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
Propulsion system for a helicopter, comprising a main engine, a main rotor, a main transmission gearbox comprising an outlet mechanically connected to the main rotor, a reducer mechanically coupled between the main engine and a first input in the main transmission gearbox, and an assistance device. The assistance device comprises a first electric machine mechanically coupled to the reducer and configured to function as an electric generator to extract energy produced by the main engine, and a second electric machine mechanically coupled to a second input in the main transmission gearbox, the second electric machine being supplied with electricity from the first electric machine and configured to function as an electric motor for delivering additional mechanical power to the main transmission gearbox.
This method for stopping at least one aircraft turbogenerator (1) comprises: - controlling the stopping (E1) of the turbogenerator (1); - passing from the nominal operating speed (Nref) of the power shaft (3, 12) to a first operating speed (N1) lower than the nominal speed (Nref), for a first predetermined duration (t2); - controlling the extinction of the combustion chamber (6) of the gas turbine (2); - maintaining the rotation of the gas turbine at a second speed (N2) for a second predetermined duration (t3), the power shaft (3, 12) being at a second speed (N2) lower than the first operating speed (N1) and, - controlling the stopping of the reversible electric machine (7) in order to no longer drive the power shaft (3, 12), in order to cause a progressive stopping (E9, E10) of the rotation of the gas turbine (2).
The invention relates to a 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 27/02 - Aircraft characterised by the type or position of power plant
5.
TURBOMACHINE HAVING A FREE TURBINE COMPRISING ELECTRIC MACHINES ASSISTING A GAS GENERATOR AND A FREE TURBINE
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
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).
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
B29C 64/379 - Handling of additively manufactured objects, e.g. using robots
B29C 64/40 - Structures for supporting 3D objects during manufacture and intended to be sacrificed after completion thereof
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
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
8.
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).
B64D 41/00 - Power installations for auxiliary purposes
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
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
H02J 4/00 - Circuit arrangements for mains or distribution networks not specified as ac or dc
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
9.
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.
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
B64C 11/00 - Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
B64D 31/12 - Initiating means actuated automatically for equalising or synchronising power plants
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
11.
METHOD FOR QUICKLY STOPPING THE ROTOR OF A HELICOPTER AFTER LANDING
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
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 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).
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 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.
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
17.
PRE-VAPORIZING PIPE, COMBUSTION ASSEMBLY PROVIDED THEREWITH AND TURBOMACHINE PROVIDED THEREWITH
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
18.
HYBRID PROPULSION SYSTEM FOR VERTICAL TAKE-OFF AND LANDING AIRCRAFT
The invention relates to a hybrid propulsion system (10) for a vertical take-off and landing (VTOL) aircraft comprising at least one combustion engine (120, 140) driving an electricity generator (122, 142), at least one electrical energy storage assembly (16; 18A, 18N) associated with each electricity generator and defining with each electricity generator an energy branch, a plurality of electric motors (200, 220, 240, 260) actuating a same plurality of rotors (202, 222, 242, 262) providing together the propulsion and/or the lift of the aircraft, and an electrical power and distribution unit (28) supplying power to the plurality of electric motors from the electricity generator and/or the electrical energy storage assembly according to a pre-established flight phase, the system comprising at least two energy branches having an asymmetrical configuration and each selectively supplying, via the power and electrical distribution unit (28), all or part of the plurality of electric motors and in which the electricity generators of the internal combustion engines have between them a power ratio of between 1.2 and 1.4 so as to allow the most powerful to alone supply the electrical power required by the aircraft in all flight conditions.
The invention relates to a system for monitoring the health of a helicopter, comprising a helicopter and a device for determining a change in state of the engine that is configured to collect data measured by engine sensors and external conditions during a stable flight phase and to process said measured data in the following way: comparing said measured data with a reference model of the engine, determining, at each time interval, an instantaneous discrepancy between each item of data measured and each item of data estimated by the reference model of the engine, determining, throughout the stable flight phase, an overall discrepancy between the measured data and the data estimated by the reference model of the engine, determining, at each time interval of the stable flight phase, an intrinsic residual corresponding to a difference between the instantaneous discrepancy and the overall discrepancy, determining a deviating portion corresponding to the part of the intrinsic residual that does not satisfy a predetermined criterion, each deviating portion containing an item of information relating to a pilot action that is not recorded, and determining a corrected residual corresponding to the instantaneous discrepancy from which the deviating portion has been subtracted, the corrected residual being analysed in order to determine whether the state of the engine has changed.
The invention relates to a hybrid propulsion system (100) for a multi-rotor rotary-wing aircraft, comprising: an internal combustion engine (12), an electric machine (14) coupled to the internal combustion engine, a rectifier (16) connected to the electric machine, conversion means (18a, 18b, 18c, 18d), an electric network (20) connecting the rectifier to the conversion means, electric motorss (22a, 22b, 22c, 22d) connected to the conversion means, rotary-blade assemblies (24a, 24b, 24c, 24d) coupled to the electric motors, the system being characterized in that it comprises: detection means configured to detect a reduction in a demand for electrical power within the system to below a predetermined value, bypass means configured to bypass the electric machine when the detection means detects a reduction in the demand for electrical power.
The invention concerns a hybrid propulsion installation (100) for an aircraft, comprising: an internal combustion engine comprising at least one combustion chamber (114); at least one electricity generator (134a, 134n); an electrical propulsion system (138); electromechanical protection means (136a, 136n) arranged between the or each electricity generator and the electrical propulsion system; characterised in that the installation also comprises: an electronic computer (132) configured to: receive at least one piece of information on a position of the electromechanical protection means; receive at least one piece of information on a speed of rotation of a shaft (124) of the internal combustion engine; and control the flow of fuel supplying the combustion chamber depending on the information received.
