.g. of undercut articles soluble or fusible
B29C 45/44 - Removing or ejecting moulded articles for undercut articles
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
The present invention relates to a blade (7) comprising a first skin (21) and a second skin (21) each having an internal face (22) delimiting therebetween an internal recess (23), the blade (7) further comprising a set of stiffeners (30) extending through the internal recess (23), being attached to the internal face (22) of the first skin (21) and of the second skin (21) so as to connect them, the internal recess (23) having a lower boundary (26) adjacent to a root (8) of the blade (7), a leading edge boundary (27) and a trailing edge boundary (28), all or part of the stiffeners (30) extending in the internal recess (23) from a lower boundary (26) to a leading edge boundary (27) to a trailing edge boundary (28).
F01D 5/18 - Hollow bladesHeating, heat-insulating, or cooling means on blades
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
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
B64C 11/06 - Blade mountings for variable-pitch blades
A blade made of a composite material for an aircraft turbomachine fan, the blade comprising means for measuring internal deformations of the blade and means for remotely storing and transmitting signals for measuring the deformation of the blade, which means are connected to the measuring means, the measuring means and the remote storage and transmission means being located in the composite material. The means for measuring can be configured to supply the means for remotely storing and transmitting.
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
B29C 37/00 - Component parts, details, accessories or auxiliary operations, not covered by group or
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 70/68 - Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts by incorporating or moulding on preformed parts, e.g. inserts or layers
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
A turbomachine-for aircraft has a primary annular flow path and a cold flow channel extending around the primary annular flow path. The turbomachine includes an impeller, a compressor, a combustion chamber, a first turbine having a first turbine rotor, a second turbine having a second turbine rotor, a first connecting shaft and a second connecting shaft, an inter-turbine stator arranged between the first turbine and the second turbine, and a first cooling circuit having successively: a first cooling inlet located between the impeller and the compressor, a first passage extending into the inter-turbine stator, and a first cooling outlet extending into the cold flow channel.
ANTI-ICING SYSTEM FOR AN AIR DUCT FOR A COMPRESSOR FOR AN AIRCRAFT TURBINE ENGINE AND ANTI-ICING METHOD FOR AN AIR DUCT FOR A COMPRESSOR FOR AN AIRCRAFT TURBINE ENGINE
The invention relates to an anti-icing system for an air duct of a compressor in an aircraft turbine engine, the system comprising: - an outer wall (30) externally delimiting the air duct, comprising an air reinjection slot (40) and an air bleed slot (50), wherein the air reinjection slot (40) is upstream of the air bleed slot (50); - an anti-icing device, the anti-icing device comprising • a duct (60) connecting the air bleed slot (50) to the air reinjection slot (40) so as to direct the air drawn in through the air bleed slot (50) to the air reinjection slot (40); • a valve (70) positioned in the duct (60) in order to interrupt the airflow inside the duct.
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
60.
METHOD FOR CHARACTERIZING THE OPTICAL FOCAL POINT OF AN X-RAY SOURCE
The invention relates to a method for characterizing the optical focal point of a source (11) of an X-ray beam (13) in the direction of a detector (12), comprising the following steps: - evaluating a blur of a radiograph of a standard object (14); and - determining a dimension of the optical focal point; evaluating the blur of the radiograph comprising: - determining a grayscale profile of a segment of interest of the radiograph, - determining, from the second derivative of the grayscale profile, points of interest of the grayscale profile; and - determining a deviation between the points of intersection between a straight line passing through the points of interest and an upper and lower level of the grayscale profile, respectively.
G01N 23/04 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by transmitting the radiation through the material and forming images of the material
A61B 6/58 - Testing, adjusting or calibrating thereof
61.
METHOD FOR MANUFACTURING A ROTATIONALLY SYMMETRICAL PART MADE OF COMPOSITE MATERIAL WITH LOCALLY OPTIMISED PROPERTIES
The invention relates to a method for manufacturing a rotationally symmetrical part (10) made of composite material having variable thickness for a gas turbine, the method comprising: producing a fibrous texture (140) in the form of a strip by means of three-dimensional or multi-layer weaving; winding the fibrous texture (140) in a plurality of overlapping layers (141, 142, 143, 144) onto a mandrel (200) having a profile corresponding to that of the casing to be manufactured, so as to obtain a fibrous preform having a shape corresponding to that of the casing to be manufactured; densifying the fibrous preform using a matrix. During the step of winding the fibrous texture (140) onto the mandrel, a textile strip (150) is placed at least between one or more adjacent turns of the fibrous texture, wherein the textile strip (150) has a structure different from a three-dimensional weave and a width smaller than the width of the fibrous texture (140) in an axial direction (DA).
