Propulsion system (2) for a helicopter (1) comprising a main engine (9), a main rotor (3), a main gearbox (4) including an output mechanically connected to the main rotor (3), a reduction gearbox (13) mechanically coupled between the main engine (9) and a first input of the main gearbox (4), and an assistance device (10).
Propulsion system (2) for a helicopter (1) comprising a main engine (9), a main rotor (3), a main gearbox (4) including an output mechanically connected to the main rotor (3), a reduction gearbox (13) mechanically coupled between the main engine (9) and a first input of the main gearbox (4), and an assistance device (10).
The assistance device (10) comprises a first electric machine mechanically coupled to the reduction gearbox (13) and configured to operate as an electric generator to take off energy produced by the main engine (9), and a second electric machine mechanically coupled to a second input of the main gearbox (4), the second electric machine being supplied with electrical power by the first electric machine and configured to operate as an electric motor to deliver additional mechanical power to the main gearbox (4).
The invention relates to a propulsion system (1, 1) for an aircraft, comprising a rotor (2) and a nacelle failing (3) that extends around said rotor in relation to an axis (X) and includes an upstream portion (10) forming an inlet section (BA) of the nacelle fairing (3) as well as a downstream portion (20), a downstream end (21) of which forms an outlet section (BF) of the nacelle fairing (3); and characterized in that the downstream portion (20) has a radially inner wall (20a) and a radially outer wall (20b), both of which are made of a deformable shape memory material, and in that the downstream end (21) includes pneumatic or hydraulic actuators (23, 23′) extending in different consecutive angular sectors about said axis (X), each actuator being independently actuatable and being configured to deform, in a direction that extends radially in relation to said axis (X) and is centered angularly in relation to its angular sector, under the effect of a predetermined control pressure.
A rotor for a centrifugal breather for an air/oil mixture of a turbomachine, this rotor including a hollow shaft extending along an axis, a pinion for rotating the hollow shaft, this pinion extending around the axis and being formed of a single part and in a first material with at least one first portion of the hollow shaft, and an annular structure extending around the axis and constrained to rotate with the shaft, this structure being produced in a second material, different from the first material, wherein the structure is made integral with the shaft by additive manufacturing of this structure directly on at least one annular surface of the pinion which forms at least one annular support surface for this additive manufacturing.
A method for igniting a continuous combustion engine including an electronic engine control member, a high energy box, a spark plug ignition circuit and a fuel solenoid valve, cooperating with a starter motor, the method being implemented by the electronic engine control member and including precharging the high energy box before an engine starting procedure, activated on an engine starting command, the precharging being controlled by switching on the electronic engine control member, or by putting the engine in idle mode.
An assembly for measuring a torque transmitted between a first member and a second member of a turbomachine includes a pinion having a first annular portion and a second annular portion joining together at a connecting portion carrying a gearing of the pinion, said first and second annular portions extending axially in opposite directions, from the connecting portion. The pinion also includes an axial power shaft with a first area coupled to the first annular portion, and a second area for coupling to the second member, and a device for measuring the torsion between the first and second areas. The device includes a first phonic wheel on the power shaft, a second phonic wheel equipping the second annular portion of the pinion and axially aligned with the first phonic wheel, and an acquisition means axially aligned with and opposite to the first phonic wheel and the second phonic wheel and outputting a signal representative of an angular variation.
G01L 3/08 - Rotary-transmission dynamometers wherein the torque-transmitting element comprises a torsionally-flexible shaft involving optical means for indicating
6.
DOUBLE WALL FOR AIRCRAFT GAS TURBINE COMBUSTION CHAMBER AND METHOD OF PRODUCING SAME
A double wall for an aircraft gas turbine combustion chamber comprising an internal wall which is configured to be in contact with the combustion reaction, and an external wall which is at a distance from the internal wall, comprising a plurality of openings so as to allow the circulation of cooling air streams outside the external wall, which cool the internal wall. The internal wall being free of perforations prevent any circulation of a cooling air stream towards the centre of the combustion chamber. The the internal wall comprises a plurality of members projecting towards the external wall, each projecting member comprising a foot portion and a cylindrical head portion with a circular cross-section, the head portion extending into an opening with a circular cross-section so as to define a calibrated cross-sectional area between the projecting member and the opening, through which area a cooling air stream can flow.
This method for stopping at least one aircraft turbogenerator (1) comprises: —controlling the stopping (E1) of the turbogenerator (1); —passing from the nominal operating speed (Nref) of the power shaft (3, 12) to a first operating speed (N1) lower than the nominal speed (Nref), for a first predetermined duration (t2); —controlling the extinction of the combustion chamber (6) of the gas turbine (2); —maintaining the rotation of the gas turbine at a second speed (N2) for a second predetermined duration (t3), the power shaft (3, 12) being at a second speed (N2) lower than the first operating speed (N1) and, —controlling the stopping of the reversible electric machine (7) in order to no longer drive the power shaft (3, 12), in order to cause a progressive stopping (E9, E10) of the rotation of the gas turbine (2).
A turbomachine includes a propeller, a propeller shaft carrying the propeller, a rotating electric machine, having at least a first configuration in which it is mechanically coupled to the propeller shaft, and a motor oil pump supplying a lubricating circuit of the turbomachine. The rotating electric machine in the first configuration is mechanically coupled to the motor oil pump in such a way that the rotating electric machine additionally drives the motor oil pump when it is supplied with current. Also disclosed is an aircraft including such a turbomachine.
Device (30) for controlling an airflow guiding system (20), comprising at least one actuator (31) configured to translate a control rod (32) between a first and a second end position of a nominal operating range in which at least one vane (21a) of the airflow guiding system (20) can be moved between a first and a second angle, the control rod (32) being connected to the vane (21a) by a control lever (33) comprising a first control rod (36) and a second control rod (37) which are hinged together. The actuator (31) is configured to bring the control rod (32) into a safety position located beyond the second end position of the nominal operating range and to orient the vane (21a) at a safe pitch angle between the first angle and the second angle.
A method for determining an efficiency fault of at least one module of a turboshaft engine of an aircraft. The method comprising a step of determining an estimated real mapping, a step of determining real indicators from the estimated real mapping, a step of determining a plurality of simulated mappings from a simulation of a theoretical model of the turboshaft engine for different efficiency configurations, a step of determining simulated indicators for each simulated mapping, a step of training a mathematical model by coupling the simulated indicators with efficiency configurations, and a step of applying said mathematical model to the real indicators so as to deduce therefrom a real efficiency configuration.
B64F 5/60 - Testing or inspecting aircraft components or systems
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
11.
METHOD FOR DETERMINING AT LEAST ONE POWER LIMIT OF A HYBRID DRIVE TRAIN FOR A TRANSPORT VEHICLE, IN PARTICULAR AN AIRCRAFT
A method for determining at least one minimum power margin of a hybrid drive train for a transport vehicle, each drive element being associated with at least one power source and at least one power consumer. The method including a step of acquiring measurements of power parameters, a step of comparing each measurement with at least one limitation threshold, so as to deduce therefrom at least one gross margin, a step of converting the gross margins into refined margins expressed according to the same common magnitude, a step of transposing into standardised margins at least at one reference point, a step of determining a source power margin and a consumer power margin at said reference point and a step of determining the minimum power margin by selecting the lowest power margin.
B64D 43/00 - Arrangements or adaptations of instruments
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
12.
FREE TURBINE TURBOMACHINE COMPRISING EQUIPMENT DRIVEN BY THE FREE TURBINE
Disclosed is a turbomachine, comprising a gas generator (13) equipped with a first shaft (18), at least one reversible electrical machine (11), a free turbine (12) provided with a second shaft (17) and caused to rotate by a gas flow generated by the gas generator (13), an accessory gear box (14) and at least one accessory (15, 16).
Disclosed is a turbomachine, comprising a gas generator (13) equipped with a first shaft (18), at least one reversible electrical machine (11), a free turbine (12) provided with a second shaft (17) and caused to rotate by a gas flow generated by the gas generator (13), an accessory gear box (14) and at least one accessory (15, 16).
