The invention relates to a turbine engine module, in particular an aircraft turbine engine (10), comprising: - an annular casing (52) having an internal wall (53) forming a channel wall; and - a nozzle (32) surrounded by the casing and comprising an annular external platform (36) and an annular internal platform (37) between which stator blades (34) extend, the external platform having an external face (36b) that faces the internal wall of the casing and comprises an annular groove (60) oriented towards the outside and housing a sealing device (64), the sealing device coming into cylindrical contact with a track (66) of the internal wall (53) of the casing, the module being characterised in that the internal wall (53) of the casing comprises a thermal barrier (70) made of ceramic material directly above the track (66), the track being arranged between the thermal barrier and the sealing device.
The invention relates to a turbomachine (100) for a hybrid aircraft, the turbomachine comprising a gas generator (12) carried by a generator shaft (14), at least one free turbine (11) carried by a turbine shaft (13) and rotated by a gas flow generated by the gas generator (12), a main rotor (60), and at least one reversible electric machine (30), the turbine shaft (13) being a through-shaft and extending axially between a first end engaged with the electric machine (30) and a second end engaged with the main rotor (60).
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 3/14 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
3.
LUBRICATION AND COOLING OF EQUIPMENT OF AN AIRCRAFT TURBOMACHINE
The invention relates to an aircraft turbomachine (10) comprising: a gas generator (12) comprising an output shaft (26) as well as a first lubricating circuit (28); and equipment (14) coupled to the output shaft (26) and comprising a rotor (38) which is rotationally guided by at least one rolling bearing (40), the equipment (14) comprising a second lubricating circuit (46) which is independent of the first lubricating circuit (28) and which is configured to lubricate the rolling bearing (40), the equipment further comprising a system (50) for cooling the rolling bearing (40), the cooling system (50) being configured to circulate oil in the region of at least one ring (40b) of the rolling bearing, characterized in that the cooling system (50) is independent of the second lubricating circuit (46) and is connected to the first lubricating circuit (28).
A propulsive assembly (100) for a multi-engine hybrid aircraft, comprising a first and a second gas turbine (10, 20) each having a gas generator (12, 22) and a free turbine (11, 21), a main rotor (62) coupled to the free turbine (11, 21) via a first and a second main coupling means (51, 52), a first and a second reversible electric machine (30, 40) each coupled to the gas generator (12, 22) via a first deactivatable coupling means (31, 41), and each coupled to the main rotor (62) via a second deactivatable coupling means (32, 42), the first deactivatable coupling means (31, 41) being activated when the electric machines (30, 40) rotate in a first direction of rotation, and the second deactivatable coupling means (32, 42) being activated when the electric machines (30, 40) rotate in a second direction of rotation opposite to the first direction of rotation.
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
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
One aspect of the invention relates to a filtering device (10) for a turbomachine, the device comprising: - a filter support (14); - a removable filter (18) co-operating with the filter support; - a cover (12); - an indicator pin (16); the device comprising a plate (13) that is translatably movable relative to the cover and relative to the indicator pin, the plate being translatably movable between two positions: - a first position such that the plate bears against the removable filter and the indicator pin is in a retracted position; - a second position such that the plate is remote from the cover due to the absence of the removable filter and the indicator pin is in an extended position.
The invention relates to a turboshaft engine (1) for an aircraft (2) comprising: - a gas generator (3) comprising a compressor (4), a combustion chamber (5) and an expansion turbine (6); - a power turbine (8) rotating a power take-off (9) by means of a reduction gear (10); - a heat exchanger (11) comprising a first circuit (12) and a second circuit (17); characterised in that the compressor (4) comprises a first shaft (22) rotated by a second shaft (23) of the expansion turbine (6) by means of a transmission mechanism (24), the transmission mechanism (24) and the reduction gear (10) forming part of a gearbox (25) which is arranged axially at a front end (20) of the turboshaft engine (1), such that the compressor (4) is arranged axially between the gearbox (25) and the power turbine (8). Drawing_references_to_be_translated
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 7/08 - Heating air supply before combustion, e.g. by exhaust gases
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
7.
DC-TO-DC CONVERTER FOR AN ELECTRICAL AIRCRAFT PROPULSION SYSTEM
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE LORRAINE (France)
Inventor
Barraco, Thomas Michel André Gérard
Klonowski, Thomas
Pierfederici, Serge Lionel
Weber, Mathieu Robert
Abstract
The invention relates to a DC-to-DC converter (209) for an electrical aircraft propulsion system designed to be connected in series with an electrical energy storage unit of the electrical propulsion system. The DC-to-DC converter (209) comprises an inverter (401), a transformer (403) and a rectifier (405), and further comprises a current source (417) that is connected to the rectifier (405) and is configured to control the power passing through the DC-to-DC converter (209). The transformer (403) comprises a primary (403a) and two secondaries (403b, 403c), the two secondaries (403b, 403c) sharing a common terminal (407) designed to be connected to a high-voltage DC bus (207) of the electrical propulsion system (201) and two other terminals (409, 411) that are connected to the rectifier (405). The rectifier (405) comprises two arms comprising at least two transistors (415a, 415b, 415c, 415d) that are each in series and are connected, on the one hand, to the two other terminals (409, 411) of the transformer (403) and, on the other hand, to the current source (417).
H02M 3/335 - Conversion of dc power input into dc power output with intermediate conversion into ac by static converters using discharge tubes with control electrode or semiconductor devices with control electrode to produce the intermediate ac using devices of a triode or a transistor type requiring continuous application of a control signal using semiconductor devices only
H02J 7/00 - Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
H02M 7/219 - Conversion of ac power input into dc power output without possibility of reversal by static converters using discharge tubes with control electrode or semiconductor devices with control electrode using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only in a bridge configuration
8.
AIRCRAFT TURBOMACHINE COMPRISING A DEVICE FOR INHIBITING THE ACCUMULATION OF COKE IN A DUCT
An aircraft turbomachine (10) having a gas generator (12) comprising, along a longitudinal axis (X), at least one compressor (14), a combustion chamber (16) and at least one turbine (18), the turbomachine (10) further having at least one duct (20) for supplying liquid to at least one member chosen from an oil jet (22) and a fuel injector (50), this duct (20) having rectilinear portions (21a, 21b) and bent portions (21c, 21d) and comprising at least one region in which the liquid is liable to coke, characterized in that it comprises at least one turbulence element (40-48) in the at least one region in the duct.
F16L 55/24 - Preventing accumulation of dirt or other matter in pipes, e.g. by traps, by strainers
B08B 9/032 - Cleaning the internal surfaces; Removal of blockages by the mechanical action of a moving fluid, e.g. by flushing
9.
METHOD FOR REGULATING THE SPEED OF ROTATION OF A PROPULSION DEVICE OF A HYBRID PROPULSION UNIT FOR AN AIRCRAFT, IN THE EVENT OF A FAILURE OF THE MAIN REGULATION SYSTEM OF THE HEAT ENGINE OF THE HYBRID PROPULSION UNIT
The invention relates to a method for regulating the speed of a propulsion device of a hybrid propulsion unit for an aircraft comprising: the propulsion device and an MGB; the heat engine and at least one electric motor, mounted in parallel on the MGB, the heat engine being provided with a fuel circuit; main and backup regulation systems for the heat engine, and a regulation system for the electric motor, each being capable of regulating the speed of the heat engine or the electric motor, respectively; and an aircraft control system, capable of sending a speed or power setpoint to each of the regulation means for the heat engine and the electric motor. The method comprises, when the main regulation system for the heat engine fails and is locked to a QCarbP* fuel flow command: - sending a speed setpoint NM2réf to the regulation system for the electric motor, the regulation system for the electric motor sending a power setpoint PM2* to the electric motor, whereby an instantaneous power PM2m of the electric motor is obtained; - simultaneously, sending a speed or power setpoint to the backup regulation system for the heat engine, the backup regulation system sending a fuel flow command QCarbAux* to the fuel circuit of the heat engine, which command is selected to vary the fuel flow QCarb injected into a combustion chamber of the heat engine according to whether the power PM1 of the heat engine is to be varied.
The invention relates to a method for checking the maximum power available to members of a propulsion system of an aircraft (1) comprising first members (M1.1, M2.1, D10, TG, BAT1) that are sized to compensate for the failure of second members of the propulsion system (M1.2, M2.2, D20, TG, BAT2) by delivering a maximum power to keep the aircraft in a safe operating range, the method comprising the following steps for each of the first members: - placing the first member in a state that is substantially equal to a maximum power state (PM); - adjusting the power delivered by the second member working in synergy with the first member so that the first member and the second member contribute to delivering the power required for the aircraft in the flight phase; - determining the power delivered by the first member placed in the maximum power state; - from the determined power, deducing information relating to the maximum power available to the first member.