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 70/32 - 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
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F01D 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
The invention relates to a device for controlling the torque output by an asynchronous electric machine, comprising: - a rotary slip ring collector system suitable for being connected to a rotor of an asynchronous electric machine and configured to collect or power at least one winding of the rotor of the asynchronous electric machine so as to vary the impedance of the rotor and thus a torque output by the asynchronous electric machine; - a variable resistor connected to the rotary slip ring collector system so as to control the rotary slip ring collector system; - a processing unit configured to, according to a frequency of a supply voltage of the asynchronous machine, a desired torque for the asynchronous electric machine and a corresponding rotor resistance, calculate a value for the variable resistor and thus control the supply voltage of the collector so that the asynchronous electric machine outputs the desired torque.
H02P 1/34 - Arrangements for starting electric motors or dynamo-electric converters for starting dynamo-electric motors or dynamo-electric converters for starting an individual polyphase induction motor by progressive reduction of impedance in secondary circuit
H02P 25/26 - Variable impedance in stator or rotor circuit with arrangements for controlling secondary impedance
The present invention relates to a fuel control system (4) comprising: a fuel source; a supply conduit; a main circuit comprising: a first centrifugal pump; a connecting conduit; a second centrifugal pump; a discharge conduit; and a secondary circuit.
A connecting rod (10), in particular for a turbomachine rotor, extending along an axial extension axis (X-X'), comprising a connecting-rod body (11) comprising a tapped thread and a connecting-rod head (12) constrained to rotate with the connecting-rod body (11) and comprising a connecting shank (13) comprising an external thread (13a) cooperating with the tapped thread of the connecting-rod body (11), the connecting rod (10) further comprising a locking system (20) for locking the relative rotation between the connecting-rod head (12) and the connecting-rod body (11). The locking system (20) further comprises a self-locking nut (23) mounted by screwing onto the connecting shank (13) of the connecting-rod head (12) and configured to block the relative rotation between the connecting-rod body (11) and the connecting-rod head (12), the self-locking nut (23) comprising a deformable portion configured to be plastically deformable in the radial direction under the action of an external stress on its outer surface between an initial configuration in which the deformable portion has a circular cross-section and a deformed configuration in which the deformable portion has a non-circular cross-section and clamps the external thread (13a) of the connecting shank (13).
The invention relates to a tool (20) for machining an abradable annular layer (10b) of a fan (1) casing (3) of an aircraft turbine engine, including: - a central centring body (30) configured to be attached to a fan disk (2) of the turbine engine; - a machining tool (50) connected to the central body (30) and including at least one machining means (51) configured to come into contact with the abradable layer; and - a first motor (52) for rotating the machining means (51), characterised in that the central body (30) comprises or carries a copying cam (70) which comprises a cam profile (72) which depends on a desired surface profile for the abradable layer (10b) after machining.
A hybrid turbine engine comprising: - an input shaft (a26) configured to be driven by a low-pressure turbine (102); - an output shaft (a64) for driving an electric machine (104) and another output shaft (a64') for driving another electric machine (104'); - a variable-speed drive (106) for regulating the speed of each output shaft (a64, a64') on the basis of a speed of the input shaft, comprising: • a first planetary gear set (T1) which has, as input, a planet carrier (108) connected to the input shaft, a ring gear (112) and a planet gear (110); as output, a first sun gear (130) connected to the output shaft, • a second planetary gear set (T2) which has, as input, the ring gear (112) and another planet gear (120); as output, a second sun gear (140) connected to the other output shaft, and • a means (MB) for blocking the ring gear or leaving it free.
The invention relates to a method for manufacturing an outer shroud for an intermediate casing of a turbofan engine made of composite material, the method comprising: - producing a fibrous texture (140) in the form of a strip by means of three-dimensional or multi-layer weaving; - winding the fibrous texture (140) in a plurality of overlapping layers (141, 142, 143, 144) onto a mandrel (200) having a profile corresponding to that of the outer shroud to be manufactured, so as to obtain a fibrous preform having a shape corresponding to that of the outer shroud to be manufactured; - densifying the fibrous preform using a matrix. During the step of winding the fibrous texture (140) onto the mandrel, a textile strip (150) is placed at least between one or more adjacent turns of the fibrous texture, wherein the textile strip (150) has a structure different from a three-dimensional weave and a width smaller than the width of the fibrous texture (140) in an axial direction (DA).