Said at least one electrical machine (11) is mechanically coupled to said second mechanical shaft (17) via the accessory gear box (14) during all phases of operation of the turbomachine (10), the accessory gear box (14) is coupled to the at least one accessory (15, 16) and the turbomachine (10) further comprises a single mechanical coupling means (20) for mechanically coupling said first mechanical shaft (18) to the accessory gear box (14) in a first configuration and mechanically uncoupling said first mechanical shaft (18) from the accessory gear box (14) in a second configuration.
A propulsion unit having a propeller for an aircraft including a nacelle; a propeller mounted in the nacelle so as to be capable of rotating about a longitudinal axis of rotation, the propeller having blades mounted by a root so as to be capable of pivoting between a deployed position, in which they extend radially relative to the axis of rotation, and a folded position, in which they are longitudinally received against the nacelle; drive means that rotate the propeller; indexing means for stopping the propeller in at least one indexed angular position relative to the nacelle; the propulsion unit wherein the indexing means consist of a stepping electric motor including a rotor that is coupled to the propeller.
Disclosed is a turbogenerator, in particular for an electrically-driven rotary wing aircraft, comprising a gas generator equipped with a first shaft, at least one reversible electrical machine, and a free turbine provided with a second shaft and caused to rotate by a gas flow generated by the gas generator. The second shaft is coupled to the at least one electrical machine during all phases of operation of the turbomachine, and the turbomachine further comprises a single mechanical coupling means for coupling the first mechanical shaft to the second mechanical shaft when the electrical machine is operating in motor mode and mechanically uncoupling the first mechanical shaft from the second mechanical shaft when the electrical machine is operating in generator mode.
A turbomachine, particularly for a rotary-wing aircraft, including a gas generator provided with a rotary shaft, a first reversible electric machine, a power turbine rotationally driven by a stream of gas generated by the gas generator, at least one accessory from among an oil pump and a fuel pump, an accessory gearbox comprising a gear train configured to drive said at least one accessory, and a second electric machine.
The second electric machine is reversible, said first electric machine is mechanically coupled to the gas generator, the accessory gearbox and the second electric machine are mechanically coupled to the power turbine, and the turbomachine is devoid of any kinematic coupling between the gear train of the accessory gearbox and the shaft of the gas generator.
A method for protecting coils from excessive heating in an aircraft electrical machine comprising a stator (12) and a rotor (14) configured to be rotationally driven with respect to one another, the stator including a plurality of notches (120) receiving one and the same plurality or otherwise of coils, the method including the following successive steps:
inserting an electrical insulator (16) into the notches or onto the teeth of the stator,
installing the coils (18) in the notches or on the teeth of the stator, casting a phase change material (20) in the notches or on the teeth equipped with the coils, the electrical insulator forming a casting mold.
H02K 15/12 - Impregnating, heating or drying of windings, stators, rotors or machines
H02K 3/34 - Windings characterised by the shape, form or construction of the insulation between conductors or between conductor and core, e.g. slot insulation
H02K 15/02 - Methods or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines of stator or rotor bodies
17.
CIRCULAR MODULAR TRAY FOR THE ADDITIVE MANUFACTURING OF A PART WITH AN AXIS OF REVOLUTION ON A POWDER BED
A circular modular tray for the additive manufacturing of a part with an axis of revolution on a powder bed, characterised in that it consists of an assembly of modules that are concentrically coupled along a contiguous axis in a radial direction, the modules including an annular peripheral module and a circular central module.
B33Y 30/00 - ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING - Details thereof or accessories therefor
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
A gas turbomachine combustion chamber includes bridges extending side by side to connect in one piece a radially inner wall and a radially outer wall towards a free end of the radially inner wall. The bridges, inner wall, and outer wall have an additive layer structure.
Hybrid turbomachine comprising an electric generator, a gas generator equipped with an air inlet and with an exhaust and an acoustic monitoring system comprising a control unit and a plurality of loudspeakers.
Hybrid turbomachine comprising an electric generator, a gas generator equipped with an air inlet and with an exhaust and an acoustic monitoring system comprising a control unit and a plurality of loudspeakers.
At least a first loudspeaker is disposed on the electric generator, and/or at least a second loudspeaker is disposed on the air inlet of the gas generator, and/or at least a third loudspeaker is disposed on the exhaust of the gas generator.
Hybrid turbomachine comprising an electric generator, a gas generator equipped with an air inlet and with an exhaust and an acoustic monitoring system comprising a control unit and a plurality of loudspeakers.
At least a first loudspeaker is disposed on the electric generator, and/or at least a second loudspeaker is disposed on the air inlet of the gas generator, and/or at least a third loudspeaker is disposed on the exhaust of the gas generator.
The control unit of the acoustic monitoring system is mounted on the electric generator and is configured to make an AC-DC electrical conversion of the electromotive force of the electric generator into an adjustable DC voltage intended to be distributed to loads or to energy storage means.
A propulsion unit with a propeller includes a nacelle; a propeller rotatably mounted in the nacelle by means of a hub. The propeller has blades mounted in a blade cuff that is pivotable about a pitch axis relative to the hub. Each blade is pivotable relative to the cuff about a folding axis. A folding device includes an actuator for folding the blades. The folding device includes a control member rotationally affixed to the blade cuff and driven by the actuator. A connecting rod is pivotably mounted, on the one hand, on a root of the associated blade and, on the other hand, on the movable control member.
A method for starting an aircraft engine in which the engine is coupled to a lubrication circuit including an oil pump system, the lubrication circuit being constructed and arranged to circulate oil in the engine, and in which an operating mode of the engine includes a stop mode and a standby mode, the starting method including, during a starting phase, measuring an oil temperature, the measurement being performed by a temperature detection device; depending on the temperature measured, compared to a threshold temperature, and depending on the operating mode of the engine, select a starting oil flow profile to be applied in said engine, the selection being performed by a calculator, and applying the selected starting oil flow profile by the oil pump system, the oil pump system being controlled by the calculator.
F16N 29/02 - Special means in lubricating arrangements or systems providing for the indication or detection of undesired conditions; Use of devices responsive to conditions in lubricating arrangements or systems for influencing the supply of lubricant
A propulsion unit with a propeller is provided, having a nacelle and a propeller rotatably mounted in the nacelle by a hub. The propeller includes blades mounted in a blade cuff pivotable about a pitch axis relative to the hub, each blade being pivotable relative to the cuff about a folding axis. The propulsion unit can further include a folding device that comprises an actuator for folding the blades. The folding device can include a control member rotationally affixed to the blade cuff and driven by the actuator, and a connecting rod pivotably mounted, on the one hand, on a root of the associated blade and, on the other hand, on the movable control member.
B64C 11/06 - Blade mountings for variable-pitch blades
B64C 27/28 - Compound rotorcraft, i.e. aircraft using in flight the features of both aeroplane and rotorcraft with forward-propulsion propellers pivotable to act as lifting rotors
B64C 29/02 - Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis vertical when grounded
Aircraft turbomachine including a centrifugal compressor, a combustion chamber, the combustion chamber being supplied by the compressor via a diffuser and via a straightener, and a heat exchanger, the exchanger including a first circuit, supplied with exhaust gas from the turbomachine, and a second circuit, which are connected by volutes on the one hand to an outlet of the diffuser and on the other hand to an inlet of the straightener, the volutes having reversed winding directions such that their connection ports to the exchanger are independent of one another and are substantially diametrically opposed, and such that the minimum cross section of each duct is situated at a larger cross section of the other duct.