The invention relates to a control device (30) for controlling an airflow guiding system (20), comprising: - at least one vane (21) that is rotatable on a shaft (21a) of the vane between a first angle and a second angle; - an actuator (31) comprising a body (32) inside which a piston (32a) is translatably mounted, the piston being rigidly connected to a drive rod (33); - a control rod (34) comprising a downstream end (34a) that is connected to the shaft (21a) of the vane, the actuator (31) being configured to drive the piston (32a) between a first end position (P1') and a second end position (P2') of a nominal operating range and the downstream end (34a) of the control rod (34) between a first end position (P1) and a second end position (P2) of a nominal operating range in which the vane (21) is movable between a first angle and a second angle, the device being characterised in that it comprises a drive mechanism (40) that connects an upstream end (33a) of the drive rod (33) to an upstream end (34b) of the control rod (34), which is opposite the downstream end (34a), and in that the drive mechanism (40) is configured so that, in the event of the control device failing, it places the downstream end (34a) of the control rod (34) in a safety position (PS) located between the first end position (P1) and a second end position (P2) of the nominal operating range and in which the vane (21) is oriented at a safety pitch angle between the first angle and the second angle.
The present invention relates to a metal powder for a powder bed additive manufacturing process, the metal powder comprising a nickel-based alloy comprising at least 0.05% carbon, at least 14.25% cobalt, at least 14% chromium, at least 4% aluminium, at least 3.9% molybdenum, at least 3% titanium, at most 0.5% iron, at least 0.012% boron, at most 0.060% zirconium, at most 0.150% manganese, at most 0.2% silicon, at most 0.1% copper, at most 0.5 ppm bismuth, at most 5 ppm silver, at most 5 ppm lead, at most 25 ppm sulphur, at most 200 ppm oxygen, and at most 60 ppm nitrogen.
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
B33Y 50/02 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
B33Y 70/00 - Materials specially adapted for additive manufacturing
B33Y 80/00 - Products made by additive manufacturing
C22C 1/04 - Making non-ferrous alloys by powder metallurgy
C22C 19/05 - Alloys based on nickel or cobalt based on nickel with chromium
C22F 1/10 - Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of nickel or cobalt or alloys based thereon
F01D 5/28 - Selecting particular materials; Measures against erosion or corrosion
13.
SYSTEM AND METHOD FOR AUTOMATICALLY VALIDATING THE RESPONSE COMPLIANCE OF A PENETRANT TESTING LINE
The invention relates to a system (1) for automatically validating the response compliance of a penetrant testing line by analysing a reference test specimen undergoing dye penetrant testing from said penetrant testing line, the system comprising a casing (2) surrounding all the elements of the system (1), a positioning base (3) for the reference test specimen undergoing dye penetrant testing, at least two ultraviolet lighting devices (4) positioned on either side of the positioning base (3) so as to be able to illuminate it with ultraviolet light, an electronic board, a motorised linear guide (6) controlled by the electronic board and suitable for being able to move the positioning base (3) and hold said base in position in the casing (2), and a monochrome camera (10) having a lens directed towards the positioning base (3).
The invention relates to a method for the quantitative measurement of an element (4) in a metal test specimen (1) which has received a thermochemical surface treatment, characterised in that the method comprises the following steps: • - a- making a cut in the test specimen; • - b- taking a photograph, using an optical microscope, of said cut on a surface S with magnification; • - c- calculating the ratio of the surfaces s occupied by the element with respect to the surface S in the photograph by means of an image-processing tool; and • - d- obtaining the content of the element in the surface S of the test specimen. A profile of the content of undesirable elements can thus be obtained without the subjective aspect of human checking, which profile is representative of a content obtained by acquiring diffraction of X-rays produced using a synchrotron.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
SAFRAN HELICOPTER ENGINES (France)
UNIVERSITE DE TOULOUSE III - PAUL SABATIER (France)
Inventor
Richard, Stéphane, Raphaël, Yves
Viguier, Christophe, Nicolas, Henri
Marragou, Sylvain
Schuller, Thierry
Abstract
The invention relates to a dihydrogen injection device (2) having a longitudinal axis (X), intended to be mounted on an annular base of an annular combustion chamber (4) of a turbomachine, comprising an inner channel (6) for circulating dihydrogen, and an annular outer channel (8) for circulating a mixture at least comprising air, the inner channel (6) and the annular outer channel (8) being coaxial, an inner swirler (14) being provided in the inner channel (6) and an outer swirler (28) being provided in the annular outer channel (8), a downstream end (16) of the inner channel (6) being arranged upstream, at a distance r, from a downstream end (24) of the annular outer channel (8). Such combustion of dihydrogen makes it possible to eliminate carbon-containing polluting emissions such as carbon monoxide, unburned hydrocarbons or fine particles and smoke particles.
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
F23D 14/24 - Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other at least one of the fluids being submitted to a swirling motion
The present invention relates to a rotor (1) for an electric motor of an aircraft, the rotor (1) comprising a shaft (2) made of a first material and a conductor assembly (4) made of a second material that differs from the first material, the rotor (1) being characterized in that the shaft (2) has a shoulder portion (6) having at least one longitudinal slot (25) and in that the conductor assembly (4) is a single-piece structure comprising at least one conducting bar (28) intended to be positioned in the at least one slot (25) and including a skin (29) intended to be fastened onto the shoulder portion.
H02K 15/00 - Methods or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines
H02K 17/16 - Asynchronous induction motors having rotors with internally short-circuited windings, e.g. cage rotors
H02K 17/18 - Asynchronous induction motors having rotors with internally short-circuited windings, e.g. cage rotors having double-cage or multiple-cage rotors
H02K 15/02 - Methods or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines of stator or rotor bodies
H02K 1/02 - DYNAMO-ELECTRIC MACHINES - Details of the magnetic circuit characterised by the magnetic material
17.
TURBOMACHINE ELEMENT COMPRISING AT LEAST ONE BLADE OBTAINED BY ADDITIVE MANUFACTURING
The present invention relates to a turbomachine element (1), comprising at least one blade (2) obtained by additive manufacturing, the blade (2) having a skin (4) and an internal lattice (6) allowing air circulation in the blade (2) and having an additive manufacturing support function for the skin (4).
The present invention relates to a rotor (1) of an aircraft electric motor comprising a shaft (2) made of a first material and a skin (4) made of a second material different from the first material, the rotor (1) being characterized in that the shaft (2) has a shoulder portion (6) to which the skin (4) is attached at the shoulder portion (6), the rotor (1) having an interpenetration layer of the first material and of the second material, the interpenetration layer comprising an alloy of the first material and of the second material.
H02K 15/00 - Methods or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines
H02K 17/16 - Asynchronous induction motors having rotors with internally short-circuited windings, e.g. cage rotors
H02K 15/02 - Methods or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines of stator or rotor bodies
H02K 1/02 - DYNAMO-ELECTRIC MACHINES - Details of the magnetic circuit characterised by the magnetic material
19.
METHOD FOR TRAINING TO COPE WITH A FAULT AFFECTING ONE POWERTRAIN OF A HYBRID PROPULSION SYSTEM
The invention relates to a method for training a pilot to cope with a fault affecting one powertrain of a hybrid propulsion system for an aircraft comprising, connected in parallel to a transmission unit, n powertrains (where n≥2), including a first and a second powertrain that are heterogeneous in nature. It involves, during a flight of the aircraft, simulating a fault affecting the first powertrain while, at the same time as performing the simulation, checking the status of the n powertrains of the propulsion system. If a fault affecting one of the n powertrains is detected, the simulation is halted and the instantaneous power delivered by at least one of either the first or the second powertrain is increased so that the sum of the instantaneous powers delivered by the n powertrains is ≥ a minimum total instantaneous power required for the aircraft to continue its flight.
G09B 9/44 - Simulators for teaching or training purposes for teaching control of vehicles or other craft for teaching control of aircraft, e.g. Link trainer providing simulation in a real aircraft flying through the atmosphere without restriction of its path
G09B 9/46 - Simulators for teaching or training purposes for teaching control of vehicles or other craft for teaching control of aircraft, e.g. Link trainer the aircraft being a helicopter
G09B 9/08 - Simulators for teaching or training purposes for teaching control of vehicles or other craft for teaching control of aircraft, e.g. Link trainer
20.