B64D 27/02 - Aircraft characterised by the type or position of power plant
B60L 50/10 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by engine-driven generators, e.g. generators driven by combustion engines
B60L 3/04 - Cutting-off the power supply under fault conditions
B64D 27/24 - Aircraft characterised by the type or position of power plant using steam, electricity, or spring force
B64D 41/00 - Power installations for auxiliary purposes
A connection (100) comprising a widened reinforcement harness (30) which comprises a reinforced harness (10), a non-deformable ferrule (46) and a reinforcement portion (40) which has a greater diameter than the reinforcement (20) of the harness (10), the non-deformable ferrule (46) being arranged around the harness (10) with the reinforcement (20) of the harness (10) folded on the outer side of the non-deformable ferrule (46), a first end (41) of the reinforcement portion (40) being attached around the folded portion (22) of the reinforcement (20) of the harness (10), the connection further comprising an integrated coupling connector (50), a second end (42) of the reinforcement portion (40) of the widened reinforcement harness (30) being attached to a section (55) of the integrated coupling connector (50). This connection (100) is particularly advantageous for the small harness (10), wherein the reinforcement of the harness (20) has a diameter which is too small to surround the section (55) of an integrated coupling connector (50).
A method for controlling a non-propulsive power generation turbine engine (T) configured to supply power to a plurality of propulsion rotors (R1-R4) of an aircraft, each propulsion rotor (R1-R4) being connected to a power distribution module (30) through at least one power supply bus (B1-B4), the turbine engine (T) supplying each power supply bus (B1-B4) via the power distribution module (30) at a supply rate, the control method comprising a step of determining the power requirement of each power supply bus (B1-B4) depending on the power requirement (P1-P4) of each propulsion rotor (R1-R4), a step of determining the basic power requirement of each power supply bus (B1-B4), a step of determining the overall power requirement (Pglob) based on all the basic power requirements of the power supply buses (B1-B4) and a step of determining an anticipation parameter (N1ap) based on the overall power requirement (Pglob).
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
B64D 27/02 - Aircraft characterised by the type or position of power plant
The invention relates to a hybrid propulsion unit (100) for an aircraft with multi-rotor rotary wings: an electrical generator (114) driven by an internal combustion engine (112), a rectifier (116) configured to convert an AC current sent by the electrical generator into DC current, means for converting (118a, 118b, 118c, 118d) DC current into AC current, an electrical network (120) connecting the rectifier to the conversion means and including a high-voltage DC current bus, electric motors (122a, 122b, 122c, 122d) powered by the propeller conversion means (124a, 124b, 124c, 124d) coupled to the electric motors, the unit being characterized in that it comprises means for storing electrical energy (126) connected to the electrical network, the storage means comprising at least one primary storage element (138) and at least one secondary storage element (140).
The invention relates to a tubular ventilation sleeve (1) for a turbomachine distributor, in particular for an aircraft, the sleeve having a generally elongate shape along an axis (A-A) and comprising a perforated tubular wall (1a) around said axis, one of the axial ends of the sleeve being open and the other being closed by a bottom wall (1b), characterized in that it further comprises support beams (7) when the sleeve (1) is made by additive manufacturing, said beams (7) extending inside the sleeve (1) between the tubular wall (1a) and the bottom wall (1b) and having a longitudinal cross-section with a generally triangular shape, two sides of which are respectively connected to the tubular wall (1a) and the bottom wall (1b) and the last side of which is free and extends inside the sleeve, perforations (2) in the tubular wall (1a) being provided between the support beams (7).
In order to master the consequences of a decoupling event on an axial toothed coupling mechanism (47) of an assembly of rotary elements (33, 35) of a gas turbine, and in a situation of initiation of axial decoupling, an excrescence (51) positioned on one of the fluted teeth of the assembly, away from the coupling axis (X), will cause the two coupling halves to pivot relative to one another about said excrescence.
F01D 5/06 - Blade-carrying members, e.g. rotors - Details thereof, e.g. shafts, shaft connections
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
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
27.
HYBRID-ELECTRIC PROPULSION ARCHITECTURE AND METHOD FOR DISSIPATING ELECTRICAL ENERGY IN SUCH AN ARCHITECTURE
Disclosed is a hybrid/electric propulsion architecture (100) for a multi-rotor rotary wing aircraft, comprising: an electricity generator (114) driven by an internal combustion engine (112), and configured to operate in motor mode, a rectifier (116) configured to convert an alternating current delivered by the electricity generator into direct current, an electrical network (120) comprising a high voltage direct current (HVDC) bus, electrical energy storage means (126) connected to the electrical network, during electrical energy regeneration on the HVDC bus, depending on the state of charge of the storage means (126): the storage means are configured to recover electrical energy, the storage means and the rectifier are configured to recover electrical energy, and the electricity generator operating in motor mode is configured to recover electrical energy.
B60L 15/20 - Methods, circuits or devices for controlling the propulsion of electrically-propelled vehicles, e.g. their traction-motor speed, to achieve a desired performance; Adaptation of control equipment on electrically-propelled vehicles for remote actuation from a stationary place, from alternative parts of the vehicle or from alternative vehicles of the same vehicle train for control of the vehicle or its driving motor to achieve a desired performance, e.g. speed, torque, programmed variation of speed
B60L 50/61 - Electric propulsion with power supplied within the vehicle using propulsion power supplied by batteries or fuel cells using power supplied by batteries by batteries charged by engine-driven generators, e.g. series hybrid electric vehicles
B60L 58/12 - Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling batteries responding to state of charge [SoC]
The invention concerns a method for monitoring the operating state of a system for positioning variable-geometry members (18) of a turbomachine (10), the members (18) being configured to travel over an operating area comprising a first position P1 and a second position P2, the method comprising the steps of: -(E23) determining a first pivoting speed V1 from the first position P1 to the second position P2; -(E25) determining a second pivoting speed from the second position P2 to the first position P1; and -(E26) determining an anomaly in the operation of the system for positioning the members (18) if the first speed V1 is lower than a first determined speed threshold and/or if the second speed V2 is lower than a second determined speed threshold.
F04D 29/56 - Fluid-guiding means, e.g. diffusers adjustable
F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
F01D 17/26 - Devices dealing with sensing elements or final actuators or transmitting means between them, e.g. power-assisted the operation or power assistance being predominantly non-mechanical fluid, e.g. hydraulic
F02C 9/54 - Control of fuel supply conjointly with another control of the plant with control of working fluid flow by throttling the working fluid, by adjusting vanes
29.