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
F02C 7/08 - Heating air supply before combustion, e.g. by exhaust gases
An aircraft turbomachine having a centrifugal compressor, an annular combustion chamber, an annular casing extending around the chamber and delimiting an annular space (E) in which the chamber is situated, and a heat exchanger. The heat exchanger can include a first circuit supplied with exhaust gas from the turbomachine, and a second circuit connected by first and second volutes respectively to an outlet of the compressor and to the annular space. The first and second volutes can be positioned at an axial distance from one another, and the second volute is can be connected to the annular space by a straightener which is situated at least in part outside the casing and which is integrated into an annular connecting pipe which connects the second volute to this casing.
A hydraulic torque measurement device for an aircraft engine unit includes a hermetically sealed enclosure having a rigid wall and a deformable wall that define an internal volume. The device further includes at least one pressure measuring element capable of measuring the pressure prevailing in the internal volume. The hydraulic torque measuring device allows a service state in which the internal volume is filled exclusively with an incompressible liquid, and in which the deformable wall is configured to be subjected to a pressure force. In some embodiments, the hydraulic torque measurement device is suitable for use with a gearbox.
G01L 3/14 - Rotary-transmission dynamometers wherein the torque-transmitting element is other than a torsionally-flexible shaft
G01L 5/12 - Apparatus for, or methods of, measuring force, work, mechanical power, or torque, specially adapted for specific purposes for measuring axial thrust in a rotary shaft, e.g. of propulsion plants
26.
Method for quickly stopping the rotor of a helicopter after landing
A method for quickly stopping the propulsion rotor of a helicopter after landing, comprising, following a request for quickly stopping the engine by a helicopter pilot, the following steps managed by the control unit of the turbomachine: Detecting the absence of the thermal stabilization phase of the gas generator of at least one turbomachine, controlling an extinction of the combustion chamber of the gas generator of at least one turbomachine, maintaining the rotation of the gas generator of which the combustion chamber is extinguished by means of said at least one electrical machine to ventilate the gas generator and stopping the main rotor of the helicopter by means of a mechanical brake.
F01D 21/12 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to temperature
A pre-vaporisation tube for a turbine engine combustion chamber includes a main body ROOM defining a first inner duct configured to have an injector mounted therein. The tube includes a first end attached to a wall of the chamber, and at least two end pieces are arranged at a second end of the body and define second inner ducts. The end pieces include first portions and second portions, respectively. The second portions each include two coaxial cylindrical walls which are inner and outer coaxial cylindrical walls, respectively, and which define an annular cavity therebetween. The inner wall defines an inner passage and has first openings for fluid communication between the passage and the annular cavity.
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F23R 3/32 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices being tubular
28.
MODULAR TRAY FOR THE POWDER BED ADDITIVE MANUFACTURING OF A PART WITH AN AXIS OF REVOLUTION
A modular tray, for the additive manufacturing of a part with an axis of revolution on a powder bed, includes: a shaft-mounted circular module including a shaft provided with a circular tray at one of the ends thereof, the shaft and the circular tray being concentric; and a main support module including, in one face, a cavity configured for receiving the shaft-mounted circular module, the shaft being completely inserted in the cavity. The assembly of the shaft-mounted circular module and of the main support module define a planar top surface that is at least partly formed by the circular tray of the shaft-mounted circular module.
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B22F 10/66 - Treatment of workpieces or articles after build-up by mechanical means
B22F 12/82 - Combination of additive manufacturing apparatus or devices with other processing apparatus or devices
B22F 10/64 - Treatment of workpieces or articles after build-up by thermal means
B33Y 30/00 - ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING - Details thereof or accessories therefor
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
B29C 64/153 - Processes of additive manufacturing using only solid materials using layers of powder being selectively joined, e.g. by selective laser sintering or melting
A turbomachine includes a compressor including variable-pitch stationary vanes each extending radially between a rotary hub and a stationary casing surrounding this rotary hub, each variable-pitch vane including a blade having a base spaced apart by a first radial gap from a stationary wall of the casing, and a tip spaced apart by a second radial gap from a rotary wall of the rotary hub. The stationary wall of the casing or the rotary wall of the rotary hub includes at the blade a shape treatment arranged to channel an air leak passing through the corresponding gap.
A combustion assembly for a gas turbine includes a flame tube and a fuel supply including a flow rate limiter supplying an injector. The flow rate limiter and the fuel injector are formed in one piece.
Turbomachine hollow blade (11) comprising at least one vane (14) having lateral walls (15) which are intended to guide a flow in a flow path around the vane and which are fixed to a first platform (12) at a first longitudinal end of the vane (14), the vane (14) further comprising an internal cavity between the lateral walls (15), which cavity is intended for passing a vane-cooling fluid, with a fluid inlet opening (19) opening through said first platform (12), characterized in that a gyroid surface network (18) fills at least part of the cavity, being arranged therein so as to guide the cooling fluid, and is in contact with at least part of the lateral walls (15).
A propulsion system for an aircraft includes a rotor and a nacelle fairing that extends around the rotor in relation to an axis. The nacelle fairing includes an upstream portion forming an inlet section of the nacelle fairing as well as a downstream portion, a downstream end of which forms an outlet section of the nacelle fairing. The downstream portion includes radially inner and outer walls, both of which are made of a deformable shape memory material. The wall has independently actuatable piston actuator mechanisms, each actuator mechanism being actuatable independently of the others and being designed to cooperate with means built into an inner surface of the wall to deform the wall in a radial direction in relation to the axis under the effect of a predetermined displacement command.
F01D 25/00 - Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
F02K 1/00 - Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
33.
ELECTROMECHANICAL INSTALLATION FOR AN AIRCRAFT WITH A TURBOGENERATOR, METHOD FOR EMERGENCY SHUTDOWN OF AN AIRCRAFT TURBOGENERATOR AND CORRESPONDING COMPUTER PROGRAM
An aircraft with an electrical network including electrical subnetworks; a turbo generator including a gas turbine, an electricity generator with permanent magnets having phase groups respectively connected to the electrical subnetworks, and, for each phase group, an isolation device; and a control device designed to detect a short circuit in at least one of the phase groups, each phase group in which a short circuit is detected being described as defective and each other phase group being described as healthy and, in response to the detection of the short circuit, to disconnect this defective phase group from its associated electrical subnetwork and to command the shutdown of the gas turbine. The control device is also designed, in response to the detection of the short circuit, to keep each healthy phase group connected to its electrical subnetwork.
B64D 41/00 - Power installations for auxiliary purposes
H02H 7/06 - Emergency protective circuit arrangements specially adapted for specific types of electric machines or apparatus or for sectionalised protection of cable or line systems, and effecting automatic switching in the event of an undesired change from norm for synchronous capacitors
A fuel supply circuit of an aircraft engine includes a centrifugal pump mechanically coupled with an engine shaft delivering mechanical power. The circuit further includes at least one electromagnetic pump including at least one stator delimiting an annular internal volume in which is present a rotor able to drive a fluid, a plurality of magnets annularly distributed on the rotor and at least a plurality of coils annularly distributed inside the stator face-to-face with the magnets. The rotor is connected to the engine shaft by a one-way clutching element.
The invention relates to a rotor blade (10) for a turbomachine, in particular of an aircraft, comprising an airfoil (12) comprising a pressure face (15) and a suction face (17) extending from a leading edge (14) to a trailing edge (16), the airfoil (12) comprising an axis of elongation extending substantially along the leading (14) and trailing (16) edges, the airfoil (12) comprising a radially firmer end for connection to a rotor and a free radially outer end. According to the invention, the airfoil (10) further comprises at least one series of fins (24) situated on said free end, each of these fins (24) comprising a pressure face (26) situated on the suction face (17) side of the airfoil (12), and a suction face (28) situated on the pressure face (15) side of the airfoil (12).
A turbogenerator (1) for an aircraft (2) comprising: a turboshaft engine (3); an electric generator (4) comprising a rotor (5) driven mechanically by the turboshaft engine (3) and a stator (6) supported by a housing (7) of the electric generator (4); characterized in that the turbogenerator (1) comprises a static separator (8) for separating an air/oil mixture coming from the turboshaft engine (3), the static separator (8) being positioned around the housing (7) of the electric generator (4).