SENSOR FOR DETERMINING A LIQUID LEVEL FOR AN AIRCRAFT TANK, TANK-AND-SENSOR ASSEMBLY, METHOD FOR USING SUCH A SENSOR
The invention relates to a sensor (1) for determining a liquid level (NE) for an aircraft tank (100), the determining sensor (1) comprising a closure device (2) for closing a port (101) of the tank (100) and a measuring device (3), removably mounted on the closure device (2), comprising a liquid line (20) configured to convey liquid from the port (100) of the tank (100), and a member (21) for automatically sealing the liquid line (20) if the measuring device (3) is not mounted on the closure device (2), the measuring device (3) comprising at least one pressure measuring member (30) configured to measure a pressure difference between the liquid pressure (P1) in the liquid line (20) and a reference pressure (P2) in order to deduce the liquid level (NE) thereof.
G01F 23/16 - Indicating, recording, or alarm devices being actuated by mechanical or fluid means, e.g. using gas, mercury, or a diaphragm as transmitting element, or by a column of liquid
G01F 22/02 - Methods or apparatus for measuring volume of fluids or fluent solid material, not otherwise provided for involving measurement of pressure
G01F 23/00 - Indicating or measuring liquid level or level of fluent solid material, e.g. indicating in terms of volume or indicating by means of an alarm
The invention relates to a turboprop (10) comprising a propeller (12), a propeller shaft (13) carrying the propeller (12), the propeller being a variable-pitch propeller having a propeller pitch, a rotating electric machine (19) having at least a first configuration in which it is mechanically coupled to the propeller shaft (13) and at least one oil pump (21, 21B) configured to supply a hydraulic circuit for adjusting the pitch of the propeller (12). The oil pump is configured to be electrically operated. The invention further relates to an aircraft comprising such a turboprop (10) and to the methods for controlling such a turboprop (10) and such an aircraft.
A propeller (2) for an aircraft propulsion assembly (1) extending longitudinally along an axis X, the propeller (2) comprising a propeller cone (21), blades (22), a guide member (4) extending longitudinally along the axis X and rotating as one with the propeller cone (21), the guide member (4) being mounted outside the propeller cone (21) in such a way as to form between them a guide path (V), the guide member (4) having an upstream opening (41) configured to convey a flow of air in the guide path (V) and a downstream opening (42) in such a way as to remove the flow of air downstream, the guide member (4) having through-orifices (40) through which extend the blades (22) of the propeller (2) and compressor vanes (5), which rotate as one with with the propeller cone (21) and which are positioned in the guide path (V) in such a way as to generate an accelerated air flow.
The invention relates to a device (136) for controlling a power-transfer system (124) for the transfer of power between a high-pressure shaft (108) and a low-pressure shaft (116) of a turbomachine (104) of an aircraft (102), comprising: - a fatigue analysis module (142) analysing the fatigue of the turbomachine (104) and designed to determine, from between two indicators (D1, D2) respectively measuring two fatigues of the turbomachine (104), which is the one that is the most advanced, which is to say which is the one at risk of being first to reach a respective upper limit (D1max, D2max); and - a control module (140) controlling the power transfer system (124) and designed to slow the fatigue measured by the more advanced indicator (D1, D2).
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 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
F02C 9/56 - Control of fuel supply conjointly with another control of the plant with power transmission control
F02C 3/113 - 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 with variable power transmission between rotors
24.
FIRE SAFETY SYSTEM FOR A TURBOMACHINE COMPRISING MEANS FOR MAINTAINING A COOLING AIR SPEED AND CORRESPONDING TURBOMACHINE
The invention relates to a fire safety system (30) for a turbomachine (1), such as an aeroplane turbofan or a turboprop engine, the turbomachine comprising at least one turbine (3) having a turbine disc (5) and an annular cavity (16) which is arranged upstream of the disc (5), the disc (5) comprising an internal bore (7) and the fire safety system (30) comprising a cooling device (14) intended to supply the cavity (16) with cooling air via injection means (17). According to the invention, the fire safety system (30) comprises means configured such that, on the one hand, the annular cavity is divided into a first cavity and a second cavity and, on the other hand, a cooling air speed is maintained at the outlet of the injection means (16) and the cooling air in the first cavity is guided to the internal bore (7) of the turbine disc (5). Said means comprise a diffuser (52) co-operating with the injection means (17) and an annular cover intended to co-operate with the diffuser (52) and to cover first attachment members (27) arranged in the cavity (16), the diffuser (52) and the cover (40) being configured such that the cooling air at the outlet of the diffuser (52) is at least partially guided by a radially outer surface (40a) of the cover (40).
F01D 5/08 - Heating, heat-insulating, or cooling means
F01D 21/14 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to other specific conditions
The invention relates to a system (100) which comprises: a gas turbine (102) designed so that a combustion chamber (110) can be ignited in a first ignition range of rotational speeds of a compressor shaft (112); and a control device (106) designed to control an electric starter (114) to accelerate the compressor shaft (112), and, when the compressor shaft (112) is accelerated, to control an attempt to ignite the combustion chamber (110). The gas turbine (102) is designed so that the combustion chamber (110) can be ignited in a second ignition range which is higher than the first ignition range, but not between these two ignition ranges, and the ignition attempt is carried out in the second ignition range.
A rotor wheel (10) for an aircraft turbine engine, said wheel comprising: - a disc (12) having a main axis (A) and provided with cells (14) at its outer periphery, the cells extending along the axis and each comprising a bottom (14a) and two side flanks (14b), - vanes (22) mounted in the cells (14) of the disc (12), each of said vanes comprising a blade (24) connected by a platform (26) to a root (28) which is configured to be interlockingly mounted in one of the cells (14), the root (28) of each of the vanes (22) comprising, at its radially inner end, a lobe (30) with a first axial end which comprises a circumferential notch (32) and a second axial end, opposite the first end, which comprises a radially inward facing stop (33) configured to axially bear on a first face (12b) of the disc (12), and - a split annular ring (34) engaged in the notches (32) of the vanes (22) and axially clamped against a second face (12a) of the disc, the second face (12a) being opposite the first face (12b), the lobe (30) of the root (28) of each of the vanes (22) comprising, between the first and second ends, a radially inward facing projecting bulb (42) configured to radially bear on the surface of the bottom (14a) of the corresponding cell (14).
The invention relates to a turbomachine comprising a rotary body comprising a motor shaft supplying mechanical power, and at least one magnetic drive pump comprising at least: one stator (110) delimiting an annular inner space and comprising a first and a second flange (1110, 1120), a rotor (120) arranged in the inner space between the first and second flanges and capable of driving fluid, the rotor being able to rotate about an axis of rotation (A), a pair of magnets (1241, 1242) having opposite polarities coaxially arranged on the rotor with the axis of rotation, a magnet (1116) arranged on the first flange (1110) in order to co-operate with one of the magnets of the pair of magnets of the rotor, a magnetic means (130) for rotating the rotor arranged on the second flange (1120), the second flange being non-magnetic.
A device (110) for guiding a main air flow (F1) for an aircraft turbine engine (22), said device comprising: a first air flowing pipe (112) of a main air flow (F1), said first pipe having a main axis (A), - a plurality of ejectors (114) of a secondary air flow (F2) located within the first pipe (112) and configured to eject a secondary air flow (F2) and force the flow of the main air flow (F1) into said first pipe (112), the ejectors (114) being distributed around the main axis (A), and - a second air flowing pipe (116) located at the outlet of the ejectors (114) and comprising one end (116a) which is connected to one end (112a) of the first pipe (112), characterised in that the second pipe (116) comprises a narrow end (134).
F02C 7/052 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with dust-separation devices
F04F 5/16 - Jet pumps, i.e. devices in which fluid flow is induced by pressure drop caused by velocity of another fluid flow the inducing fluid being elastic fluid displacing elastic fluids
The invention relates to a transmission device (1) for hybrid aircraft comprising a turboshaft engine having a gas generator (12), a free turbine (14), and a main rotor (52), the device comprising a first reversible electric motor (30) coupled to a shaft (16) of the free turbine (14) via a first disengageable coupling means (32), and to the main rotor (52), and a second reversible electric motor (40) coupled to a shaft (18) of the gas generator (12) via a second disengageable coupling means (44), and coupled to the main rotor (52) via a third disengageable coupling means (42), the second disengageable coupling means (44) being operable when the second electric motor (40) rotates in a first direction of rotation, and the third disengageable coupling means (42) being operable when the second electric motor (40) rotates in a second direction of rotation opposite to the first direction of rotation.