ELECTRIC MACHINE WITH DEVICE FOR FORCIBLY DEMAGNETISING PERMANENT MAGNETS
The present invention relates to an electric machine for an aircraft, comprising a stator and a rotor that is rotationally mobile with respect to the stator, the rotor or the stator comprising a plurality of permanent magnets, the machine comprising a device for demagnetising a permanent magnet, suitable for achieving a temporary increase in the temperature of the permanent magnet, in order to limit, during the temporary increase in temperature, an exciting magnetic flux generated by the permanent magnet. The present invention furthermore relates to an assembly comprising an assembly comprising such an electric machine and a hot-fluid source suitable for delivering hot fluid to the demagnetising device of the electric machine. The hot-fluid source may be a gas stream of a turbine engine.
The invention relates to an assembly for a multistage turbine (10) of a turbomachine, the assembly comprising a static sealing device, a turbine stator (22) comprising a radially outer end and an outer casing (16) surrounding the stator (22), the static sealing device being arranged radially between a radially outer end of the stator (22) and the outer casing (16), and comprising an annular seal (44) borne by the stator (22) and an annular structure (60) defining a plurality of radial annular walls (68) axially spaced apart from one another, at least one first wall of said radial annular walls (68) being in annular contact radially inwardly with the annular seal (44) and its longitudinal dimension being less than the longitudinal dimension of the seal (44).
The invention concerns a centrifugal degasser of an air-oil mixture of a turbine engine comprising a hollow shaft (12) extending along an X axis and a pinion (28) for rotating the hollow shaft (12), and an annular chamber (10) for centrifugal separation of said mixture. The chamber (10) is arranged around the hollow shaft (12) and forms a fluid passage area, an inlet (18) of which is oriented axially for feeding the chamber (10) with said mixture, and a first outlet (20) of which is oriented radially inwards for the outlet of the de-oiled air separated from said mixture. The chamber (10) also includes at least one second oil outlet (22) oriented radially outwards and intended for discharging the oil separated from said mixture to the outside of the degasser. The degasser is characterised in that the chamber (10) and the pinion (28) form a single piece.
The invention relates to an aircraft propulsion system, comprising a main transmission unit (12) and at least two turbojet engines connected to the main transmission unit (12), respectively a first turbojet engine (14a) and a second turbojet engine (14b), each turbojet engine comprising a free turbine (24a, 24b), characterized in that the first turbojet engine (14a) comprises a heat exchanger (30) configured to recover some of the thermal energy from the exhaust gas at the outlet of the free turbine, and in that the propulsion system comprises at least one computer (28a, 28b) configured to control the two turbojet engines and to limit the acceleration and the deceleration of the first turbojet engine (14a) when neither of the turbojet engines is broken down, in order to limit the reactor power transients at the heat exchanger (30).
The invention relates to an aircraft power architecture comprising a power transmission gearbox (12), located in a first compartment (30), a gas turbine (14), located in a second compartment (32) comprising a gas generator (18) and a free turbine (22) connected to the power transmission gearbox (12) by a power shaft (26) of the gas turbine (14), and an accessory gearbox (16), the gas turbine (14) being set into the main transmission gearbox (12), characterized in that the gas turbine (14) comprises a first electric machine (38), and in that the accessory gearbox (16) is placed in the first compartment (30) and comprises a second electric machine (42) configured to supply energy to the accessory equipment and to receive electrical energy transmitted via the first electric machine (38).
B64D 27/02 - Aircraft characterised by the type or position of power plant
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
Assembly for a turbomachine (1), comprising an annular channel (6) designed to form a flow duct for a flow of gas between two turbine stages (2, 3) of the turbomachine (1), said channel (6) being bounded by a radially inner annular wall (7) and a radially outer annular wall (8), said walls (7, 8) being connected by hollow arms (16) that extend radially, a support (17) comprising a radially outer annular part (9) that is located radially outside the outer annular wall (8) of the annular channel (6), and a radially inner annular part (12) that is located radially inside the inner annular wall (7) of the annular channel (6), the outer and inner parts (9, 12) of the support (17) being connected by connecting parts (18) that extend radially, each connecting part (18) passing through one of the hollow arms (16) of the annular channel (6), in which at least one of the connecting parts (18) of the support (17) and the corresponding hollow arm (16) are connected to one another by at least one connecting partition (22), said connecting partition (22) comprising a frangible part (23) that is able to rupture when the mechanical stresses in said connecting partition (22) are greater than a certain value.
An electrical generator (9) is housed in an annular cavity (14) between the casing (15) and the propeller shaft (3) of a turboprop, while imposing little or no additional space requirement and with lightweight ancillary equipment (16) and (25). The rotor (10) of the generator is mounted on an autonomous shaft end (16). A flange (20) of the outer casing (5) is removable in order to access the generator and to enable its easy removal and remounting.
The invention relates to a component for a centrifugal breather for removing gas from a turbomachine air/oil mixture, which component is intended to rotate about an axis of symmetry (X), comprising a structural part (2, 3) configured to delimit a flow path (4) for the flow of said mixture comprising: a circumferential chamber (10) for the centrifugal separation of said mixture; an axial inlet (6) for said air/oil mixture into said chamber (10); radial oil outlets (9), an outlet (16) for oil-free air, further comprising at least one cellular structure (5) made for example from a material of the metallic foam type, configured to filter the oil while allowing the air to pass and occupying at least a space (17) in said flow path (4) which closes the communication between the axial inlet (6) and the internal radial outlet (12), characterized in that the structural part (2, 3) and the cellular structure (5) are formed as a single component. The invention also relates to a method for the additive manufacture of said component.
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
B01D 46/26 - Particle separators, e.g. dust precipitators, using rigid hollow filter bodies rotatable
F01M 13/04 - Crankcase ventilating or breathing having means for purifying air before leaving crankcase, e.g. removing oil
37.