A method for manufacturing a blade with a first portion and a second portion, the method includes forming the first portion that includes forming a model of the first portion from removable material, forming a first shell mould from the model of the first portion, and forming the single-crystal or columnar first portion m a first metal alloy in the first shell mould from a single-crystal seed, and forming the second portion in which the second portion is formed on the first portion, and in which the first portion and the second portion are made from different materials, the second portion being polycrystalline and formed from a second metal alloy. The blade includes a single-crystal or columnar first portion made from a first metal alloy and a polycrystalline second portion made from the second metal alloy different from the first metal alloy.
Rolling bearing with three contact points, wherein the inner raceway, the third contact point being located in a sector of the outer raceway delimited by a proximal end of the outer raceway on the one hand, and by the rolling plane on the other hand, characterized in that the outer ring comprises a drain provided in the outer raceway, and opening from an outer surface of the outer ring, said drain opening eccentrically on the outer raceway with respect to the rolling plane, in a sector of the outer raceway delimited by the rolling plane on the one hand, and a distal end of the outer raceway on the other hand, the drain and the third contact point being disjoint.
F16C 33/66 - Special parts or details in view of lubrication
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
39.
Device for supplying fuel to a combustion chamber of a gas generator
A device for supplying fuel to a combustion chamber of a gas generator includes an injection wheel (14) for injecting fuel into the combustion chamber (18),—a fuel supply rail (20) including an internal fuel circuit (30) with a fuel outlet means (32) supplying fuel to an annular spray chamber (24) formed between the rail (28, 44, 48, 52, 56) and the injection wheel (14),—at least one dynamic annular seal (26) adapted to provide a seal between an annular face (34) of the fuel supply rail (28, 44, 48, 52, 56) and the injection wheel (14), wherein the internal fuel circuit (30) of the fuel supply rail includes an annular fuel flow part arranged radially at the dynamic annular seal (26).
A hybrid propulsion chain for an aircraft, the hybrid propulsion chain comprising a plurality of propulsion rotors connected to an electrical distribution module by a plurality of electrical connections, the electrical distribution module being connected, on the one hand, to a non-propulsion turbine engine via an electrical generation system and, on the other hand, to an electric battery, each propulsion rotor comprising a stator member and at least one rotor shaft which is configured to be rotated with respect to the stator member when the stator member is electrically powered, the hybrid propulsion chain comprising an auxiliary mechanical drive system mechanically connected to the non-propulsion turbine engine, the auxiliary mechanical drive system comprising a plurality of mechanical connections for mechanically rotating at least one rotor shaft of each propulsion rotor.
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
B64D 35/04 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission driving a plurality of propellers or rotors
H02K 7/18 - Structural association of electric generators with mechanical driving motors, e.g.with turbines
41.
SYSTEM AND METHOD FOR BALANCING AT LEAST ONE PARAMETER TO BE BALANCED OF AN ELECTRIC MOTOR OF A PROPULSION SYSTEM
A system (11) for balancing at least one parameter to be balanced of an electric motor of a propulsion system (1), in particular of an aircraft, includes at least two electric motors (3, 4) and a propulsion member (2) driven in rotation by said electric motors. The balancing system is configured to calculate a correction of the speed setpoint (Corr_Cons_VI, Corr_Cons_V2) as a function of a correction factor (F1, F2) of the speed setpoint depending on a parameter (P1, P2) of the associated electric motor that is intended to be balanced and on a speed setpoint (Cons_VH) of the propulsion member (2).
B64D 31/12 - Initiating means actuated automatically for equalising or synchronising power plants
H02P 5/50 - Arrangements specially adapted for regulating or controlling the speed or torque of two or more electric motors for speed regulation of two or more dynamo-electric motors in relation to one another by comparing electrical values representing the speeds
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
A propulsion system (1) for a helicopter, comprising a turboshaft engine (2) with a linked turbine and an electric machine (3) capable of operating as an electric motor, the turboshaft engine (2) and the electric machine (3) being capable of driving in rotation at least one main rotor (5) intended to be coupled to a rotating wing (6) characterised in that it comprises means of coupling and decoupling (14) in rotation between a rotor (3a) of the electric machine (3) and a rotor (2a) of the turboshaft engine (2), the means of coupling and decoupling (14) being capable of allowing the rotor (2a) of the turboshaft engine (2) to be driven in rotation with the aid of the electric machine (3), in a first state of the propulsion system (1), and capable of allowing the rotor (2a) of the turboshaft engine (2) and the rotor (3a) of the electric machine (3) to be decoupled in rotation, in a second state of the propulsion system (1).
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
A turbomachine comprises a rotating spool comprising a drive shaft delivering mechanical power. The turbomachine comprises an electromagnetic pump mechanically decoupled from the drive shaft. The electromagnetic pump comprises at least one stator delimiting an annular internal volume in which is present a rotor able to drive a fluid, a plurality of magnets distributed annularly on the rotor and at least one plurality of coils distributed annularly inside the rotor. The coils of the plurality of coils face magnets along an axial direction.
F04C 2/10 - Rotary-piston machines or pumps of intermeshing-engagement type, i.e. with engagement of co-operating members similar to that of toothed gearing of internal-axis type with the outer member having more teeth or tooth-equivalents, e.g. rollers, than the inner member
F04C 15/00 - Component parts, details or accessories of machines, pumps or pumping installations, not provided for in groups
F04C 7/00 - Rotary-piston machines or pumps with fluid ring or the like
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
44.
BLADE FOR A TURBINE ENGINE, AND ASSOCIATED TURBINE ENGINE
A blade for a turbine engine includes an aerodynamic airfoil extending radially outwards, a blade tip, a blade root, and a platform connected to the root by an upper end. A maximum radial distance between the blade tip and the platform defines a maximum extent of the bearing surface of the airfoil. The blade is partially covered by a protective strip to combat oxidation and corrosion. The strip extends between a lower limit situated on at least one portion of the platform and an upper limit situated on the blade root. The blade tip is not covered by the protective strip.
Disclosed is a movable vane (1) for a wheel (2) of an aircraft turbine engine, the vane (1) comprising a blade (4) delimited by an outer heel (8) comprising a first seal (14), the vane (1) comprising an internal circuit (16) suitable for receiving a first minor gas flow (f1), this circuit (16) comprising a supply cavity (17) opening at the root (9) via at least one inlet opening (18), characterised in that the circuit (16) comprises at least two channels (19) connected with the supply cavity (17) and each opening on an outer surface of the first seal (14) via a discharge opening such that a gas jet (J) of the first minor gas flow (f1) is capable of being discharged from each discharge opening, each channel (19) being oriented such that the corresponding gas jet (J) is capable of being projected towards a second minor gas flow (f2) escaping between the heel (8) and a directly adjacent member (22).
F01D 5/22 - Blade-to-blade connections, e.g. by shrouding
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 11/00 - Preventing or minimising internal leakage of working fluid, e.g. between stages
46.
CONSTANT-VOLUME COMBUSTION SYSTEM WITH SYNCHRONIZED INJECTION
A constant volume combustion system for a turbomachine comprises a plurality of combustion chambers distributed annularly about an axis, each combustion chamber comprising an inlet orifice and an outlet orifice, a selective shut-off element movable in rotation relative to the combustion chambers, the selective shut-off element comprising a shroud facing the inlet and outlet orifices of the combustion chambers. The shroud including on a first annular portion at least one intake aperture intended to cooperate with the inlet orifice of each combustion chamber during the rotation of the selective shut-off element and on a second annular portion at least one exhaust aperture intended to cooperate with the outlet orifice of each combustion chamber during the rotation of the selective shut-off element. Each combustion chamber comprises a fuel injection device whose opening and closing are synchronized by the shut-off element.