B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type
B60K 6/50 - Architecture of the driveline characterised by arrangement or kind of transmission units
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
B64D 35/02 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the type of power plant
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
The system comprises: - a fuel tank; - a suction pipe that is connected to the tank and is located higher than the tank; - an electric pump (36); - a supply pump configured to be mechanically driven by an accessory gear box and be connected, at the outlet, to a fuel supply circuit for the engine; and - an air evacuation drain (26). The electric pump (36) is in communication with the suction pipe independently of the supply pump, and with the drain. The supply pump is in communication with the suction pipe independently of the electric pump.
The present invention relates to a device (1) for controlling an electric aircraft-propelling assembly, said propelling assembly comprising a propeller (3) and at least one electric motor (4) that is powered by an electric supply voltage and that delivers a torque and a rotation speed to drive the propeller (3). The control device (1) comprises at least a unit (11) for measuring an electric supply voltage, and a control unit (12) suitable for making a signal delivered to the electric motor vary as a function of said electric supply voltage, with a view to making the rotation speed of the propeller vary.
The invention relates to a turbomachine comprising a gas generator (13) equipped with a first shaft (18), an accessories box (14), at least one reversible electrical machine (11) coupled with the accessories box (14), a free turbine (12) equipped with a second shaft (17) and driven in rotation by a gas flow from the gas generator (13), and at least one accessory (15, 16) coupled to the accessories box (14). The turbomachine (10) comprises a first mechanical coupling means (20) configured to mechanically couple the first mechanical shaft (18) to the accessories box (14) in a first configuration and to mechanically uncouple said first mechanical shaft (18) from the accessories box (14) in a second configuration, and a second mechanical coupling means (25) configured to mechanically couple the second mechanical shaft (17) to the accessories box (14) in a first configuration and to mechanically uncouple said second mechanical shaft (17) from the accessories box (14) in a second configuration, the electrical machine (11) being sized to drive the gas generator and the at least one accessory when the turbomachine is started.
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 6/20 - Adaptations of gas-turbine plants for driving vehicles
A fixed-wing combat aircraft (10) comprising an electrical power source (20), a propulsion system (30), a low-power non-propulsion assembly (40) comprising a flight control system, a high-power non-propulsion assembly (50) comprising an electrical weapon system (52), and a management unit (60) configured to selectively establish on command multiple operating modes comprising: a flight mode, in which the management unit distributes the electrical power supplied by the electrical power source to the propulsion system and to the low-power non-propulsion assembly, and an attack mode, in which the management unit limits the electrical power supplied by the electrical power source to the propulsion system and to the low-power non-propulsion assembly to the power required to allow the aircraft to glide, and reserves a majority of the available electrical power for the high-power non-propulsion assembly.
B64D 7/00 - Arrangement of military equipment, e.g. armaments, armament accessories, or military shielding, in aircraft; Adaptations of armament mountings for aircraft
F41H 13/00 - Means of attack or defence not otherwise provided for
34.
SYSTEM FOR PUMPING AND METERING A FLUID FOR A TURBINE ENGINE AND METHOD FOR CONTROLLING SUCH A SYSTEM
The invention relates to a system (400) for pumping and metering a fluid (10) for a turbine engine, which system comprises at least one pump (412, 414) for the fluid and an electronic computer (430) configured to determine the flow rate of the fluid to be delivered to the turbine engine, the pumping and metering system being characterised in that it further comprises a first electric motor and a second electric motor, which are each configured to drive the at least one pump, and in that the electronic computer comprises a first control loop (432) for controlling at least the first electric motor and a second control loop (434) for controlling at least the second electric motor.
A turbine (1) for a turbomachine of longitudinal axis (X), comprising: an alternating arrangement of annular rows of movable blades (64) and of fixed blades (65) and a radially inner annular cavity (68) formed radially inside the movable and fixed blades (64, 65), and a supply circuit (32) for supplying cooling air to the inner annular cavity (68), the downstream end of the supply circuit (32) comprising an inner annular row of orifices (81) and an outer annular row of orifices (82) opening into the radially inner annular cavity (68), the turbine further comprising means (85) for controlling the flow rate of supply air to the orifices of the inner and outer annular rows of orifices (81, 82).
Disclosed is a rotor (121) for a centrifugal breather for an air/oil mixture of a turbomachine, this rotor comprising: a hollow shaft (114) extending along an axis (X), a pinion (115) for rotating the hollow shaft (114), this pinion (115) extending around the axis (X) and being formed of a single part and in a first material with at least one first portion (114a) of the hollow shaft (114), and an annular structure (117) extending around the axis (X) and constrained to rotate with the shaft (114), this structure (117) being produced in a second material, different from the first material, characterised in that said structure (117) is made integral with the shaft (114) by additive manufacturing of this structure directly on at least one annular surface (122) of the pinion (115) which forms at least one annular support surface for this additive manufacturing.
F01M 11/08 - Separating lubricant from air or fuel-air mixture before entry into cylinder
B01D 39/20 - Other self-supporting filtering material of inorganic material, e.g. asbestos paper or metallic filtering material of non-woven wires
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 5/08 - Centrifuges for separating predominantly gaseous mixtures
Ignition method for a continuous combustion engine comprising an electronic engine control member (7), a high-power housing (2), a spark plug ignition circuit and a solenoid fuel valve engaging with a starter, the method being implemented by the electronic engine control member (7) and comprising a pre-charging step (03) of pre-charging the high-power housing (2) prior to an engine-starting step (E3) initiated upon an instruction to start the engine (02), the pre-charging step (03) being controlled: - by switching on the electronic engine control member (7), - or by putting the engine in idle mode (OV).
A double wall (2) for an aircraft gas turbine combustion chamber (1) comprising an internal wall (21) which is configured to be in contact with the combustion reaction, and an external wall (22) which is at a distance from the internal wall (21), comprising a plurality of openings (3) so as to allow the circulation of cooling air streams (F), outside the external wall (22), which cool the internal wall (21), the internal wall (21) being free of perforations so as to prevent any circulation of a cooling air stream (F) towards the centre of the combustion chamber (1). The double wall (2) is characterised in that the internal wall (21) comprises a plurality of members (4) projecting towards the external wall (22), each projecting member (4) comprising a foot portion (4a) and a cylindrical head portion (4b) with a circular cross-section, the head portion (4b) extending into an opening (3) with a circular cross-section so as to define a calibrated cross-sectional area between the projecting member (4) and the opening (3), through which area a cooling air stream (F) can flow.
The invention relates to an assembly (1) for transmitting and measuring a torque transmitted between a first member and a second member (16) of a turbomachine, comprising a pinion (2) which includes a first annular portion (5) and a second annular portion (6) joining together at a connecting portion (7) supporting a gear system (4) of the pinion (2), the first and second annular portions (5, 6) extending axially along an axis of rotation (X) of the pinion (2) in directions (D1, D2) opposite from the connecting portion (7), the rotation of at least one of the the annular portions (5, 6) being guided via a bearing (8, 9) for guiding the pinion (2), the pinion further comprising an axially extending power shaft (10), the power shaft (10) comprising a first area (11) rotatably coupled to the first annular portion (5) of the pinion (2), and a second area (12) intended to be rotatably coupled to the second member (16) of the turbomachine, the first area being axially spaced apart from the second area, and means for measuring the torsion of the power shaft (10) which are arranged in a third area of the power shaft (10) located between the first and second areas (11, 12), the measuring means comprising at least one tone wheel (20, 25) and acquisition means capable of engaging with the tone wheel (20, 25), characterised in that the measuring means comprise a first tone wheel (20) provided in the third area of the power shaft (10) and a second tone wheel (25) provided in the second annular portion (6) of the pinion (2) and located axially opposite the first tone wheel (20), the acquisition means being arranged axially opposite the first tone wheel (20) and the second tone wheel (25) and being configured to supply a signal representative of an angular variation between the first tone wheel (20) and the second tone wheel (25).
G01L 3/10 - Rotary-transmission dynamometers wherein the torque-transmitting element comprises a torsionally-flexible shaft involving electric or magnetic means for indicating
Disclosed is a propulsion system (2) for a helicopter (1) comprising a main engine (9), a main rotor (3), a main gearbox (4) comprising 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 function as an electric generator in order 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 electricity by the first electric machine and configured to function as an electric motor in order to deliver additional mechanical power to the main gearbox (4).
ref1ref223211) and, - controlling the stopping of the reversible electric machine (7) in order to no longer drive the power shaft (3, 12), in order to cause a progressive stopping (E9, E10) of the rotation of the gas turbine (2).