INJECTEUR DE CARBURANT A JET PLAT POUR UNE TURBOMACHINE D'AERONEF
A flat-jet fuel injector for an aircraft turbine engine, comprising a body (112) having a generally elongate shape having a longitudinal axis A, said body comprising a main pipe (118) having a generally elongate shape having a longitudinal axis B substantially perpendicular to said longitudinal axis A, the two longitudinal ends (120) of said main pipe being connected directly and respectively to longitudinal ends of two secondary pipes (119) having a generally elongate shape having a longitudinal axis C at least substantially parallel to said longitudinal axis A, and being configured to form, respectively, two separate fuel flow inlets intended to meet substantially at the middle of said main pipe which comprises at least one ejection slot (124) for ejecting said fuel jet, characterised in that at least one of said main and secondary pipes defines a flow area, at least one geometric parameter of which, such as the shape or a dimension, varies along said pipe and/or is different from the same geometric parameter defined by a flow area of another of said pipes.
Electric machine comprising a stator (1) and a rotor (2) configured to be driven in rotation in relation to one another, the rotor (2) comprising a plurality of permanent magnets (5), the stator further comprising a magnetic circuit (3) comprising poles (7) extending toward the rotor (2), the machine comprising windings of conducting elements around each pole (7) and at least one heat sink (8) arranged inside a conducting element and/or between the conducting elements, the heat sink comprising a phase change material.
H02K 3/22 - Windings characterised by the conductor shape, form or construction, e.g. with bar conductors consisting of hollow conductors
H02K 3/24 - Windings characterised by the conductor shape, form or construction, e.g. with bar conductors with channels or ducts for cooling medium between the conductors
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
39.
METHOD FOR CHECKING THE MAXIMUM AVAILABLE POWER OF A TURBINE ENGINE OF AN AIRCRAFT EQUIPPED WITH TWO TURBINE ENGINES
The invention concerns a method (100) for checking the maximum available power of a turbine engine (11, 12) of an aircraft (10) equipped with two turbine engines (11, 12) configured to operate in parallel and together to supply a necessary power (P1 +2) to the aircraft during a flight phase, said method comprising the following steps:- placing one of the turbine engines (11, 12) in a maximum take-off power regime (PMD), and - adjusting a power (P2 , P1) supplied by the other turbine engine (12, 11), such that the turbine engines (11, 12) continue to supply the necessary power (P1 +2) to the aircraft during the flight phase; - determining a power (P1, P2) supplied by the turbine engine (11, 12) placed in the maximum take-off power regime (PMD), and - processing the supplied power (P1, P2) determined in this way, in order to deduce a piece of information relating to the maximum available power.
A device for temporarily increasing power (13) for increasing the power of at least one first turbomachine (5A) and at least one second turbomachine (5B), said device comprising a coolant reservoir (14), a first injection circuit (16A), connected to said reservoir (14) and opening onto at least one injection nozzle (22) configured to be installed upstream from the first turbomachine (5A), a second injection circuit (16B), connected to said reservoir (14) and opening onto at least one injection nozzle (22) configured to be installed upstream from the second turbomachine (5B), the first injection circuit (16A) and the second injection circuit (16B) each comprising at least one first valve (23) and at least on second valve (29) arranged upstream from said at least one first valve (23), and a bypass pipe (40) connecting the first injection circuit (16A) and the second injection circuit (16B), upstream from the first valve (23) of same and downstream from the second valve (29) of same.
F02C 6/02 - Plural gas-turbine plants having a common power output
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 3/30 - Adding water, steam or other fluids to the combustible ingredients or to the working fluid before discharge from the turbine
F02C 7/143 - Cooling of plants of fluids in the plant of working fluid before or between the compressor stages
41.
TEST SPECIMEN FOR VALIDATING OPERATING PARAMETERS OF A METHOD FOR THE ADDITIVE MANUFACTURING OF A PART BY LASER MELTING ON POWDER BEDS
The invention relates to a test specimen (200) for validating operating parameters of a method for the additive manufacturing of a part by laser melting on powder beds, said test specimen having at least one upper face (210), at least one lower face (212) and side faces (214, 216, 218, 220), including a front face (214) and a rear face (216) that are substantially on opposite sides from one another, said test specimen also having at least one recess (217) opening onto said front and rear side faces, said recess having a substantially triangular shape in cross section and being delimited by three internal faces, including a first, lower internal face (222) oriented upwards, a second, upper internal face (224) oriented downwards, and a third, upper internal face (226), which defines a narrow wall (228) with one other (220, 220b) of said side faces, which is inclined.
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
B28B 1/00 - Producing shaped articles from the material
The invention relates to a centrifugal deaerator for an air/oil mixture of a turbomachine, comprising: an annular housing (10) for the centrifugal separation of said mixture, arranged around a hollow shaft (11) and defined by an outer annular wall (12) and an inner annular wall (11); axial inlets (14) for the inflow of said air/oil mixture into said housing (10); a pinion (20) for rotating said housing (10), comprising a web (21) that is securely connected to said hollow shaft (11) and to said inner (13) and outer (12) annular walls; radial oil outlets (15) in said outer wall (12); and oil-free air outlets (16) in said inner wall (13), characterised in that said axial mixture inlets (14) and said radial oil outlets (15) are axially arranged either said of said web (21) of said pinion (20) so as to prevent said axial mixture inlets (14) reintroducing oil evacuated by said radial oil outlets (15) into the enclosure (10). In one of the modes of embodiment, the cells of the deaerator can contain a metal filter.
The invention relates to a centrifugal deaerator for an air/oil mixture of a turbomachine, comprising: an annular housing (10) for the centrifugal separation of said mixture, arranged around a hollow shaft (11); axial inlets (14) for the flow of said air/oil mixture into said housing (10); a pinion (20) for rotating said housing (10); and radial oil outlets (15) and oil-free air outlets (16), characterised in that it comprises a metal foam (50) housed in said housing (10) by partially extending along the axial direction (XX') so as to define two successive spaces, a foam-free space (51) and a foam-lined space (52), said foam-free space opening up towards said axial inlets. A device can be advantageously added to the hollow shaft in order to reduce the free-vortex phenomenon and to thereby significantly reduce the load losses of the deaerator.