F02C 5/12 - Gas-turbine plants characterised by the working fluid being generated by intermittent combustion the combustion chambers having inlet or outlet valves, e.g. Holzwarth gas-turbine plants
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
The present invention relates to a steel part for use in aeronautics, comprising a substrate, the substrate comprising at least carbon, cobalt, aluminium and nickel, and having an average atomic fraction of carbon between 0.09% and 0.17%, an average atomic fraction of cobalt between 15.5% and 18.5%, an average atomic fraction of aluminium less than 0.1%, an average atomic fraction of nickel between 7.2% and 9.8%, the part being case-hardened and also comprising a nitrided layer, the nitrided layer at least partially covering the substrate and having a thickness between 5 μm to 180 μm, preferably between 50 μm and 150 μm.
C21D 9/32 - Heat treatment, e.g. annealing, hardening, quenching or tempering, adapted for particular articles; Furnaces therefor for gear wheels, worm wheels, or the like
C22C 38/10 - Ferrous alloys, e.g. steel alloys containing cobalt
C22C 38/06 - Ferrous alloys, e.g. steel alloys containing aluminium
C23C 8/34 - Solid state diffusion of only non-metal elements into metallic material surfaces; Chemical surface treatment of metallic material by reaction of the surface with a reactive gas, leaving reaction products of surface material in the coating, e.g. conversion coatings, passivation of metals using gases more than one element being applied in more than one step
48.
Device and method for cold expansion of through bore
The invention relates to a cold expansion device (36) for work hardening a through bore (34) in a turbine engine part (32), comprising at least one chuck (38) supporting a burnisher (40) and a means (41) for pushing the chuck (38) in an axial direction, and is characterised in that it comprises at least: —a tubular guide (42) configured to guide the burnisher (40) to the bore (34), —a first magnetic attachment means (51) arranged at a free end (44) of the chuck (38), —the burnisher (40), comprising a work surface (46) and an end (48) comprising second magnetic attachment means (49) complementary to the first magnetic attachment means (51), and in that the axial pushing means (41) is configured to push the burnisher through the bore (34) until it emerges from stud bore.
B24B 39/02 - Burnishing machines or devices, i.e. requiring pressure members for compacting the surface zone; Accessories therefor designed for working internal surfaces of revolution
A fuel injector for an aircraft turbine engine includes a tubular body having an axis of elongation. A first longitudinal end configured to be Supplied with fuel and a second longitudinal end configured to elect a jet of fuel. The body further includes an integrated purge-air circuit that has an internal cavity which is connected to air inlet orifices situated on the body and to at least one air outlet situated at said second end. Aire-flow disruptors are provided, projecting into said cavity.
An aeronautical turbogenerator for hybrid electric propulsion includes a heat engine and an electrical generator coupled mechanically to the heat engine and including a rotor and a stator, the rotor extending in an axial direction and including a common magnetized rotor yoke comprising a plurality of permanent magnets defining at least three axially distributed movable annular rings, the stator including a magnetic stator yoke comprising a plurality of electrical windings defining axially and/or circumferentially distributed stationary sectors, at least two stationary sectors, one of which covers axially at least two movable annular rings, being arranged angularly so as not to mutually coincide and thus deliver at least two distinct and independent voltage levels.
H02K 7/18 - Structural association of electric generators with mechanical driving motors, e.g.with turbines
H02K 21/14 - Synchronous motors having permanent magnets; Synchronous generators having permanent magnets with stationary armatures and rotating magnets with magnets rotating within the armatures
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type
B64D 27/02 - Aircraft characterised by the type or position of power plant
51.
RING FOR A TURBOMACHINE OR A TURBOSHAFT ENGINE TURBINE
The invention relates to a ring (1) for a turbomachine turbine or a turboshaft engine turbine, intended to surround an impeller (2) of a turbine rotor, the said ring (1) extending circumferentially about an axis and comprising an annular and continuous support part (9), radially external, and a part (10) delimiting a circulation passage (6) of a gas flow, radially internal and comprising a plurality of angular segments (13) distributed over the periphery and situated adjacent to one another so as to form an annular part delimiting the passage (6), characterised in that circumferential clearances (j) are formed between the circumferential ends of the adjacent segments (13) located opposite each other, each segment (13) being connected to the support part (9) by means of a connecting zone (14), an annular channel (15) for the circulation of cooling fluid being delimited radially between the outer support part (9) and the inner part (10) delimiting the passage.
F01D 11/04 - Preventing or minimising internal leakage of working fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
52.
Component for a turbomachine centrifugal degasser with adapted longitudinal walls
A component for a centrifugal degasser for an air/oil mixture of a turbomachine is disclosed. The degasser rotates about an axis of symmetry, forming an annular chamber for centrifugal separation of the mixture. The chamber forms a fluid passage duct, one inlet of which is oriented axially for supplying the chamber with the mixture, and a first outlet of which is oriented radially inwards for discharging the deoiled air separated from the mixture. The chamber also includes at least one second oil outlet oriented radially outwards and intended to discharge the oil separated from the mixture to the outside of the degasser. The chamber has longitudinal walls passing radially therethrough, at least one of the surfaces of the longitudinal walls having surface structures and/or corrugations arranged to form obstacles to a flow of the mixture along the surface of the longitudinal walls.
B01D 45/14 - Separating dispersed particles from gases or vapours by gravity, inertia, or centrifugal forces by centrifugal forces generated by rotating vanes, discs, drums or brushes
B04B 7/18 - Rotary bowls formed or coated with sieving or filtering elements
B04B 11/00 - Feeding, charging, or discharging bowls
A method for manufacturing a flame tube for a turbomachine, the flame tube extending about an axis and comprising an annular radially internal wall and an annular radially external wall, connected to each other by an end wall or head wall, the internal wall, the external wall and the end wall defining an internal volume, at least one part of the said end wall forming a double wall comprising a first part and a second part connected to each other and spaced apart from each other so as to delimit a flow channel for a flow of cooling air opening into the said internal volume, the said flow channel comprising at least one air-inlet opening, the first and second parts of the double wall being connected by connecting zones or bridges extending into the flow channel for the cooling-air flow, the flame tube being manufactured by additive manufacturing.
The invention relates to a pre-vaporizing pipe (5) for a combustion chamber of a turbomachine, comprising a generally elongate main body (50), comprising a first end portion (51) delimiting an inlet duct (52) in which a fuel injector is to be mounted, the body (50) having an external surface (500) having a first length (L1) that extends from the portion (51) to at least one first fuel outlet end piece (53), and a second length (L2) that extends the first length (L1) from the end piece (53) and that extends from the end piece (53) to a second fuel outlet end piece (54). According to the invention, inside the body (50) there is a first channel (55) connecting the duct (52) to the end piece (53) and a second channel (56) which is distinct from the first channel (55) and which connects the duct (52) to the end piece (54).
F23R 3/32 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices being tubular
A turbine vane of a turbine engine is described. The turbine vane includes a blade and a root. The root includes a stilt having lateral flanks with a curvilinear profile. The stilt includes a frangible zone suitable for undergoing a breakage of the stilt if radial forces higher than a threshold are exerted on the vane, in particular centrifugal forces during an overspeed state of the turbine. The frangible zone includes at least one oblong frangibility recess formed on at least one of the lateral flanks of the stilt, the oblong recess extending in an axial direction of the stilt along a longitudinal axis parallel to or included in a minimum cross-sectional plane which contains a minimum cross-section of the stilt.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
56.
ANNULAR COMPONENT FOR SUPPORTING A TURBINE ENGINE BEARING
A component (1, 2) for supporting at least one bearing (3) for a turbine engine (10) comprising: two coaxial walls, internal (4) and external (5) walls respectively, defining a gas flow vein (6) between them and interconnected by a row of arms (7); an external ferrule (50) comprising an internal peripheral edge (51) connected to the external wall (5) and an external peripheral edge (52) connected to an external mounting flange (53); an internal ferrule (40) comprising an external peripheral edge (41) connected to the internal wall (4) and an internal peripheral edge (42) comprising an internal mounting flange (43); at least one of the ferrules (4, 5), which at the peripheral edge (41, 51) thereof is connected to the corresponding wall (4, 5), having a general shape which is corrugated about an axis (X-X) of the component (1, 2).