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
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
68.
INTER-COMPRESSOR CASING FOR A TURBINE ENGINE WITH A HYBRID VBV DOOR AND AN OPTIMISED EXTERIOR PROFILE
The invention relates to an inter-compressor assembly for a turbine engine, wherein a VBV door (62) is mounted so as to pivot about a pivot axis (64) so as to be movable between a closed position in which the VBV door closes an opening (60) and an open position in which the VBV door allows air to flow through an upstream portion (60A) and a downstream portion (60B) of the opening. The VBV door (62) is profiled with a downstream surface (76) having a convex axial section extending from one of its ends to the other of its ends, and an angle between the downstream surface (76) and an inner surface (72) of the VBV door is between 120 and 140 degrees. The risk of aerodynamically disturbed air from the VBV door in the open position deviating from its path and negatively affecting the operation of another member located downstream, such as a turbine engine compressor, instead of being drawn into the downstream portion (60B) of the opening, is thus reduced.
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
F02C 3/13 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having variable working fluid interconnections between turbines or compressors or stages of different rotors
69.
TURBINE ENGINE MONITORING METHOD AND CORRESPONDING TURBINE ENGINE
The invention relates to a method for monitoring a turbine engine, which method comprises the following steps: - performing a differential measurement (ΔT) between a first temperature (T1) upstream of a flow of a heat transfer fluid and a second temperature (T2) downstream of the flow of heat transfer fluid, the flow of heat transfer fluid passing through an air-fluid heat exchange member (10) of the turbine engine which is exposed to an air flow capable of transporting water in the solid state or in the liquid state, - comparing the differential measurement (ΔT) with a reference value (ΔTref), - and, on the basis of this comparison, determining the presence of water in the solid state taken into the air-fluid heat exchange member (10) or of frost affecting the air-fluid heat exchange member (10).
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for
F01D 21/10 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to unwanted deposits on blades, in working-fluid conduits, or the like
F01D 25/02 - De-icing means for engines having icing phenomena
F02C 7/14 - Cooling of plants of fluids in the plant
F01D 21/12 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to temperature
F02C 7/05 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
70.
RTM INJECTION METHOD AND MOULD USING SYMMETRIC ANTI-PINCHING SECTORS
An injection mold for the manufacture of an axisymmetric part of composite material comprising a mandrel supporting a fibrous preform and comprising an annular wall, and a plurality of counter-mold angular sectors assembled to the mandrel and intended to close the mold and to compact the fibrous preform wound on the mandrel. Each angular sector comprises an annular base intended to come into contact with the fibrous preform. The plurality of angular sectors includes a first series of angular sectors and a second series of angular sectors, first and second lateral edges of the first series of sectors each including a protruding lower portion, the first and the second lateral edges of the second series of sectors each including a recessed lower portion.
B29C 70/54 - Component parts, details or accessoriesAuxiliary operations
B29C 70/48 - Shaping or impregnating by compression for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM]
The invention relates to a power supply architecture (1) for a hybrid-electric propulsion aircraft (0), comprising a gas turbine (2), a propulsion power supply network (5), and a non-propulsion power supply network (6); the propulsion power supply network (5) comprising a first AC generator (9), at least one first AC/DC converter (10), and at least one first battery (11) for supplying power to the propulsion equipment (7) of the aircraft; the non-propulsion power supply network (6) comprising a second AC generator (20), at least one second AC/DC converter (21), and at least one second battery (22) for supplying power to the non-propulsion equipment (8) of the aircraft.
B60L 50/12 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by engine-driven generators, e.g. generators driven by combustion engines using AC generators and DC motors
B60L 50/61 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells using power supplied by batteries by batteries charged by engine-driven generators, e.g. series hybrid electric vehicles
72.
VANE FOR AN AIRCRAFT TURBINE ENGINE, ASSOCIATED CORE AND MANUFACTURING METHOD
The invention relates to a vane comprising: - a blade extending along an axis of extension (Y), - at least one internal circulation cavity (20) for a flow of cooling air (F3) to circulate, delimited by an internal wall (18) of the blade, - at least one first air outlet opening (22) provided in at least one of the pressure-side and suction-side faces, - at least one internal channel (30) connected to the first outlet opening (22) and extending along an axis transverse to the axis of extension (Y), characterized in that the at least one internal channel (30) is also connected to at least one second air outlet opening (32) provided in the internal wall (18) and opening into the internal cavity (30), the second outlet opening (32) having a back wall (34) connected to the internal channel (30) and inclined in relation to the axis of the internal channel (30).