The invention relates to a temperature measurement device comprising at least three probes, a computer, the computer being configured to estimate a temperature on the basis of voltage measurements at the output of the probes, characterised in that the probes are connected together in a plurality of meshes mounted in series, with at least one mesh comprising at least two probes mounted in parallel and at least one other mesh comprising a probe or a plurality of probes mounted in parallel.
G01K 7/02 - Measuring temperature based on the use of electric or magnetic elements directly sensitive to heat using thermoelectric elements, e.g. thermocouples
G01K 13/024 - Thermometers specially adapted for specific purposes for measuring temperature of moving fluids or granular materials capable of flow of moving gases
G01K 7/12 - Arrangements with respect to the cold junction, e.g. preventing influence of temperature of surrounding air
45.
PROCEDE ET SYSTEME DE COMMANDE D'UN DISPOSITIF D'URGENCE
The invention relates to a method for controlling an emergency device of a helicopter, said helicopter comprising a rotor suitable for being rotated, said emergency device being suitable for supplying additional emergency propulsion power to the helicopter, said method comprising a step (10) of measuring the rotation speed of the helicopter rotor, a step (12) of calculating the drift of the measured rotation speed, a step (20) of continuously verifying conditions such that the speed of rotation of the rotor is higher than a predetermined value, referred to as arming speed, and the drift of the rotation speed is lower than a predetermined value, referred to as arming drift, and a step (22) of activating the emergency device if the verified conditions are validated.
Propulsion system (8) for a single-engine helicopter (1), comprising: a main engine (9) connected to a front drive shaft (5) and a rear drive shaft (7), respectively, suitable for driving a main gearbox (4) referred to as MGB (4) and a tail gearbox (6) referred to as TGB (6); an assistance device (10) attached to the main engine (9); characterised in that said propulsion system (8) is designed in order that the assistance device (10) can mechanically drive the TGB and MGB (6, 4) by introducing power to the rear drive shaft (7).
The invention relates to a blade of a high-pressure turbo-engine turbine, comprising: - a blade airfoil (13) extending in a span direction (EV), terminating in an apex (S) and comprising a lower surface (16) and an upper surface (17) joined by a leading edge (18) and trailing edge (19), and - an internal cooling circuit (25) having only an upstream duct (26) and a central chamber (27) for cooling the blade by circulating air; - wherein the upstream duct (26) and the central chamber (27) are separately supplied with air; - the upstream duct (26) cools the leading edge (18) and the upper surface (17); - and the central chamber (27) cools the lower surface (16) and the trailing edge (19), and is provided with bridge elements (51) interconnecting the lower surface (16) and the upper surface (17).
The invention relates to a wheel blade (40) for a turbomachine compressor. The blade (40) comprises a body (41) extending between the root (P) of the blade in the direction of the height (Y-Y) of the blade. The body (41) extends from a leading-edge (BA) to a trailing edge (BF) in a direction of the camber line (S-S) of the blade. The blade (40) comprises a winglet (50) located in the extension of the body (41) in the direction of the height (Y-Y). The winglet (50) is located at the tip (S) of the blade and at the leading edge (BA) of the blade. The body (41) has a curvature oriented in a first direction of rotation with respect to the direction of the camber line (S-S), the winglet (50) having a curvature in a second direction of rotation counter to the first direction of rotation.
The invention relates to a method for producing a shielding cartridge (1) of a turbomachine element (10), said method comprising a step of winding a carbon wire (3) comprising a plurality of fibres about a longitudinal axis and a step of trapping and preserving the outside surroundings of the wound-up wire (3) in a housing (2).
B29C 70/32 - Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core on a rotating mould, former or core
B29C 70/86 - Incorporating in coherent impregnated reinforcing layers
B32B 15/14 - Layered products essentially comprising metal next to a fibrous or filamentary layer
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
F41H 5/04 - Plate construction composed of more than one layer
The invention relates to the field of pyrotechnic devices and more specifically to a pyrotechnic device (1) comprising a main pyrotechnic charge (3), a firing device (2) for firing the main pyrotechnic charge (3), a discharge channel (5) for discharging gases generated by firing the main pyrotechnic charge (3) and an injection device (9) configured to inject a cooling fluid (10) into said gas discharge channel (5) so as to supply gases, particularly for driving turbines, at relatively low temperatures. The invention also relates to a method for cooling the gases generated by the firing of the main pyrotechnic charge (3) by injecting the cooling fluid (10).
Method for collecting operational data of an aircraft (1), this aircraft comprising a screen (2) for displaying operational data (3). The method is characterised in that: - by means of a mobile terminal (4) provided with at least one optical sensor, images (5) of said display screen (2) are acquired so that at least some of the images acquired show operational data displayed on the display screen; by means of an optical character recognition (OCR) module associated with said mobile terminal (4), strings of characters (C1, C2), each representing one of said operational data (3) displayed on the display screen (2) are identified in the images (5) acquired; - by means of a report generating module and said strings of characters (C1, C2) identified, an electronic report (6) containing at least some of said operational data (3) of the aircraft (1) is generated.
The invention relates to a process for manufacturing a turbomachine blade (P) of the type having at least one 3D cavity (7), characterized in that said blade is produced by a succession of depositions and selective consolidations of layers of a metal additive manufacturing powder based on an alloy of copper and nickel, said alloy comprising from 2% to 7% of nickel. It also relates to a turbomachine blade, characterized in that it is manufactured by metal additive manufacturing using said process.
OFFICE NATIONAL D'ETUDES ET DE RECHERCHES AEROSPATIALES (France)
SAFRAN HELICOPTER ENGINES (France)
Inventor
Rame, Jeremy
Belaygue, Philippe
Caron, Pierre
Delautre, Joel
Jaquet, Virginie
Lavigne, Odile
Abstract
The invention relates to a superalloy based on nickel, comprising, in mass percentages, 4.0 to 6.0 % of chromium, 0.4 to 0.8 % of molybdenum, 2.5 to 3.5 % of rhenium, 6.2 to 6.6 % of tungsten, 5.2 to 5.7 % of aluminium, 0.0 to 1.6 % of titanium, 6.0 to 9.9 % of tantalum, 0.0 to 0.7 % of hafnium, and 0.0 to 0.3 % of silicon, the rest consisting of nickel and potential impurities. The invention also relates to a monocrystalline blade (20A, 20B) comprising such an alloy and to a turbomachine (10) comprising such a blade (20A, 20B).