The invention relates to a 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 comprising a step of acquiring (ET1) measurements (Pv) of power parameters (PARi), a step of comparing (ET2) each measurement (Pv) with at least one limitation threshold, so as to deduce therefrom at least one gross margin (Pmb), a step of converting (ET3) the gross margins (Pmb) into refined margins (Pma) expressed according to the same common magnitude, a step of transposing (ET4) into standardised margins (Pmn) at least at one reference point, a step of determining (ET5) a source power margin (MIndS) and a consumer power margin (MIndC) at said reference point and a step of determining (ET6) the minimum power margin (MInd) by selecting the lowest power margin.
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
B64D 43/00 - Arrangements or adaptations of instruments
B64D 27/02 - Aircraft characterised by the type or position of power plant
43.
DEVICE FOR CONTROLLING AN AIRFLOW GUIDING SYSTEM, IN PARTICULAR IN AN AIRCRAFT TURBINE ENGINE
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.
The invention relates to a turbomachine (10) comprising a propeller (12), a propeller shaft (13) carrying the propeller (12), a rotating electric machine (19), having at least a first configuration in which it is mechanically coupled to the propeller shaft (13), and a motor oil pump (21) supplying a lubricating circuit (20) of the turbomachine (10). The rotating electric machine (19) in the first configuration is mechanically coupled to the motor oil pump (21) in such a way that the rotating electric machine (19) additionally drives the motor oil pump (21) when it is supplied with current. The invention further relates to an aircraft comprising such a turbomachine (10).
A turbomachine (10), in particular for a rotary-wing aircraft, having a gas generator (13) provided with a rotary shaft (18), a reversible first electric machine (11), a free turbine (12) driven in rotation by a flow of gas generated by the gas generator (13), at least one accessory (15, 16) selected from an oil pump and a fuel pump, an accessories gearbox (14) comprising a gear train configured to drive said at least one accessory (15, 16), and a second electric machine (19). The second electric machine (19) is reversible, said first electric machine (11) is mechanically coupled to the gas generator (13), the accessories gearbox (14) and the second electric machine (19) are mechanically coupled to the free turbine (12), and the turbomachine (10) does not have a kinematic link between the gear train of the accessories gearbox (14) and the shaft (18) of the gas generator (13).
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 7/32 - Arrangement, mounting, or driving, of auxiliaries
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
The invention relates to a method for determining an efficiency fault (R11-R15) of at least one module (11-15) of a turboshaft engine (T) of an aircraft (A), the method for determining comprising: • A step of determining an estimated real mapping (CARE), • A step of determining real indicators (IRE) from the estimated real mapping (CARE), • A step of determining (E3) a plurality of simulated mappings from a simulation of a theoretical model of the turboshaft engine (T) for different efficiency configurations, • A step of determining (E4) simulated indicators (ISx) for each simulated mapping (CARSx), • A step of training (E5) a mathematical model (CLASS) by coupling the simulated indicators (ISx) with efficiency configurations (CR), and • A step of applying (E6) said mathematical model (CLASS) to the real indicators (IRE) so as to deduce therefrom a real efficiency configuration (CR).
Disclosed is a turbomachine comprising a gas generator (13) provided 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.
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
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 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
Disclosed is a turbogenerator (10), in particular for an electrically-driven rotary wing aircraft, comprising a gas generator (13) provided with a first shaft (18), at least one reversible electrical machine (11), and a free turbine (12) provided with a second shaft (17) and caused to rotate by a gas flow generated by the gas generator (13). The second shaft (17) is coupled to the at least one electrical machine (11) during all phases of operation of the turbogenerator (10), and the turbogenerator (10) further comprises a single mechanical coupling means (20) for coupling said first mechanical shaft (18) to said second mechanical shaft (17) when the electrical machine (11) is operating in motor mode and mechanically uncoupling said first mechanical shaft (18) from said second mechanical shaft (17) when the electrical machine (11) is operating in generator mode.
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
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
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
Hybrid turbomachine (1) comprising an electrical generator (2), a gas generator (3) provided with an air intake (31) and an exhaust (32) and an acoustic control system (4) comprising a control unit (41) and a plurality of loudspeakers. At least a first loudspeaker (43) is arranged on the electrical generator (2) and/or at least a second loudspeaker (45) is arranged at the air intake (31) of the gas generator (3) and/or at least a third loudspeaker (47) is arranged at the exhaust (32) of the gas generator (3). The control unit (41) of the acoustic control system (4) is mounted on the electrical generator (2) and is configured to perform an AC/DC electrical conversion of the electromotive force of the electrical generator (2) into an adjustable direct-current voltage which is intended to be distributed to loads or to energy storage means.
Disclosed is a method for protecting windings from excessive heating in an aircraft electrical machine comprising a stator (12) and a rotor (14) configured to be driven in rotation with respect to one another, the stator comprising a plurality of notches (120) receiving the same or different plurality of windings, the method comprising the following successive steps: inserting an electrical insulator (16) into the notches or on the teeth of the stator, positioning the windings (18) in the notches or on the teeth of the stator, casting a phase-change material (20) into the notches or onto the teeth provided with the windings, the electrical insulator forming a casting mould.
H02K 15/10 - Applying solid insulation to windings, stators or rotors
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
C09K 5/06 - Materials undergoing a change of physical state when used the change of state being from liquid to solid or vice-versa
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
An aspect of the invention relates to a start-up method (100) for an aircraft engine (1), wherein the engine (1) is connected to a lubrication circuit (6) comprising, in particular, an oil pump system (7), the lubrication circuit (6) being constructed and arranged to circulate oil in the engine (1), and wherein an operating mode of the engine (1) comprises a stop mode and a standby mode, the start-up method (100) being characterised in that it comprises, during a start-up phase, the steps of: - measuring (101) an oil temperature, the measuring step (101) being carried out by a temperature detection device; - as a function of the temperature measured compared with a threshold temperature, and as a function of the operating mode of the engine, selecting (102) an oil flow profile for start-up to be applied in the engine, the step of selecting (102) being carried out by a computer; - applying (103) the oil flow profile for start-up selected by means of the oil pump system, the oil pump system being controlled by the computer.
The invention relates to 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 comprising an annular peripheral module (3) and a circular central module (2).
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
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B22F 10/37 - Process control of powder bed aspects, e.g. density
B22F 10/47 - Structures for supporting workpieces or articles during manufacture and removed afterwards characterised by structural features
B22F 10/60 - Treatment of workpieces or articles after build-up
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
53.
PROPULSION UNIT WITH FOLDABLE PROPELLER BLADES AND METHOD FOR FOLDING THE BLADES
The invention relates to a propulsion unit (22) with a propeller (26), comprising: - a nacelle (24); - a propeller (26) rotatably mounted in the nacelle (24) by means of a hub (28), the propeller (26) comprising blades (32) mounted in a blade cuff (38) pivotable about a pitch axis (Y) relative to the hub (28), each blade (32) being pivotable relative to the cuff (38) about a folding axis (Z); - a folding device (50) that comprises an actuator (52) for folding the blades (32); the propulsion unit (22) being characterised in that the folding device (50) comprises a control member (56) rotationally affixed to the blade cuff (38) and driven by the actuator (52), and a connecting rod (58) pivotably mounted, on the one hand, on a root (34) of the associated blade (32) and, on the other hand, on the movable control member (56).
Aircraft turbomachine (10), comprising: a centrifugal compressor (14), a combustion chamber (24), the combustion chamber being supplied by the compressor via a diffuser (34) and via a straightener (36), and a heat exchanger (38), the exchanger comprising: a first circuit (38a), supplied with exhaust gas from the turbomachine, and a second circuit (38b), which are connected by volutes (40a, 40b) on the one hand to an outlet (34b) of the diffuser (34) and on the other hand to an inlet (36a) of the straightener (36), the volutes (40a, 40b) having reversed winding directions such that their connection ports to the exchanger (38) are independent of one another and are substantially diametrically opposed, and such that the minimum cross section (S2) 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
The invention proposes that bridges (9) extending adjacent to each other integrally link a radially inner wall (5) and a radially outer wall (3) of a combustion chamber of a gas turbine engine towards a free end (5a) of the radially inner wall (5), the whole being an additive layer structure.