The present disclosure relates to a stator assembly of a turbine engine in which an air flow circulates, the assembly comprising: - an inner shroud extending along a longitudinal axis X-X, - an outer shroud radially outside the inner shroud, and - a stator vane extending radially from the inner shroud to the outer shroud, the stator vane comprising a leading edge and a trailing edge downstream of the leading edge in the direction of circulation of the air flow, and a suction-side wall and a pressure-side wall extending respectively from the leading edge to the trailing edge. The stator vane further comprises a through-slot formed at least partially in the suction-side wall and extending over the entire height of the stator vane.
A turbomachine includes a low-pressure shaft having an internal passage for a first air stream and a first segment. A second segment includes a first frustoconical portion, and an intermediate lubrication chamber wherein at least one bearing is arranged. The intermediate lubrication chamber is arranged in an intermediate casing. An annular air chamber is configured to be supplied with air from a low-pressure compressor and is at least partly axially delimited by the first frustoconical portion. A radial separating wall is mounted around the first segment, and the first frustoconical portion further includes an opening to enable the first air stream to pass from the air chamber to the internal passage.
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
F01D 25/20 - Lubricating arrangements using lubrication pumps
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 7/14 - Cooling of plants of fluids in the plant
F02C 7/16 - Cooling of plants characterised by cooling medium
F02C 9/18 - Control of working fluid flow by bleeding, by-passing or acting on variable working fluid interconnections between turbines or compressors or their stages
F02C 9/52 - Control of fuel supply conjointly with another control of the plant with control of working fluid flow by bleeding or by-passing the working fluid
75.
TOOL FOR BONDING AN ANTI-WEAR STRIP TO A TURBOMACHINE BLADE
A tool (40) for bonding at least one anti-wear strip (22) to a turbomachine blade (10), comprising: - a support (50), - at least one first fastening member (60) which is fastened to the support (50) and is able to cooperate with the root (14) of the blade (10) in order to immobilize it with respect to the support (50), and - at least one second member (70) for holding said at least one anti-wear strip (22) on the blade (10) and for heating this anti-wear strip (22) for the bonding thereof, this second member (70) being fastened to the support (50) and comprising, on the one hand, at least one expandable bag (72), and, on the other hand, at least one heating element (80) able to heat said at least one anti-wear strip (22) for the bonding thereof.
Disclosed is an assembly for a turbomachine comprising: ⋅—a second connection element (8) having a third end (81) which is connected to a turbine shaft (2) and is mounted on a second bearing (5), and a fourth end (82) connected to a mechanical transmission device (6), the second connection element (8) having a second radial compliance (SR2); and ⋅—a third connection element (9) having a fifth end (91) fixedly mounted on a casing (1), and a sixth end (92) connected to the mechanical transmission device (6), the third connection element (9) having a third radial compliance (SR3), wherein a ratio of the first radial compliance (SR1) to the third radial compliance (SR3) is strictly lower than 10% and/or a ratio of the first radial compliance (SR1) to the second radial compliance (SR2) is strictly lower than 4%.
The invention relates to a secondary flow stator vane (1) for a turbomachine, comprising skins (2, 3) having end portions (21, 31) delimiting an opening (5) and a cavity (4) therebetween, which skins diverge from one another in a direction (Y). A filler material (41) is located in the cavity (4) at a distance from the opening (5). A reinforcement (6), the density of which is higher than the density of the material (41), comprises a first reinforcement portion (61) located in the cavity (4) against the material (41) and a second reinforcement portion (62) for closing the opening (5) between the portions (21, 31) to form a single piece.
This connecting assembly comprises: a connecting member (4) comprising a rod (12), a bearing and a head (14b) provided with two opposing faces traversed by a bore forming a housing for the bearing, a female clevis comprising two arms arranged to either side of the head and each provided with a hole, and a pin extending through the arms and the bearing, a washer (10) provided with a body (34) arranged axially between one arm of the female clevis and one of the two opposing faces of the head (14b), at least one tab (42a, 42b) extending from the body (34) and axially beyond said face of the head, and at least one additional tab (44a) extending from the body (34) and being in contact with said face of the head (14b).