The invention relates to a method for monitoring the soundness of helicopters comprising the determination of the severity of a plurality of flight missions of a plurality of helicopters, comprising a step for acquiring and storing flight data from helicopter flight missions, and a step for acquiring and storing maintenance data from the plurality of helicopters. The method is characterised in that said determination comprises a mission-type construction step, comprising a sub-step for constructing descriptors, a sub-step for partitioning the descriptors and a sub-step for allocating a mission type to each flight by associating the descriptor of said flight and a sub-set, in which this descriptor is found, and a step for interpreting the severity of the mission types, comprising a sub-step for estimating the severity models, and a sub-step for associating a severity model with each mission type determined in the mission type construction step.
The electrical machine according to the invention is a dual machine, and comprises a first machine (32, 33, 34) which can be reversed and a second machine (30, 35) functioning through induction between two windings. The second machine can be used for de-icing a propeller, the induction-receiving winding (30) being mounted on the propeller shaft (2). The first electrical machine can function as an electrical engine to 'taxi' the aircraft. In certain embodiments, the winding mounted on the stator (31) is common to the two machines and different magnetic flows are utilised to control them. A separate functioning of the two machines and a very good integration into the aircraft engine, with space-saving and low mass, are possible.
The invention relates to a system for deoiling an air-oil pressurising mixture (11, 12) for pressurised enclosures (13) of a turbine engine, said system comprising: at least one centrifugal degasser (20); at least one mechanical casing (30) of the turbine engine comprising a plurality of mechanical parts (21, 22), at least one of which is mechanically connected to a drive shaft of said degasser (20) so as to be able to rotate same, said casing (30) being configured to be able to contain an oil mist (39) for lubricating said mechanical parts (21, 22), characterised in that said deoiling system also comprises means (40) for isolating said air-oil mixture (11, 12) from said oil mist (39) of said mechanical casing (30) so that said oil mist (39) of said mechanical casing (30) cannot enrich with oil said mixture to be separated by said degasser.
The invention relates to an electric actuator for air intake flaps of a gas turbine engine comprising a fixed part (104) attached to a stationary part of the engine and a moving part (102) mechanically connected to the air intake flaps, this electric actuator (100) comprising, uniformly distributed at the periphery of the moving part and fixed to the fixed part, fixed electromagnets (108a-108h) and, arranged between these fixed electromagnets, mobile electromagnets (110a-110h) each one secured on each side to first (114a-114h) and second (116a-116h) piezoelectric elements, the mobile electromagnets and the first and second piezoelectric elements having a degree of freedom to move with respect to the moving part.
Adjustable-length connecting rod (40) for an aircraft turbomachine (10), this connecting rod having a generally elongate shape and comprising two longitudinal opposite ends (42, 50) for attachment to elements to be connected, characterised in that: - a first (42) of these longitudinal ends is connected to a first threaded shaft (44) screwed into a first tubular portion (46) of an adjustment sleeve (48), - a second (50) of these longitudinal ends is connected to a second threaded shaft (52) screwed into a second tubular portion (54) of the adjustment sleeve, and further comprising a housing (56) in which at least a portion of the first shaft is configured to slide.
The invention concerns an assembly comprising: a fuel supply circuit (15, 15a, 15b) configured to supply fuel to a turbine heat engine, an electronic module (14, 14a, 14b), a power source (13, 13a, 13b) for supplying power to the electronic module (14, 14a, 14b), and a heat exchanger (16, 16a, 16b) positioned to allow a flow of heat from the electronic module (14, 14a, 14b) to the fuel supply circuit (15, 15a, 15b), the assembly being characterised in that the electronic module (14, 14a, 14b) comprises a phase-change material (PCM), configured to change state when the temperature of same reaches a predetermined phase-change temperature (Tf).
The invention relates to a fuel supply system (20) for a turbomachine. The supply system (20) comprises a starter circuit (120), at least one first hydraulic resistance (124, 136) and a purge circuit (130). The purge circuit (130) comprises a duct (131) comprising an opening (133) at the exterior of the supply system (20). The purge circuit (130) is configured to make purge air flow between a starter injector (126) and the opening (133) through the first hydraulic resistance (124, 136). The supply system (20) comprises means for measuring a value representative of pressure (141), which means are configured to measure a value representative of pressure (Pm) between the starter injector (126) and the first hydraulic resistance (124, 136) when purge air flows between the starter injector (126) and the first hydraulic resistance (124, 136).
Actuating system for an aircraft including: an electro mechanical actuator, (25) which includes a non volatile memory (60) in which are kept stored data (61) including configuration data (62) specific to the electro mechanical actuator; a control unit (22) using configuration data to implement a feedback control loop having as an output signal a digital command signal of the electro mechanical actuator's electric motor; at least one digital transmission channel (50) connecting the control unit and the electro mechanical actuator.
The invention relates to an aircraft engine part (1) comprising at least one metal substrate (2) and an erosion protective coating (3) present on the substrate, the coating comprising at least one phase (4) comprising at least chromium in an atomic content greater than or equal to 45% and carbon in an atomic content between 5% and 20%, said phase comprising chromium carbides Cr7C3 and Cr23C6. The invention also relates to a process for manufacturing such a part in which a coating composition is deposited on the part by electrodeposition and the part is heat treated at a temperature between 250°C and 700°C.
C25D 5/50 - After-treatment of electroplated surfaces by heat-treatment
B32B 15/04 - Layered products essentially comprising metal comprising metal as the main or only constituent of a layer, next to another layer of a specific substance
B32B 15/16 - Layered products essentially comprising metal next to a particulate layer
B32B 33/00 - Layered products characterised by particular properties or particular surface features, e.g. particular surface coatings; Layered products designed for particular purposes not covered by another single class
C25D 3/06 - Electroplating; Baths therefor from solutions of chromium from solutions of trivalent chromium
63.