Sealing ring for a wheel of an aircraft turbomachine turbine, the ring having an annular body extending around an axis of revolution and including an outer surface and an inner surface which is coated with an annular layer of an abradable material, the ring further including an annular wall extending around the annular body and at a radial distance from the body, the annular wall including openings through which cooling air flows by impact on the outer surface, wherein the body and the wall are integrally formed.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 11/08 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator
F01D 11/14 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
A module for an aircraft turbine engine including at least one annular casing of an annular combustion chamber, at least one sealing ring for a turbine wheel, and at least one annular bearing support, wherein this module is made in one piece.
A hybrid propulsion installation for an aircraft, including an internal combustion engine having at least one combustion chamber; at least one electricity generator; an electrical propulsion system; electromechanical protection mechanism arranged between the or each electricity generator and the electrical propulsion system; wherein the installation also includes: an electronic computer configured to: receive at least one piece of information on a position of the electromechanical protection mechanism; receive at least one piece of information on a speed of rotation of a shaft of the internal combustion engine; and control the flow of fuel supplying the combustion chamber depending on the information received.
B60L 50/10 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by engine-driven generators, e.g. generators driven by combustion engines
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
A propulsion assembly for aircraft includes a single-shaft engine turbomachine including a combustion chamber and a rotatably mounted shaft that turns at a turbomachine rating; an electrical generator coupled to the shaft; and a control system. The control system includes a fuel pump that brings fuel into the combustion chamber at a fuel flow rate which is a direct function of an ambient pressure and of the turbomachine rating, and power electronics that are coupled to the electrical generator and that control an electrical power drawn off the electrical generator so as to attain a target turbomachine rating.
A process for three-dimensional printing of a workpiece (10), comprising a succession of steps for producing a layer (18) of the workpiece (10) by means of a nozzle (28) which has an output cross-section along a path (32) of the nozzle (28), wherein, during at least one step for producing a layer (18), use is made of at least one nozzle (28), an output portion (42) of which has an output cross-section which is variable between a first maximum cross-section and a second minimum cross-section, and wherein the cross-section of at least one part of an output portion (42) of the nozzle is varied along at least one portion (32a, 32b) of the path (32).
B29C 64/118 - Processes of additive manufacturing using only liquids or viscous materials, e.g. depositing a continuous bead of viscous material using filamentary material being melted, e.g. fused deposition modelling [FDM]
An electrical energy supply network of an aircraft equipped with a plurality of electrical loads to be supplied includes at least two generators of a turbogenerator of the aircraft, each suitable for providing an electrical energy source, and at least one stator associated with a rectifier. The stators of the generators are mounted in parallel on at least two distribution buses designed to supply the plurality of electrical loads. The supply network also includes contactors suitable for electrically connecting or disconnecting the distribution buses from each other.
The present invention relates to a pinion for a gear train of an aircraft turbine engine, the pinion comprising: —a cylindrical body (2) extending along an axis and configured to engage with a shaft received in the cylindrical body, —a rim (4) concentric with the cylindrical body, —a web (3) defined axially by a front wall (32) and a rear wall (33) and extending radially from the cylindrical body to the rim, each of the front and rear walls having a density, the pinion comprising a cross-linked structure (5) around the cylindrical body between the front wall and the rear wall, the cross-linked structure comprising a unit cell repeated along three axes of a three-dimensional coordinate system, a density of the unit cell being strictly less than each of the densities of the front and rear walls.
A connection includes a widened reinforcement harness which comprises a reinforced harness, a non-deformable ferrule and a reinforcement portion which has a greater diameter than the reinforcement of the harness, the non-deformable ferrule being arranged around the harness with the reinforcement of the harness folded on the outer side of the non-deformable ferrule, a first end of the reinforcement portion being attached around the folded portion of the reinforcement of the harness, the connection further comprising an integrated coupling connector, a second end of the reinforcement portion of the widened reinforcement harness being attached to a section of the integrated coupling connector. This connection is particularly advantageous for the small harness, wherein the reinforcement of the harness has a diameter which is too small to surround the section of an integrated coupling connector.
A hybrid propulsion system for a vertical take-off and landing aircraft comprising at least one combustion engine driving an electricity generator, at least one electrical energy storage assembly associated with each electricity generator and defining, with each electricity generator, an energy branch, a plurality of electric motors actuating a same plurality of rotors providing together the propulsion and/or the lift of the aircraft, and an electrical power and distribution unit supplying power to the plurality of electric motors from the electricity generator and/or from the electrical energy storage assembly according to a pre-established flight phase, the system including at least two energy branches having an asymmetric configuration and each supplying power selectively, by means of the electrical power and distribution unit, all or part of the plurality of electric motors, and in that the electricity generators of the combustion engines have between them a power ratio comprised between
An assembly including a winding and a cooler in contact with conductors of the winding, the cooler including a container forming a heat dissipator and including a phase change material having the ability to absorb a surplus quantity of heat when the conductors of the winding are subject to an increase in their current density and the heat dissipator is composed of at least two hollow parts separate and nestable with one another and intended to house the phase change material.
H02K 9/22 - Arrangements for cooling or ventilating by solid heat conducting material embedded in, or arranged in contact with, the stator or rotor, e.g. heat bridges
H02K 9/20 - Arrangements for cooling or ventilating for machines with closed casing and closed-circuit cooling using a liquid cooling medium, e.g. oil wherein the cooling medium vaporises within the machine casing
A system monitors the health of a helicopter, and includes a device for determining a change of state of the engine and is configured to collect data measured by engine and external conditions sensors during a stable flight phase and to process the measured data.
B64F 5/60 - Testing or inspecting aircraft components or systems
B64D 45/00 - Aircraft indicators or protectors not otherwise provided for
B64D 43/00 - Arrangements or adaptations of instruments
F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
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
68.
METHOD FOR CONTROLLING AN ELECTRICAL POWER SUPPLY NETWORK FOR AN AIRCRAFT
Methods for controlling an electrical power supply network for an aircraft, include pre-charging at least one capacitor in a normal operating mode when a gas turbine is available, pre-charging the at least one capacitor in a backup operating mode when the gas turbine is not available, and discharging the at least one capacitor at the end of use of the electrical power supply network when the at least one capacitor is charged.
A toothed coupling mechanism for an assembly of rotating elements of an aircraft gas turbine engine includes a pair of coupling halves having an axial toothed coupling interface therebetween. Each coupling half has a plurality of splined teeth inter-engaged about an axis for transmitting torque therebetween. A protrusion is located on one of the splined teeth of one of the coupling halves. A splined tooth of the other coupling half comes into contact with the protrusion in a situation of uncoupling of said coupling halves.
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
F01D 5/06 - Blade-carrying members, e.g. rotors - Details thereof, e.g. shafts, shaft connections
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F16D 1/10 - Quick-acting couplings in which the parts are connected by simply bringing them together axially
70.
Method and system for regulating a non-propulsion electrical generation turbomachine
A method for controlling a non-propulsive power generation turbine engine configured to supply power to a plurality of propulsion rotors of an aircraft, each propulsion rotor being connected to a power distribution module through at least one power supply bus, the turbine engine supplying each power supply bus via the power distribution module at a supply rate, the control method comprising a step of determining the power requirement of each power supply bus depending on the power requirement of each propulsion rotor, a step of determining the basic power requirement of each power supply bus, a step of determining the overall power requirement based on all the basic power requirements of the power supply buses and a step of determining an anticipation parameter based on the overall power requirement.
B64D 31/00 - Power plant control; Arrangement thereof
B64D 41/00 - Power installations for auxiliary purposes
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
A hybrid propulsion unit for an aircraft with multi-rotor rotary wings includes an electrical generator driven by an internal combustion engine, a rectifier configured to convert an AC current sent by the electrical generator into DC current, a DC-AC converter, an electrical network connecting the rectifier to the converter and including a high-voltage DC current bus, electric motors powered by propeller converters coupled to the electric motors, electrical energy storage connected to the electrical network, the electrical storage including at least one primary storage element and at least one secondary storage element.