A gas turbine shroud assembly comprises a turbine shroud comprising shroud sectors (28) made of ceramic matrix composite material, a casing (22) made of metal alloy and a sectorized cooling device (52) comprising a plenum chamber (52B) delimited by an air diffuser (54) having holes for diffusing jets of cooling air over a radially outer surface (30B) of the turbine shroud. The assembly further comprises two metal alloy bushings (60) housed in two cavities (62) present at the circumferential ends of the air diffuser (54). Each of the bushings is in the form of a stepped cylinder with a base (60A), a lower face (601) of which bears on the radially outer surface (30B) of the turbine shroud, and a stud surmounting the base (60B). The stud is intended to receive a spring (58) compressed between an upper face (602) of the base and an upper face (620) of the cavity.
An assembly for an aircraft turbine engine propeller (10), the assembly comprising a variable-pitch vane (14), a system (76) for controlling the pitch of the vane, which comprises a bushing (80) engaged in a metal body (102) forming the root of the vane, and a tie rod (120) which passes through the bushing (80) and the metal body (102) in order to link the vane to the control system (76) and prestress the metal body (102).
One aspect of the invention relates to a method for determining at least one microstructural property of a metal alloy on the surface of a part manufactured in said metal alloy.
G01N 23/223 - Investigating or analysing materials by the use of wave or particle radiation, e.g. X-rays or neutrons, not covered by groups , or by measuring secondary emission from the material by irradiating the sample with X-rays or gamma-rays and by measuring X-ray fluorescence
The invention relates to a turbine shroud assembly (2) comprising a plurality of shroud segments (10) made of ceramic matrix composite material, which form a turbine shroud (4) and define an axial direction (DA), a radial direction (DR), and a circumferential direction (DC), as well as a plurality of inter-segment junctions (100) in the circumferential direction (DC), wherein the turbine shroud assembly (2) further comprises a shroud support structure (6) having an upstream annular flange (20) and a downstream radial flange (64), between which each shroud segment (10) is axially and radially held, and wherein the upstream annular flange (20) is formed as a single annular part. The upstream annular flange (20) comprises, opposite each shroud segment (10), a circumferential expansion slot (80).
The invention relates to a turbine ring assembly provided with a sectorized CMC turbine ring and with a ring support structure (6), each sector (10) comprising a base (12) from which an upstream attachment tab (16) and a downstream attachment tab (14), which are axially spaced apart, extend radially outwards, the support structure (6) comprising an upstream radial flange (62) and a downstream radial flange (64) between which the attachment tabs (14, 16) are held. Opposite each inter-sector junction (100) between two adjacent ring sectors (10) in the circumferential direction (De), the downstream radial flange (64) comprises a fit-adjustment portion (70) machined in the thickness of the downstream radial flange (64), from an upstream face (642) of the downstream radial flange (64), and extending radially from the inner radial end (643) of the downstream radial flange (64), the thickness of the downstream radial flange (64) being measured in the axial direction (DA), and the fit-adjustment portion (70) of the downstream radial flange (64) being axially distant from the inter-sector junction (100) opposite which it is positioned.
A turbine engine module, this module having a longitudinal axis (X) and comprising: - a propeller (3) comprising a plurality of variable-pitch vanes (30), - a speed reducer (34), - a system (50) for changing the pitch of the vanes (30), - a fluid transfer device (94) configured to supply fluid to the hydraulic actuator (52) of the pitch-changing system (50), this transfer device (94) comprising a stator part (96) engaged in a rotor part (97), and - a member (99) for mechanically connecting the stator part (96) of the transfer device (94) to the planet carrier (38) of the speed reducer (34), and fluidically connecting a fluid inlet port (98a) of the transfer device (94) to fluid pipes (95) configured to be connected to a fluid tank (91).
A rolling bearing (401) for an aircraft turbine engine which comprises an inner ring and an outer ring (403) which are coaxial about a rotation axis, and rolling elements and a cage separating the rolling elements, located between the inner ring and the outer ring (403). The outer ring (403) comprises at least one drainage channel (405) which extends from an inner face (403ba) of the outer ring (403) to a downstream face (403bb) of the outer ring (403) and which has no edge and no right angle between different portions of its length.