PARTICLE-TRAPPING DEVICE FOR A TURBOMACHINE AND TURBOMACHINE WITH SUCH A DEVICE
The invention relates to a particle-trapping device (2) for a turbomachine, said particles being contained in an air stream flowing inside a turbomachine, in particular the air stream flowing in the bypass region (17) of the combustion chamber (13) of said turbomachine. The device is characterised in that it comprises: - at least two particle deflectors (3, 3a, 3b, 3c), - a member (5) for collecting and storing the particles deflected by said deflector, - and means (6) for attaching said trapping device (2) to a part of the turbomachine.
F01D 25/32 - Collecting of condensation water; Drainage
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
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
The invention relates to the field of magnetohydrodynamic generators, and more specifically to such a generator (10) including a flow duct (11) for a working fluid, delimited by a first wall (12) and a second wall (13), a working fluid ionization device (14), a pair of arms (15), each connecting the first and second walls (12, 13) downstream of said ionization device (14) so as to define, between said arms (15) and said walls (12, 13), a channel (16) in the flow duct (11), said channel (16) being arranged so as to have some of the working fluid passing therethrough after the ionization thereof, a magnet for generating a magnetic field (B) oriented perpendicularly to the flow of the working fluid in the channel (16) delimited by the pair of arms (15) and said walls (12, 13), and at least one pair of electrodes (17). Each electrode (17) of each pair is arranged on one side of the channel (16) delimited by the pair of arms (15) and said walls (12, 13). Said electrodes (17) of each pair are spaced apart from one another perpendicularly to said magnetic field (B) and the flow of the working fluid in the channel (16) delimited by the pair of arms (15) and said walls (12, 13).
The invention relates to a device for controlling a propeller, having variable-pitch blades, of a turboprop engine, including a first hydromechanical device (10) for controlling the pitch of the blades of the propeller and a second hydromechanical device (20) for controlling the speed of rotation of the propeller, characterised in that said device includes a single electromechanical actuator (30) including a movable actuator member (31, 36) mechanically connected both to said first hydromechanical device (10) for controlling the pitch, in order to manage the pitch setpoint, and to said second hydromechanical device (20) for controlling the speed, in order to manage the speed setpoint.
The invention relates to a torque meter comprising a casing (12) in which a translationally movable part (14) may move in a longitudinal direction under the effect of an axial pressure representative of the torque to be measured. The torque meter is characterised in that it comprises: a bearing surface (22) connected to the casing (12) via at least one of its ends and lying at least partially in a plane substantially perpendicular to the longitudinal direction, one longitudinal end of said movable part (14), which end is called the contact end (24), being suitable for being brought into contact with said bearing surface (22) so that a longitudinal movement of the movable part (14) leads to a deformation of the bearing surface (22); and means (26) for measuring the deformation of the bearing surface (22).
G01L 3/14 - Rotary-transmission dynamometers wherein the torque-transmitting element is other than a torsionally-flexible shaft
F02C 7/00 - Features, component parts, details or accessories, not provided for in, or of interest apart from, groups ; Air intakes for jet-propulsion plants
67.
TURBOSHAFT ENGINE MOUNTABLE IN A REDUCTION GEARBOX
The invention relates to a turboshaft engine for a helicopter, comprising a case (5) inside which a gas generator (6) and a turbine (7) are accommodated, the turbine (7) being mounted on a power shaft (8) that extends along a longitudinal direction (X'X), the turboshaft engine further comprising means for removably mounting the power shaft (8) into a reduction gearbox (10) inside which at least one gear (11) of a first reduction stage is accommodated; the means for removably mounting the power shaft include a pinion (12) having a central bore, the shape of which is adapted to that of the power shaft (8) in such a way that the pinion (12) can slide over the power shaft (8); furthermore, the contour of the pinion (12) is adapted to the shape of the gear (11) of the first stage in such a way that the pinion (12) can form a leading input pinion of the gearbox in said gear once the power shaft (8) has been mounted in the reduction gearbox (10).
B64D 35/00 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions
F01D 25/16 - Arrangement of bearings; Supporting or mounting bearings in casings
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 flow limiter which comprises a body (1) comprising a fluid intake (2) and a fluid outlet (3), in which body a pipe (5) for circulating the fluid is provided, comprising a series of chambers having different cross-sections, connected to the fluid intake as well as to the fluid outlet, the pipe being generally tubular, and the pipe having a bottom which has a regular curve so that the pipe does not have one or more areas that retain the fluid circulating in the pipe.
The invention relates to a device for controlling an auxiliary engine (8) including a gas generator and a free turbine suitable for being connected mechanically to the rotor (12) of a helicopter in order to supply same with thrust power, characterised in that said control device includes a proportional-plus-integral controller (30) having proportional gain (Kp) and integral gain (Ki), which are functions of the speed of rotation of said gas generator, said controller (30) being configured to receive an error signal representing a speed error of said free turbine, and to generate a signal (Sc) for correcting the drive speed of said gas generator obtained by adding a signal which is proportional to said error signal according to said proportional gain (Kp), and an integrated signal (Si), resulting from adding a signal which is proportional to said error signal according to said integral gain (Ki) and a memory signal (Sm), provided by a feedback loop (31) of said integrated signal (Si), said memory signal (Sm) being dependent on a measurement that represents the speed of rotation of said free turbine.
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 6/02 - Plural gas-turbine plants having a common power output
70.
TWIST-BREAKABLE MECHANICAL FUSE AND COOLING UNIT OF A TURBINE ENGINE FITTED WITH SUCH A FUSE
The invention relates to a mechanical fuse intended to be rigidly mounted between a drive unit (8, 9) and a receiver unit (10), each rotating about a same axis (7) of rotation, said fuse comprising a body (13) extending in a longitudinal direction parallel to said axis (7) of rotation, once the fuse is mounted between said drive unit (8, 9) and receiver unit (10). The invention is characterized in that said body (13) comprises a plurality of longitudinal bars (14), each bar (14) being deformable by bending, such as to form a twist-breakable mechanical fuse.