The invention relates to a propulsion unit (22) having a propeller (26) for an aircraft (20), comprising: - a nacelle (24); - a propeller (26) mounted in the nacelle (24) so as to be capable of rotating about a longitudinal axis (X) of rotation, the propeller (26) having blades (32) mounted by a root (34) so as to be capable of pivoting between a deployed position, in which they extend radially relative to the axis (X) of rotation, and a folded position, in which they are longitudinally received against the nacelle (24); - drive means that rotate the propeller (26); indexing means for stopping the propeller (26) in at least one indexed angular position (θ i) relative to the nacelle (24); the propulsion unit (22) being characterised in that the indexing means consist of a stepping electric motor (82) comprising a rotor (84) that is coupled to the propeller (26).
ELECTROMECHANICAL INSTALLATION FOR AN AIRCRAFT WITH A TURBOGENERATOR, METHOD FOR EMERGENCY SHUTDOWN OF AN AIRCRAFT TURBOGENERATOR AND CORRESPONDING COMPUTER PROGRAM
This electromechanical installation (100) for an aircraft comprises: - an electrical network (102) comprising electrical subnetworks (104, 106); - a turbogenerator (118) comprising a gas turbine (120), an electricity generator (122) with permanent magnets having phase groups (126, 128) respectively connected to the electrical subnetworks (104, 106), and, for each phase group (126, 128), an isolation device (134, 136); and - a control device (140) 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 (126, 128) from its associated electrical subnetwork (104, 106) and to command the shutdown of the gas turbine (120). The control device (140) is also designed, in response to the detection of the short circuit, to keep each healthy phase group (126, 128) connected to its electrical subnetwork (104, 106).
H02J 4/00 - Circuit arrangements for mains or distribution networks not specified as ac or dc
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
H02P 3/00 - Arrangements for stopping or slowing electric motors, generators, or dynamo-electric converters
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
58.
PROPULSION UNIT WITH FOLDABLE PROPELLER BLADES AND METHOD FOR FOLDING THE BLADES
The invention relates to a propulsion unit (22) with a propeller (26), comprising: - a nacelle (24); - a propeller (26) rotatably mounted in the nacelle (24) by means of a hub (28), the propeller (26) comprising blades (32) mounted in a blade cuff (38) pivotable about a pitch axis (Y) relative to the hub (28), each blade (32) being pivotable relative to the cuff (38) about a folding axis (Z); - a folding device (50) that comprises an actuator (52) for folding the blades (32); the propulsion unit being characterised in that the folding device (50) comprises a control member (56) rotationally affixed to the blade cuff (38) and driven by the actuator (52), and a connecting rod (58) pivotably mounted, on the one hand, on a root (34) of the associated blade (32) and, on the other hand, on the movable control member (56).
Aircraft turbomachine (11), comprising: - a centrifugal compressor (12), - an annular combustion chamber (24), - an annular casing (29) extending around the chamber and delimiting an annular space (E) in which the chamber is situated, and - a heat exchanger (38), this exchanger comprising a first circuit (38a) supplied with exhaust gas from the turbomachine, and a second circuit (38b) connected by volutes (40a, 40b) respectively to an outlet (34b) of the compressor (12) and to said space (E), characterized in that the volutes (40a, 40b) are at an axial distance from one another, and in that the second volute (40b) is connected to the space (E) by a straightener (42) which is situated at least in part outside the casing (29) and which is integrated into an annular connecting pipe (44) which connects the second volute (40b) to this casing (29).
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
The invention relates to a hydraulic torque measurement device for an aircraft engine unit, comprising: a hermetically sealed enclosure (20a) comprising a rigid wall (21) and a deformable wall (22) defining an internal volume (26), and at least one pressure measuring element (28) capable of measuring the pressure prevailing in the internal volume (26), the hydraulic device (20) allowing a service state in which the internal volume (26) is filled exclusively with an incompressible liquid, and in which the deformable wall (22) is configured to be subjected to a pressure force. The invention further relates to a gearbox comprising such a hydraulic device.
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
G01L 3/14 - Rotary-transmission dynamometers wherein the torque-transmitting element is other than a torsionally-flexible shaft
61.
TURBOMACHINE COMPRESSOR HAVING A STATIONARY WALL PROVIDED WITH A SHAPE TREATMENT
The invention relates to a turbomachine having a compressor comprising variable pitch stator vanes (11) each extending radially between a rotating hub (6) and a stationary casing (3) surrounding this rotating hub (6), each variable pitch vane (11) comprising a blade (12) having a base (14) spaced by a first radial clearance (J1) from a stationary wall (2) of the casing (3), and a tip (17) spaced by a second radial clearance (J2) from a rotating wall (4) of the rotating hub (6). The stationary wall of the casing (3) or the rotating wall (4) of the rotating hub (6) has a shape treatment opposite the blade (12) for channeling an air leak passing through the corresponding clearance.
The invention relates to an oro-nasal respiratory mask (1) comprising an airtight body (2) comprising an edge (3) intended to be sealably applied onto the face of a user, the body (2) comprising at least one opening (4) allowing the passage of an inhalation air flow and an exhalation air flow, a filter (5) being detachably and sealably mounted in the opening (4) of the body (2) so that the inhalation and exhalation air flows fully pass through said filter (5).
The invention relates to a modular tray (12) for the additive manufacturing of a part with an axis of revolution on a powder bed, characterised in that it comprises: - a circular shaft module (13) comprising a shaft provided with a circular tray at one end thereof, the shaft and the circular tray being concentric; and - a main support module (16) having a cavity (17) in one face configured to receive the circular shaft module (13), the shaft being inserted completely into the cavity; the assembly of the circular shaft module and the main support module defining a planar upper surface which is formed at least in part by the circular tray of the circular shaft module. A modular tray (12) for the additive manufacturing of a part with an axis of revolution on a powder bed, characterised in that it comprises: a circular shaft module (13) comprising a shaft provided with a circular tray at one end thereof, the shaft and the circular tray being concentric; and a main support module (16) having a cavity (17) in one face configured to receive the circular shaft module (13), the shaft being inserted completely into the cavity; the assembly of the circular shaft module and the main support module defining a planar upper surface which is formed at least in part by the circular tray of the circular shaft module.
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
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 80/00 - Products made by additive manufacturing
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
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).
The invention relates to a combustion assembly for a gas turbine, comprising a flame tube (14) and fuel supply means having a flow limiter (18) which supplies an injector (19), the flow limiter (18) and the fuel injector (19) being formed in a single piece.
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 inner 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).
The invention relates to a fuel supply circuit of an aircraft engine, comprising a centrifugal pump which is mechanically coupled to a drive shaft (18) delivering a mechanical power. The circuit further comprises at least one electromagnetic pump (100) having at least one stator (110) which delimits an annular internal volume, in which a rotor (120) capable of driving a fluid is present, a plurality of magnets (130) which are distributed annularly around the rotor (120), and at least a plurality of coils (140) which are distributed annularly inside the stator (110) opposite the magnets (130). The rotor (120) is connected to the drive shaft (18) by means of a unidirectional clutch element (150).
Disclosed is a method for manufacturing a vane (10) comprising a first portion and a second portion, the method comprising a step of forming the first portion (E1) that comprises forming a model of the first portion from removable material, then forming a first shell mould from the model of the first portion, then forming the single-crystal or columnar first portion from a first metal alloy in the first shell mould from a single-crystal seed, a step of forming the second portion (E2) in which a 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. Also disclosed is a vane (10) comprising a single-crystal or columnar first portion made from a first metal alloy and a polycrystalline second portion made from a second metal alloy different from the first metal alloy.
B22C 9/24 - Moulds for peculiarly-shaped castings for hollow articles
B22D 19/04 - Casting in, on, or around, objects which form part of the product for joining parts
B22D 19/16 - Casting in, on, or around, objects which form part of the product for making compound objects cast of two or more different metals, e.g. for making rolls for rolling mills
B22D 21/02 - Casting exceedingly oxidisable non-ferrous metals, e.g. in inert atmosphere
B22D 27/04 - Influencing the temperature of the metal, e.g. by heating or cooling the mould
C30B 11/00 - Single-crystal-growth by normal freezing or freezing under temperature gradient, e.g. Bridgman- Stockbarger method
C30B 29/60 - Single crystals or homogeneous polycrystalline material with defined structure characterised by the material or by their shape characterised by shape
C30B 29/66 - Crystals of complex geometrical shape, e.g. tubes, cylinders
69.