F16C 19/16 - Bearings with rolling contact, for exclusively rotary movement with bearing balls essentially of the same size in one or more circular rows for both radial and axial load with a single row of balls
B29C 64/00 - Additive manufacturing, i.e. manufacturing of three-dimensional [3D] objects by additive deposition, additive agglomeration or additive layering, e.g. by 3D printing, stereolithography or selective laser sintering
B33Y 80/00 - Products made by additive manufacturing
F16C 27/04 - Ball or roller bearings, e.g. with resilient rolling bodies
The invention relates to an air jet cooling device (100) for a casing (48) of a turbine, in particular a casing of a low-pressure turbine (28), comprising at least one air supply housing (120), at least one cooling rail (110) intended to be disposed around the casing to be cooled, the or each cooling rail comprising two cooling tubes (112) that are disposed on either side of said housing, the supply housing (120) comprising two lateral walls (125, 126), each lateral wall comprising air outlet orifices (127), each orifice having a main axis and being configured and dimensioned to receive an associated one of the cooling tubes, the cooling device comprising a fastening system intended to fasten the cooling tubes to said casing, characterized in that each cooling tube is secured to said fastening system and movable in translation along the main axis (P) in the associated orifice.
A device for centering and rotationally guiding a turbomachine shaft comprises blocking means (54A, 54B; 130) which are movable between: a standby position, toward which the blocking means are urged by the pressure of a lubricating liquid supplying a film compression damping cavity, and in which the blocking means are spaced apart from at least one of a bearing support (34) and an outer ring (38) so as to leave these free to move radially relative to one another; and a blocking position, toward which the blocking means are urged by elastic return means (90A, 90B; 156; 180), and in which the blocking means are interposed with contact between an outer annular surface (48) of the outer ring (38) and an inner annular surface (50) of the bearing support (34) so as to prevent these from moving radially relative to one other.
F16C 19/24 - Bearings with rolling contact, for exclusively rotary movement with bearing rollers essentially of the same size in one or more circular rows, e.g. needle bearings for radial load mainly
F16C 27/04 - Ball or roller bearings, e.g. with resilient rolling bodies
The present invention relates to an end piece (1') for applying a product to a screw, this end piece (1') comprising a proximal end (11) configured to be able to be fastened to an apparatus for dispensing this product and an opposite, distal end (12), this end piece (1') having a longitudinal axis (X-X'). This end piece is characterized in that it comprises at least one chamber (2) for distributing the product, having the form of a cavity that opens out at the distal end (12) of the end piece and is configured to be able to receive the shank of a screw, a first chamber (3) for introducing the product, opening out at the proximal end (11) of the end piece, and a plurality of channels (4') for distributing the product, connecting the first introduction chamber (3) to the distribution chamber (2), each distribution channel (4') opening out into the distribution chamber (2) via at least one distribution orifice (41') and these distribution orifices (41') being disposed at the periphery of said distribution chamber (2).
B05C 17/005 - Hand tools or apparatus using hand-held tools, for applying liquids or other fluent materials to, for spreading applied liquids or other fluent materials on, or for partially removing applied liquids or other fluent materials from, surfaces for discharging material through an outlet orifice by pressure
90.
FLUID TRANSFER DEVICE WITH HYDRAULIC AND MECHANICAL CONNECTION MEANS
A fluid transfer device for a turbomachine with longitudinal axis X, which includes a stator portion intended to be connected to a stator equipment of the turbomachine and a rotor portion engaged in the stator portion. The stator equipment including at least one conduit fluidly connected to at least one pipeline of the stator portion, the stator portion and the stator equipment including an attachment interface intended to releasably receive attachment members, the attachment interface and the attachment members being configured so as to make a sealed coincidence of the pipelines and conduit, and the transfer device comprising at least one passage allowing the access of an external tool upstream of the transfer device for accessing the attachment members and passing through the transfer device in either side.
The invention relates to a turbine engine subassembly comprising: - an inter-compressor casing (17) and a bearing support (18) mounted on the inter-compressor casing (17); - a high-pressure body (11) and a low-pressure body (9); - an electric machine (M) comprising an electric stator (S) mounted on the bearing support (18) and an electric rotor (R) surrounding the electric stator (S) while being mounted on the high-pressure body (11); - a downstream bearing (22) mounted on the inter-compressor casing (17) for rotationally guiding the high-pressure body (11); - an upstream bearing (26) mounted on the bearing support (18) for rotationally guiding the low-pressure body (9); - an additional bearing (39) mounted on the bearing support (18), and being located longitudinally between the upstream bearing (26) and the downstream bearing (22) in order to guide the rotation of the low-pressure body (9).