B60L 53/20 - Methods of charging batteries, specially adapted for electric vehicles; Charging stations or on-board charging equipment therefor; Exchange of energy storage elements in electric vehicles characterised by converters located in the vehicle
F02B 61/00 - Adaptations of engines for driving vehicles or for driving propellers; Combinations of engines with gearing
72.
HYBRID PROPULSION SYSTEM AND METHOD FOR CONTROLLING SUCH A SYSTEM
A hybrid propulsion system for a multi-rotor rotary-wing aircraft, including: an internal combustion engine, an electric machine coupled to the internal combustion engine, a rectifier connected to the electric machine, a converter, an electric network connecting the rectifier to the converter, electric motors connected to the converter, rotary-blade assemblies coupled to the electric motors, wherein the system includes: a detector configured to detect a reduction in a demand for electrical power within the system to below a predetermined value, a bypass circuit configured to bypass the electric machine when the detector detects a reduction in the demand for electrical power.
HVDC), either one of these two feedback loops delivering an RMS current setpoint Idref and Iqref for a feedback loop on the basis of a current (Igen) of the electrical source delivering an AC voltage.
H02M 5/458 - Conversion of ac power input into ac power output, e.g. for change of voltage, for change of frequency, for change of number of phases with intermediate conversion into dc by static converters using discharge tubes or semiconductor devices to convert the intermediate dc into ac using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
H02J 3/36 - Arrangements for transfer of electric power between ac networks via a high-tension dc link
B60L 50/15 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by engine-driven generators, e.g. generators driven by combustion engines with additional electric power supply
B64D 27/02 - Aircraft characterised by the type or position of power plant
74.
HYBRID-ELECTRIC PROPULSION ARCHITECTURE AND METHOD FOR DISSIPATING ELECTRICAL ENERGY IN SUCH AN ARCHITECTURE
A hybrid/electric propulsion architecture for a multi-rotor rotary wing aircraft, including an electricity generator driven by an internal combustion engine, and configured to operate in motor mode, a rectifier configured to convert an alternating current delivered by the electricity generator into direct current, an electrical network including a high voltage direct current (HVDC) bus, electrical energy storage means connected to the electrical network, during electrical energy regeneration on the HVDC bus, depending on the state of charge of the storage means: the storage means are configured to recover electrical energy, the storage means and the rectifier are configured to recover electrical energy, and the electricity generator operating in motor mode is configured to recover electrical energy.
A method for manufacturing a compressor impeller or rotor including a hub that carries blades, involving a step of manufacturing a hub which includes all or some of the blades, and a step of additive manufacture by adding localised material using a method such as the LMD process to form or finish each blade.
B22F 7/08 - Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting of composite workpieces or articles from parts, e.g. to form tipped tools with one or more parts not made from powder
F04D 29/28 - Rotors specially adapted for elastic fluids for centrifugal or helico-centrifugal pumps
A lattice formed of entwined links in a cavity encased by the external skin of a blade and provided with parts of low mechanical strength referred to as breakable, which are capable of rupturing when sufficient loads are applied thereto during operation of the blade. This arrangement, which concentrates the ruptures at given locations, will leave the rest of the lattice intact and therefore will not hamper or will scarcely hamper its properties of heat evacuation by conduction or convection or pressure loss, if for example ventilation air passes through it.
A method for controlling a turbomachine including a temporary power-increasing device, the control method including a step wherein the flow rate of the coolant injected is adjusted as a function of the atmospheric pressure and/or of the ambient temperature and/or of at least one parameter such as the speed of rotation of a gas generator, the speed of rotation of a low-pressure turbine or of a power turbine, the gas pressure at the outlet of a compressor stage, the temperature at the inlet of the low-pressure turbine or of the power turbine, the engine torque, and/or the collective pitch of a helicopter rotor or the pitch of a propeller of a turboprop.
The present invention relates to an electric machine for an aircraft, comprising a stator and a rotor that is rotationally mobile with respect to the stator, the rotor or the stator comprising a plurality of permanent magnets, the machine comprising a device for demagnetising a permanent magnet, suitable for achieving a temporary increase in the temperature of the permanent magnet, in order to limit, during the temporary increase in temperature, an exciting magnetic flux generated by the permanent magnet. The present invention furthermore relates to an assembly comprising an assembly comprising such an electric machine and a hot-fluid source suitable for delivering hot fluid to the demagnetising device of the electric machine. The hot-fluid source may be a gas stream of a turbine engine.
The invention concerns a method for monitoring the operating state of a system for positioning variable-geometry members (18) of a turbomachine (10), the members (18) being configured to travel over an operating area comprising a first position P1 and a second position P2, the method comprising the steps of: —(E23) determining a first pivoting speed V1 from the first position P1 to the second position P2; —(E25) determining a second pivoting speed from the second position P2 to the first position P1; and—(E26) determining an anomaly in the operation of the system for positioning the members (18) if the first speed V1 is lower than a first determined speed threshold and/or if the second speed V2 is lower than a second determined speed threshold.
F01D 17/26 - Devices dealing with sensing elements or final actuators or transmitting means between them, e.g. power-assisted the operation or power assistance being predominantly non-mechanical fluid, e.g. hydraulic
F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
F02C 9/20 - Control of working fluid flow by adjusting vanes
F02C 9/54 - Control of fuel supply conjointly with another control of the plant with control of working fluid flow by throttling the working fluid, by adjusting vanes
80.
Impact-cooling tubular insert for a turbomachine distributor
A tubular ventilation sleeve for a turbomachine distributor, in particular for an aircraft, the sleeve having a generally elongate shape along an axis (A-A) and including a perforated tubular wall around said axis, one of the axial ends of the sleeve being open and the other being closed by a bottom wall, wherein it further includes support beams when the sleeve is made by additive manufacturing, the beams extending inside the sleeve between the tubular wall and the bottom wall and having a longitudinal cross-section with a generally triangular shape, two sides of which are respectively connected to the tubular wall and the bottom wall and the last side of which is free and extends inside the sleeve, perforations in the tubular wall being provided between the support beams.
The invention concerns a turbomachine rotor (1), characterised in that it comprises a threaded or tapped part (3, 6) and a damping nut (8) screwed onto the threaded or tapped part (3, 6) so as to allow the threads of the nut (8) and of the threaded or tapped part (3, 6) to rub against each other in the event of vibration of the rotor (1).
The present disclosure relates to an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines connected to a power transmission gearbox, characterized in that it comprises: at least one hybrid turboshaft engine capable of operating in at least one standby mode during a stable cruise flight of the helicopter; at least two systems for controlling each hybrid turboshaft engine, each system comprising an electric machine connected to the hybrid turboshaft engine and suitable for rotating the gas generator thereof, and at least one source of electrical power for the electric machine, each reactivation system being configured such that it can drive the turboshaft engine in at least one operating mode among a plurality of predetermined modes.
The invention relates to an electrical machine comprising a stator (1) and a rotor (2) designed to be rotated in relation to each other, said rotor (2) or said stator comprising a plurality of permanent magnets (5), at least one permanent magnet comprising at least one fluid-propagation channel (10) extending longitudinally inside the permanent magnet, the propagation channel comprising a fluid inlet and a fluid outlet, the fluid inlet being bell-mouthed and oriented in a preferential direction of rotation of the permanent magnet.
The invention relates to an aircraft power architecture comprising a power transmission gearbox (12), located in a first compartment (30), a gas turbine (14), located in a second compartment (32) comprising a gas generator (18) and a free turbine (22) connected to the power transmission gearbox (12) by a power shaft (26) of the gas turbine (14), and an accessory gearbox (16), the gas turbine (14) being set into the main transmission gearbox (12), characterized in that the gas turbine (14) comprises a first electric machine (38), and in that the accessory gearbox (16) is placed in the first compartment (30) and comprises a second electric machine (42) configured to supply energy to the accessory equipment and to receive electrical energy transmitted via the first electric machine (38).