THREE-POINT CONTACT ROLLING BEARING WITH IMPROVED DRAIN
Three-point contact rolling bearing (1) wherein the internal raceway (12), the third contact point (P3) being situated in a sector of the external raceway (22) delimited by a proximal end of the external raceway (22) on one side, and by the rolling plane (Pr) on the other side, characterized in that the external ring (20) comprises a drain (50) that is provided in the external raceway (22) and opens out at an external surface of the external ring (20), said drain (50) opening onto the external raceway (22) in an off-centre manner with respect to the rolling plane (Pr), in a sector of the external raceway (22) that is delimited by the rolling plane (Pr) on one side, and by a distal end of the external raceway (22) on the other side, the drain (50) and the third contact point (P3) being separated from one another.
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
F16C 33/66 - Special parts or details in view of lubrication
70.
TURBOGENERATOR FOR AIRCRAFT, COMPRISING AN IMPROVED OIL SYSTEM
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).
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
B01D 45/08 - Separating dispersed particles from gases or vapours by gravity, inertia, or centrifugal forces by utilising inertia by impingement against baffle separators
71.
SYSTEM AND METHOD FOR BALANCING AT LEAST ONE PARAMETER TO BE BALANCED OF AN ELECTRIC MOTOR OF A PROPULSION SYSTEM
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, comprising 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 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
B64D 31/12 - Initiating means actuated automatically for equalising or synchronising power plants
B64C 11/00 - Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
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
H02P 5/68 - Arrangements specially adapted for regulating or controlling the speed or torque of two or more electric motors controlling two or more dc dynamo-electric motors
72.
DEVICE FOR SUPPLYING FUEL TO A COMBUSTION CHAMBER OF A GAS GENERATOR
The document relates to a device for supplying fuel to a combustion chamber of a gas generator, comprising: - a wheel (14) for injecting fuel into the combustion chamber (18), - a fuel supply rail (20) comprising an internal fuel circuit (30) which comprises fuel outlet means (32) supplying fuel to an annular spray chamber (24) formed between the ramp (28, 44, 48, 52, 56) and the injection wheel (14), - at least one dynamic annular seal (26) designed to provide a seal between an annular face (34) of the fuel supply rail (28, 44, 48, 52, 56) and the injection wheel (14), characterised in that the fuel circuit (30) of the fuel supply rail comprises an annular fuel flow portion arranged radially in the dynamic annular seal (26).
F02C 3/16 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant the combustion chambers being formed at least partly in the turbine rotor
F02C 7/224 - Heating fuel before feeding to the burner
F23R 3/38 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising rotary fuel injection means
73.
TURBOMACHINE PROVIDED WITH AN ELECTROMAGNETIC PUMP WITH AXIAL MAGNETIC FLUX
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
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
F04C 7/00 - Rotary-piston machines or pumps with fluid ring or the like
F04C 15/00 - Component parts, details or accessories of machines, pumps or pumping installations, not provided for in groups
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 said 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
The invention relates to a propulsion system (1, 1') for an aircraft, comprising a rotor (2) and a nacelle fairing (3) that extends around said rotor in relation to an axis (X) and includes an upstream portion forming an inlet section of the nacelle fairing 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) includes 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 wall (20b) has piston actuator mechanisms (23, 23'), each actuator mechanism being actuatable independently of the others and being designed to cooperate with means (24, 24') built into an inner surface (20b') of the wall (20b) so as to deform the wall (20b) in a radial direction in relation to said axis (X) under the effect of a predetermined displacement command.
The invention relates to a propulsion system (1, 1') for an aircraft, comprising a rotor (2) and a nacelle fairing (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.
Disclosed is a method for quickly stopping the propulsion rotor of a helicopter after landing, comprising, following a request (100) by a pilot of the helicopter to quickly stop the engine, the following steps managed by the control unit of the turbomachine: - detecting (110) the absence of a thermal stabilisation phase of the gas generator of at least one turbomachine, - controlling (130) the extinguishing of the combustion chamber of the gas generator of at least one turbomachine, - maintaining the rotation of the gas generator (140) whose combustion chamber has been extinguished by means of the at least one electrical machine in order to ventilate the gas generator, and - stopping (160) 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
The invention relates to a propulsion system (1, 1') for an aircraft, in particular a vertical take-off and landing (VTOL) aircraft, comprising a rotor (2) and a nacelle fairing (3) that extends around said rotor in relation to an axis (X) and includes an upstream portion forming an inlet section of the nacelle fairing 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), at least one part of one of said walls being made of a deformable shape memory material, said wall being provided with heating elements (23a, 23c) extending in different consecutive angular sectors about said axis (X), each heating element being independently actuatable and being designed to deform said wall in a direction that extends radially in relation to said axis (X) and is angularly centered in relation to its angular sector, under the effect of a predetermined voltage command.
F02K 1/10 - Varying effective area of jet pipe or nozzle by distorting the jet pipe or nozzle
B64D 27/16 - Aircraft characterised by the type or position of power plant of jet type
B64C 15/02 - Attitude, flight direction or altitude control by jet reaction the jets being propulsion jets
B64C 11/00 - Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
B64C 27/20 - Rotorcraft characterised by having shrouded rotors, e.g. flying platforms
B64D 33/04 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
F02K 1/34 - Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto using fluid jets to influence the jet flow for attenuating noise
F02K 1/00 - Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
The invention relates to a propulsion system (1) for a helicopter, comprising a turbine engine (2) with a coupled turbine and an electric machine (3) able to operate as an electric motor, the turbine engine (2) and the electric machine (3) being able to rotate at least one main rotor (5) intended to be coupled to a rotary wing (6), characterized in that said system comprises means (14) for rotational coupling and decoupling between a rotor (3a) of the electric machine (3) and a rotor (2a) of the turbine engine (2), the coupling and decoupling means (14) being able to allow the rotor (2a) of the turbine engine (2) to be rotated by means of the electric machine (3), in a first state of the propulsion system (1), and being able to allow the rotor (2a) of the turbine engine (2) and the rotor (3a) of the electric machine (3) to be rotationally decoupled, 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
The present invention relates to a pre-vaporisation tube (4) for a turbine engine combustion chamber (1), comprising: a main body (40) defining a first inner duct (41) in which an injector (3) is intended to be mounted, and comprising a first end to be attached (40a) to a wall (13) of the chamber, at least two end pieces (42) arranged at a second end (40b) of the body (40) and defining second inner ducts (43), the end pieces comprising first portions (44) and second portions (45), respectively; the second portions (45) each comprise two coaxial cylindrical walls which are inner (452) and outer (454) coaxial cylindrical walls, respectively, which define an annular cavity (450) therebetween, and the inner wall (452) defines an inner passage and comprises first openings (50) for fluid communication between said passage and the annular cavity (450).
F23R 3/32 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices being tubular
81.
BLADE FOR A TURBINE ENGINE, AND ASSOCIATED TURBINE ENGINE
A blade (10) for a turbine engine comprises an aerodynamic airfoil (12) extending radially outwards, a blade tip (14), a blade root (16), and a platform (18) connected to the root (16) by an upper end (18a), the maximum radial distance (H) between the blade tip (14) and the platform (18) defining a maximum extent of the bearing surface of the airfoil (12). The blade (10) is partially covered by a protective strip (30) to combat oxidation and corrosion, said strip (30) extending between a lower limit (L1) situated on at least one part of the platform (18) and an upper limit (L2) situated on the blade root (16). The blade tip (14) is not covered by the protective strip (30).
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 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 11/12 - 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 5/20 - Specially-shaped blade tips to seal space between tips and stator
83.
RING FOR A TURBOMACHINE OR TURBOSHAFT ENGINE TURBINE
The invention relates to a ring (1) for a turbomachine turbine, intended to surround a bladed disc (2) of a turbine rotor, said ring (1) extending circumferentially about an axis and comprising a radially outer annular support part (9) and a part (10) delimiting a radially inner gas-flow circulation duct (6) and comprising a plurality of angular segments (13) distributed over the periphery and situated adjacently with respect to one another so as to form an annular part delimiting the duct (6), circumferential clearances (j) being formed between the circumferential ends of the adjacent segments (13) situated facing one another, each segment (13) being connected to the support part (9) via a connecting region (14), an annular cooling-fluid circulation passage (15) being delimited radially between the outer support part (9) and the inner part (10) delimiting the duct.