The invention relates to a ventilation circuit (80) suitable for being mounted in a turbine engine (1) with a first flow path (V1) with a combustion chamber (5), a second flow path (V2) and an item of equipment (90). The ventilation circuit (80) comprises a first duct (10) which connects the equipment (90) downstream to the atmosphere, a second duct (20) which connects the equipment (90) upstream to the first flow path (V1) upstream of the combustion chamber (5) or to the second flow path (V2) and a drive mechanism (30), and is such that: - in a first configuration wherein the turbine engine (1) is operational, the air circulates in the downstream direction in the second duct (20) and then in the first duct (10); and - in a second configuration wherein the turbine engine (1) is stopped, the drive mechanism (30) operates and circulates the air in the upstream direction in the first duct (10) and then in the second duct (20).
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (CNRS) (France)
UNIVERSITE DE BORDEAUX (France)
INSTITUT POLYTECHNIQUE DE BORDEAUX (France)
Inventor
Cavarroc, Marjorie Christine
Poulon, Angélique Nadine Jeanne
Joulia, Aurélien
Michau, Dominique Paul Abel
Abstract
The invention relates in particular to a method for coating a silicon-containing ceramic matrix composite substrate (1) with at least one coating (4) that forms an environmental barrier, which comprises the following steps: E1: forming a silicon-based bonding undercoat (2) on top of said ceramic matrix composite substrate (1); E2: forming a coating (4) that forms an environmental barrier on top of said silicon-based bonding undercoat (2), characterized in that it comprises an additional step Eadd of forming at least one layer of TaN in the hexagonal phase (5) on top of or under said silicon-based bonding undercoat (2).
The invention relates to a device (10) for transferring a plurality of fluid paths, which comprises: a central portion (20); a peripheral portion (22) rotating about the central portion; and transfer chambers (24A-24C) between the central and peripheral portions. The central portion defines first fluid paths (FP1A-FP1C) connecting fluid inlets (14A, 14C) to the transfer chambers. The peripheral portion (22) defines second fluid paths (FP2A-FP2C) connecting fluid outlets (16A-16C) to the transfer chambers. The central portion comprises, in order to define the first fluid paths (FP1A-FP1C): fluid channels (36A-36C) – one for each first fluid path – extending concentrically and connected, from a first axial side (S1), to the fluid inlets; and, for each of the fluid channels, at least one fluid tapping (38) connected to the fluid channel in question, on a second axial side (S2), and opening into the corresponding transfer chamber (24A-24C).
The invention relates to a device (10) for transferring a plurality of channels of fluid, which device comprises: a central portion (20); a peripheral portion (22) rotating about the central portion; and transfer chambers (24A-24C) between the central and peripheral portions. The central portion defines first fluid paths (FP1A-FP1C) connecting fluid inlets (14A, 14C) to the transfer chambers. The peripheral portion (22) defines second fluid paths (FP2A-FP2C) connecting fluid outlets (16A-16C) to the transfer chambers. The peripheral portion comprises a receiving bushing (DR) and at least two rings (70A, 70B) mounted tightly in the receiving bushing (DR) and each defining, together with the latter, at least some of the second fluid paths (FP2A-FP2C).
F16L 39/04 - Joints or fittings for double-walled or multi-channel pipes or pipe assemblies allowing adjustment or movement
F16L 39/06 - Joints or fittings for double-walled or multi-channel pipes or pipe assemblies of the multiline swivel type, e.g. comprising a plurality of axially mounted modules
F01D 9/06 - Fluid supply conduits to nozzles or the like
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
96.
MULTI-WAY FLUID-TRANSFER DEVICE WITH PERIPHERAL PORTION PROVIDED WITH A HOOP, A SLEEVE AND RINGS JOINTLY DEFINING FLUID PATHS
The invention relates to a device (10) for transferring a plurality of fluid pathways, which device comprises: a central portion (20); a peripheral portion (22) rotating about the central portion; and transfer chambers (24A-24C) between the central portion and the peripheral portion. The central portion defines first fluid paths (FP1A-FP1C) connecting fluid inlets (14A, 14C) to the transfer chambers. The peripheral portion (22) comprises a transfer tube (DT), and a hoop (90) tightly mounted around the transfer tube (DT). The transfer tube (DT) and the hoop (90) jointly define second fluid paths (FP2A-FP2C) connecting fluid outlets (16A-16C) to the transfer chambers; - the transfer tube (DT) is formed by a sleeve (50) delimiting a bore (52) extending along the axis (8), and at least two rings (70A, 70B) having respective outer surfaces (72) mounted tightly in the bore (52) and respective inner surfaces (74) forming corresponding portions of the inner surface (22A) of the peripheral portion; - the sleeve (50) defines the tube interface surface (54) and comprises, in order to contribute to defining each of the fluid-transfer paths, at least one corresponding fluid passage (64) for each fluid-transfer path, each fluid passage (64) having an inner end that opens through the bore (52), and an outer end that opens through the tube interface surface (54); - each ring (70A, 70B) comprises fluid passages (76A, 76B), each having an inner end (78) that opens through the inner surface (74) of the ring into one of the corresponding transfer chambers (24A, 24C) and an outer end (84) that opens through the outer surface (72) of the ring so as to form another portion of one of the corresponding fluid-transfer paths.