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
An assembly for a multistage turbine of a turbomachine has a static sealing device and a nozzle with a radially outer end and an outer casing surrounding the nozzle. The static sealing device is arranged radially between a radially outer end of the nozzle and the outer casing. The static sealing device includes an annular seal borne by the nozzle and an annular structure that defines a plurality of radial annular walls. The walls are axially spaced apart from one another, and at least one first wall is in annular contact radially inwardly with the annular seal. A longitudinal dimension of the annular contact is less than a longitudinal dimension of the seal.
A propulsion system for a helicopter includes a linked turbine engine that is configured to drive a main rotor configured to be coupled to a rotary wing. The propulsion system further includes an electric machine that is configured to form an electric motor. The electric machine is coupled directly or indirectly to the main rotor.
A method for carburizing a steel workpiece includes implementing between 1 and 30 consecutive carburizing cycles, each carburizing cycle including injecting carburizing gas in such a way as to increase the surface carbon rate until a predetermined higher rate is reached, a step of injecting a neutral gas so as to decrease the surface carbon rate of the workpiece until it reaches a predetermined lower rate, followed by a second phase of injection of the neutral gas.
A tensile testing machine comprising a test specimen whose elongation is to be measured along a tensile axis, slide plates, an intermediate plate, and first and second parallel guide rods, which freely guide the slide plates axially past them.
G01N 3/14 - Investigating strength properties of solid materials by application of mechanical stress by applying steady tensile or compressive forces generated by spring tension
G01N 3/06 - Special adaptations of indicating or recording means
G01N 3/08 - Investigating strength properties of solid materials by application of mechanical stress by applying steady tensile or compressive forces
90.
Connection between a ceramic matrix composite stator sector and a metallic support of a turbomachine turbine
A turbine of a turbomachine includes a ceramic matrix composite sector of a stator includes an outer platform and an inner platform connected via at least one vane. The outer platform has means for attaching to a sector of a metallic support, the attachment means having at least one central rim and two lateral rims. The central rim is radially offset with respect to said lateral rims along a directrix line such that the central rim is radially on one side of said directrix line and the lateral rims on the other. The central rim and said central hook bear radially against one another and are located radially on either side of said directrix line. The lateral rim and said corresponding lateral hook bear radially against one another and are located radially on either side of said directrix line.
A turbomachine rotor has a disk carrying vanes, each vane having a blade linked by a platform to a root. For at least one vane, a recess is defined between the platform and the disk, and a vibration damper is mounted in the recess. The vibration damper includes a first structural portion configured to contact the platform of which the vibrations are to be dampened, and a second mass portion configured to dampen these vibrations. The second mass portion is a powder and the first structural portion is a box containing the powder.
In the field of multi-rotor rotary-wing aircraft a propulsion system for such an aircraft includes a reconfigurable electric network for supplying the electric motors driving the rotors. The system includes: a supply source, a supply bus connected to the supply source, at least four drive units each including an electric motor and its control circuit, and an electrical ring network including: an electrical line which is interrupted at each drive unit and whose ends are connected to the supply bus, and, for each drive unit, a first switch and a second switch that are connected between the control circuit and the electrical line, on either side of the interruption.
A centrifugal degasser of an air-oil mixture of a turbine engine includes a hollow shaft extending along an X axis and a pinion for rotating the hollow shaft, and an annular chamber for centrifugal separation of the mixture. The chamber is arranged around the hollow shaft and forms a fluid passage area, an inlet of which is oriented axially for feeding the chamber with the mixture, and a first outlet of which is oriented radially inwards for the outlet of the de-oiled air separated from said mixture. The chamber also includes at least one second oil outlet oriented radially outwards and configured for discharging the oil separated from said mixture to the outside of the degasser. The chamber and the pinion form a single piece.
F01M 11/08 - Separating lubricant from air or fuel-air mixture before entry into cylinder
B01D 45/14 - Separating dispersed particles from gases or vapours by gravity, inertia, or centrifugal forces by centrifugal forces generated by rotating vanes, discs, drums or brushes
B04B 7/18 - Rotary bowls formed or coated with sieving or filtering elements
B33Y 80/00 - Products made by additive manufacturing
94.
Method for checking the maximum available power of a turbine engine of an aircraft equipped with two turbine engines
A method for checking the maximum available power of a turbine engine of an aircraft equipped with two turbine engines configured to operate in parallel and together to supply a necessary power to the aircraft during a flight phase includes: placing one of the turbine engines in a maximum take-off power regime, and adjusting a power supplied by the other turbine engine, such that the turbine engines continue to supply the necessary power to the aircraft during the flight phase; determining a power supplied by the turbine engine placed in the maximum take-off power regime, and processing the supplied power determined in this way, in order to deduce a piece of information relating to the maximum available power.
A device for temporarily increasing power in order to increase the power from at least one first turbine engine and from at least one second turbine engine, the device including a tank of coolant liquid, a first injection circuit connected to the tank and leading to at least one injection nozzle configured to be installed upstream from the first turbine engine, a second injection circuit connected to the tank and leading to at least one injection nozzle configured to be installed upstream from the second turbine engine, each of the first and second injection circuits including at least one first valve and at least one second valve arranged upstream from said at least one first valve, and a bridge pipe connecting together the first injection circuit and the second injection circuit upstream from their respective first valves and downstream from their respective second valves.
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 6/02 - Plural gas-turbine plants having a common power output
F02C 7/143 - Cooling of plants of fluids in the plant of working fluid before or between the compressor stages
A turbomachine blade formed of a hollow airfoil having a leading edge and a trailing edge opposed to each other and connected by an intrados wall and an extrados wall each extending along a radial axis of the blade, between a blade root and a blade tip, and including a cooling circuit supplied with air and delivering air jets ensuring through multiple perforations of the cooling circuit an impingement cooling of the inner surface of the airfoil, the cooling circuit includes superimposed cooling channels over the height of the blade, each integrated into the inner surface of the airfoil while matching its contour, the multiple perforations being drilled in the cooling channels terminating in a purge cavity of the airfoil are able to ensure a purge of the air having impinged the inner surface of the airfoil after its passage through the perforations.
A labyrinth seal lip (5) comprises, standing on one of its lateral faces (7), a deflector (13) which channels the flow of gas tangent to the lip (5) towards the facing seal (4) so as to disturb the leakage flow (22) through the gap and decrease the flow rate.
A stator vane of a turbine of a gas turbine engine, including an outer platform and an inner platform between which there extends an outer wall forming an outer skin, wherein it includes an inner wall, forming an inner skin, facing the outer wall so as to define an inter-skin cavity between the outer wall and the inner wall, the inner wall including a plurality of cooling orifices for impingement cooling of the outer wall, the outer wall and inner wall being produced by additive manufacturing.
An electrical generator is housed in an annular cavity between the casing and the propeller shaft of a turboprop, while imposing little or no additional space requirement and with lightweight ancillary equipment. The rotor of the generator is mounted on an autonomous shaft end. A flange of the outer casing is removable in order to access the generator and to enable its easy removal and remounting.
An assembly for a turbomachine including an annular channel designed to form a flow duct for a flow of gas between two turbine stages of the turbomachine. The channel is bounded by a radially inner annular wall and a radially outer annular wall. The walls are connected by hollow arms that extend radially, a support having a radially outer annular part that is located radially outside the outer annular wall of the annular channel, and a radially inner annular part that is located radially inside the inner annular wall of the annular channel. The outer and inner parts of the support are connected by connecting parts that extend radially and pass through one of the hollow arms of the annular channel. The connecting parts may be connected by a connecting partition having a frangible part that ruptures when the mechanical stresses in the connecting partition are above a threshold.
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