The invention relates to an aeronautical turbogenerator for hybrid electric propulsion comprising a heat engine and an electric generator (10) mechanically coupled to the heat engine and comprising a rotor and a stator, the rotor extending in an axial direction and comprising a common magnetised rotor yoke (100) comprising a plurality of permanent magnets (102) defining at least three axially distributed mobile annular rings (102A, 102B, 102C), the stator comprising a magnetic stator yoke (200) comprising a plurality of electrical windings (202) defining stationary sectors (202A, 202B, 202C) distributed axially and/or circumferentially, at least two stationary sectors, one of which axially covers at least two mobile annular rings, being arranged angularly so as not to coincide with one another and thus to output at least two separate, independent voltage levels.
H02K 16/00 - Machines with more than one rotor or stator
H02K 21/28 - Synchronous motors having permanent magnets; Synchronous generators having permanent magnets with rotating armatures and stationary magnets with armatures rotating within the magnets
85.
FUEL INJECTOR WITH A PURGE CIRCUIT FOR AN AIRCRAFT TURBINE ENGINE
Fuel injector (110) for an aircraft turbine engine, comprising a tubular body (112) having an axis of elongation (A) and comprising a first longitudinal end (114) for supplying with fuel and a second longitudinal end (116) for ejecting a jet of fuel, said body further comprising an integrated purge-air circuit which comprises an internal cavity (122) which is connected to air inlet orifices (124) situated on the body and to at least one air outlet situated at said second end, characterized in that air-flow disruptors (150, 156) are provided, projecting into said cavity.
A hybrid propulsion chain (1) for an aircraft, the hybrid propulsion chain (1) comprising a plurality of propulsion rotors (R1-RX) connected to an electrical distribution module (10) by a plurality of electrical connections (E1-E4), the electrical distribution module (10) being connected, on the one hand, to a non-propulsion turbine engine (T) via an electrical generation system (20) and, on the other hand, to an electric battery (BAT), each propulsion rotor (R1-RX) 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 (1) comprising an auxiliary mechanical drive system (30) mechanically connected to the non-propulsion turbine engine (T), the auxiliary mechanical drive system (30) comprising a plurality of mechanical connections (M1-M4) for mechanically rotating at least one rotor shaft (32) of each propulsion rotor (R1-RX).
B64D 27/02 - Aircraft characterised by the type or position of power plant
B64D 27/10 - Aircraft characterised by the type or position of power plant of gas-turbine type
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
87.
TURBOGENERATOR WITH SIMPLIFIED REGULATING SYSTEM FOR AN AIRCRAFT
The present invention relates to a propulsion assembly for an aircraft, comprising a coupled-turbine turbomachine comprising a combustion chamber and a shaft that is mounted so as to rotate and is designed to rotate at a turbomachine rotational speed; an electrical generator (2) coupled to the shaft of the turbomachine and from which electrical power is to be drawn; characterized in that it further comprises a regulating system comprising: a fuel pump (4) that is intended to convey fuel into the combustion chamber at a fuel flow rate, said fuel flow rate being a direct function of an ambient pressure and the turbomachine rotational speed; and power electronics (6) coupled to the electrical generator (2), the power electronics (6) being configured to regulate the electrical power that is drawn so as to achieve a target turbomachine rotational speed.
A constant-volume combustion system (1) for a turbomachine comprises a plurality of combustion chambers (100) distributed annularly about an axis (XX'), each combustion chamber comprising an inlet orifice (102) and an outlet orifice (103), a selective shutoff element (200) that is rotatable with respect to the combustion chambers (100), the selective shutoff element comprising a shroud (210) facing the inlet and outlet orifices (102, 103) of the combustion chambers (100). The shroud (210) has, in a first annular portion (211), at least one intake hole intended to cooperate with the inlet orifice (102) of each combustion chamber during the rotation of the selective shutoff element (200) and, in a second annular portion (212), at least one exhaust hole intended to cooperate with the outlet orifice (103) of each combustion chamber during the rotation of the selective shutoff element. Each combustion chamber (100) comprises a fuel injection device (140), the opening and closing of which are synchronized by the shutoff element (200).
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
89.
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).
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]
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
91.
METHOD FOR MANUFACTURING A FLAME TUBE FOR A TURBOMACHINE
The invention relates to a method for manufacturing a flame tube (1) for a turbomachine, the flame tube (1) extending around an axis (X) and comprising a radially inner annular wall (2) and a radially outer annular wall (3), connected to one another by an end wall (4) or head wall, the inner wall (2), the outer wall (3) and the end wall (4) defining an internal space (5), at least one portion of said end wall (4) forming a double wall (6) comprising a first portion (7) and a second portion (8) connected to one another and spaced apart from one another so as to define a flow channel (9) for a cooling air flow opening into said internal space (5), the flow channel (9) comprising at least one air inlet opening (13), the first and second portions (7, 8) of the double wall (6) being connected by connecting areas or bridges (10) extending into the flow channel of the cooling air stream (9), the flame tube (1) being produced 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
93.
COMPONENT FOR A TURBOMACHINE CENTRIFUGAL DEGASSER WITH ADAPTED LONGITUDINAL WALLS
The invention relates to a component for a centrifugal degasser for an air/oil mixture of a turbomachine, the degasser being intended to rotate about an axis of symmetry, forming an annular chamber for centrifugal separation of the mixture, the chamber forming a fluid passage duct, one inlet (5) 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 comprising at least one second oil outlet (8) oriented radially outwards and intended to discharge the oil separated from the mixture to the outside of the degasser, the invention being characterised in that the chamber comprises longitudinal walls (7) passing radially therethrough, at least one of the surfaces of the longitudinal walls (7) 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
94.
TURBINE VANE PROVIDED WITH A RECESS FOR EMBRITTLEMENT OF A FRANGIBLE SECTION
The invention relates to a turbine vane of a turbine engine which comprises a blade (11) and a root (12), the root comprising a stilt (13) having lateral flanks with a curvilinear profile, said stilt comprising 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 comprises at least one oblong frangibility recess (17) formed on at least one of the lateral flanks of the stilt, said oblong recess extending in an axial direction of the stilt along a longitudinal axis (X-X') parallel to or included in a minimum cross-sectional plane (P) which contains a minimum cross-section of the stilt.
F01D 21/02 - Shutting-down responsive to overspeed
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
95.
SEALING RING FOR A WHEEL OF A TURBOMACHINE TURBINE
Sealing ring (36, 38) for a wheel (26) of an aircraft turbomachine turbine, said ring comprising an annular body (51) extending around an axis of revolution (A) and comprising an outer surface (51a) and an inner surface (51b) which is coated with an annular layer (53) of an abradable material, the ring further comprising an annular wall (52) extending around the annular body and at a radial distance from said body, said annular wall comprising openings (54) through which cooling air flows by impact on the outer surface, characterised in that 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 11/14 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
The invention relates to a module (50) of an aircraft turbine engine, said module comprising: - at least one annular housing (24, 32) of an annular combustion chamber (16), - at least one sealing ring (36, 38) for a turbine wheel (26), and - at least one annular bearing support (42), characterized in that this module is made in one piece.
Turbine module (50) for an aircraft turbomachine (10), said module comprising: - arms (52) for connecting annular walls (50a, 50b), said arms comprising upstream leading edges (52a) which are located in a plane P1 perpendicular to an axis, - a nozzle (54) located downstream of the arms and comprising stator vanes (54a) which comprise upstream leading edges (54b) located in a plane P2 perpendicular to the axis and located downstream of the plane P1, and downstream trailing edges (54c) located in a plane P3 perpendicular to the axis and located downstream of the plane P2, the module being characterised in that it is integrally formed, the vanes extending substantially radially between the walls, and the arms extending downstream and each being closely bonded with one of the vanes.
The invention relates to a device (110) for additive manufacturing of at least one part by laser powder bed fusion, comprising: a support plate (134) that can move in vertical translation and is configured to be covered by powder layers (132) and for supporting the part during its manufacture, a closed chamber (140) extending above the support plate, and - a laser beam generator (120) configured to generate a laser beam (122) for scanning the powder layers for localised fusion of the powder and layer-by-layer manufacture of the part, characterised in that it comprises at least one heating system (150) located above the powder layers, said heating system being configured to simultaneously heat more than 10% of the upper surface S.
B22F 3/105 - Sintering only by using electric current, laser radiation or plasma
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
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
The invention relates to a lattice structure (1) for a mechanical component (100) for force transmission which is suitable for production by means of additive manufacture. The structure comprises a lattice having recesses (11), and comprising a plurality of basic structures (12), each basic structure (12) comprising a plurality of beam elements (14) connected to form a polyhedron, wherein the beam elements (14) of each basic structure (12) are at an incline with respect to a reference plane, referred to as the manufacturing plane (P).
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.
patents