F16H 57/04 - Features relating to lubrication or cooling
F16L 39/06 - Joints or fittings for double-walled or multi-channel pipes or pipe assemblies of the multiline swivel type, e.g. comprising a plurality of axially mounted modules
F16L 27/08 - Adjustable jointsJoints allowing movement allowing adjustment or movement only about the axis of one pipe
The invention relates to a flow guide assembly (6) comprising: - two outer (62) and inner (64) annular shrouds defining therebetween a first flow path (V2) for a flow (F2), wherein the shrouds are each sectioned into outer and inner shroud sectors (620, 640) arranged circumferentially about an axis (A); and - discharge ducts (66) located inside the inner shroud (62) and leading into openings (65) formed in this inner shroud (62), wherein these ducts (66) define internal discharge flow passages (F3) intended to be injected into the first flow path (V2) through these openings (65), and wherein each of the ducts (66) is formed as a single part of composite material with one of the sectors of the inner shroud (640) that comprises the opening (65) into which this duct (66) leads.
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups Air intakes for jet-propulsion plants
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
98.
SYSTEM FOR CONTROLLING A HYBRIDISED TURBINE ENGINE VIA OBSERVER-BASED RECONFIGURATION BY DETECTION OF CONTROL SATURATION
The invention relates to a system for controlling a hybridised turbine engine (2) provided with at least one rotational speed sensor, the system comprising: a. a means (3) for determining a rotational speed setpoint; b. an observation means determining a control disturbance vector; c. a detection means (6) that issues a fault signal when a value of the disturbance vector is greater than a predetermined threshold; and - a means (11) for calculating the torque control commands for the hybridised turbine engine (2) according to the setpoint and the rotational speed measurement, and a fault signal; - a means (12a) for saturating the torque control command intended for the electrical portion (2a) of the hybridised turbine engine according to the fault signal; - a subtracter (13a) that ussues the fault signal in the event of a difference between the torque control command intended for the electrical portion (2a) of the turbine engine before and after saturation.
The invention relates to a component (11) for an aircraft turbomachine, the component (11) comprising: - a body (12) comprising a composite material comprising an organic matrix and fibres embedded in the matrix, and - a fire barrier (13) positioned on at least one surface of the body (12), characterized in that the fire barrier (13) is multilayer and comprises: - a perforated first layer (14) comprising glass fibres, and - a second layer (15) comprising fibres that differ from glass fibres, the first layer (14) being located between the body (12) and the second layer (15).
B32B 3/04 - 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 features of form at particular places, e.g. in edge regions characterised by a layer folded at the edge, e.g. over another layer
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 3/26 - 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 particular shape of the outline of the cross-section of a continuous layerLayered 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 layer with cavities or internal voids
B32B 5/02 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by structural features of a layer comprising fibres or filaments
B32B 5/12 - Layered products characterised by the non-homogeneity or physical structure of a layer characterised by structural features of a layer comprising fibres or filaments characterised by the relative arrangement of fibres or filaments of adjacent layers
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
The invention relates to an one-piece acoustic component (120) comprising a plurality of hollow acoustic elements (121) having a shape that gradually narrows from a base (121a) to an apex (121b), wherein the bases (121a) of the hollow acoustic elements (121) are connected to one another by connecting edges (123), wherein the acoustic component (120) is characterised in that it comprises a plurality of partitions (122) linking the hollow acoustic elements (121) so as to form acoustic cells that extend between the hollow acoustic elements (121), and wherein the apices (121b) of the hollow acoustic elements (121) each comprise an opening (121c) that leads into at least one acoustic cell, such that a single opening (121c) of the hollow acoustic element (121) leads into each acoustic cell.