F16D 9/06 - Couplings with safety member for disconnecting by breaking due to shear stress
B64D 33/08 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
F01P 5/14 - Safety means against, or active at, failure of coolant-pump drives, e.g. shutting engine down; Means for indicating functioning of coolant pumps
F16D 1/06 - Couplings for rigidly connecting two coaxial shafts or other movable machine elements for attachment of a member on a shaft or on a shaft-end
71.
AIR SUPPLY PLENUM WITH STEPS, LIPS AND AIR INTAKE PROTECTION GRID
The invention relates to an air supply plenum (12) for an engine, the plenum being disposed upstream of an air intake (2) of said engine, the air intake being provided in a casing (3) of said engine. The air supply plenum (12) comprises a first lateral wall (121) and a second lateral wall (122) which together form a conduit in which an air flow (7) flows as the engine functions. The invention is characterized in that each lateral wall (121, 122) comprises a step (8) which forms a transverse recess relative to the direction of the air flow (7) in the air supply plenum (12), such that an aerodynamic separation (6) occurs in the step (8) when the engine is operating.
B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
F02C 7/055 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with intake grids, screens or guards
The invention relates to a propulsion unit (10), notably for an aircraft, comprising - a turbomachine (12); - a propeller (14) that can be selectively coupled to said turbomachine (12); - a rotary electrical machine (16) able to drive at least the turbomachine (12), characterized in that it comprises means (20) of selective coupling of the rotary electric machine (16) to the propeller (14) and/or the turbomachine (12), which means are able to couple only the propeller (14) to the rotary electric machine (16) during a defined stage of operation of the propulsion unit (10) and which are able to couple or not couple the rotary electric machine (16) to a gas generator and/or a free turbine of the turbomachine (12).
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
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
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
73.
HYBRID PROPULSION SYSTEM FOR A MULTI-ENGINE AIRCRAFT
The hybrid propulsion system for a multi-engine aircraft comprises a plurality of free turbine turbomachines each one equipped with a gas generator, these including at least one first turbomachine (1) or hybrid turbomachine able to operate in at least a standby state during a stabilized flight of the aircraft, while other turbomachines of the plurality of turbomachines operate alone during the course of this stabilized flight. The hybrid turbomachine (1) is associated with first and second identical electrotechnical sequences each comprising an electric machine (2, respectively 3) that can operate as a starter and as a generator, and is itself connected to a power electronics module (4, respectively 5) itself selectively connected to a specific electrical energy supply network (8), such as an onboard network, and to at least one electrical energy storage member (6, respectively 7). Each of the electrotechnical sequences is designed to deliver a maximum power at least equal to half the total power (Prr) needed for rapid reactivation of the hybrid turbomachine (1).
B64D 31/06 - Initiating means actuated automatically
B64D 33/00 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for
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 41/00 - Power installations for auxiliary purposes
REMOVABLE PACK FOR REACTIVATING A TURBOSHAFT ENGINE, ARCHITECTURE FOR A MULTI-ENGINE SYSTEM FOR PROPELLING A HELICOPTER, PROVIDED WITH SUCH A PACK, AND CORRESPONDING HELICOPTER
The invention relates to a removable pack for reactivating a helicopter turboshaft engine (6), including a gas generator (7) provided with a drive shaft (12). Said turboshaft engine (6) is capable of operating in at least one standby mode during a stabilized flight of the helicopter. Said removable pack includes: a removable housing (30) including a housing output shaft (31); controlled means (32, 33) for rotating said housing output shaft (31), said means being a so-called means for reactivating said turboshaft engine; and mechanical means (14) for reversibly coupling said housing output shaft (31) with said shaft (12) for driving said gas generator.
The invention relates to an architecture for a propulsion system of a multi-engine helicopter including turboshaft engines connected to a power gearbox. Said architecture is characterized in that it includes: - at least one hybrid turboshaft engine (20) capable of operating in at least one standby mode during a stabilized coasting flight of the helicopter; - at least two chains (30, 40) for controlling each hybrid turboshaft engine (20), each chain (30, 40) including an electrical machine (31, 41) connected to the hybrid turboshaft engine (20) and being suitable for rotating a gas producer thereof; and - at least one source (33, 43) for supplying electrical power to said electrical machine (31, 41). Each reactivation chain (30, 40) is configured so as to be able to drive said turboshaft engine (20) in at least one operation mode from among a plurality of predetermined modes.
B64C 27/24 - Compound rotorcraft, i.e. aircraft using in flight the features of both aeroplane and rotorcraft with rotor blades fixed in flight to act as lifting surfaces
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
F01D 13/00 - Combinations of two or more machines or engines
F01D 19/00 - Starting of machines or engines; Regulating, controlling, or safety means in connection therewith
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
Method for notifying an authorization to completely shut down an aircraft gas turbine engine and aircraft gas turbine engine comprising a device for such notification, the method comprising, following the detection (E10) that the engine has switched to low idle: (a) a step (E20) of evaluating, from a value of a first engine operating parameter, a value of a second parameter T45MG characterizing a thermal behaviour of an engine component subject to coking; (b) a step (E30) of comparing the value of the second parameter T45MG against a predefined threshold value T45threshold corresponding to a value of the second parameter that does not cause coking of said component; and (c) a step (E50) of notifying authorization to completely shut down the engine if the value of the second parameter T45MG is below the predefined threshold value T45threshold, steps (a) to (c) being reiterated if not.
The invention relates to a device for testing the integrity of a rapid restart system for a helicopter turbine engine (5) comprising a pneumatic turbine connected mechanically to said turbine engine (5) and supplied with gas, on demand, by a pneumatic supply circuit (8) so as to be able to spin said turbine engine (5) and reactivate it, said test device being characterized in that it comprises means (21, 22) of bleeding pressurized air from the turbine engine (5); a pipe (23) for conveying this bled air to said pneumatic circuit (8) for supplying gas to said pneumatic turbine; and means for determining the rotational speed of said pneumatic turbine.
The present invention concerns a flight control system of an aircraft comprising: -a first processing unit (1), -a second processing unit (2), -communication means configured to establish a first two-way digital link (3) and a second two-way digital link (4) between the first processing unit (1) and the second processing unit (2), said second link (4) being redundant with the first link (3), and said first link (3) and second link (4) being likely to be active concomitantly, said system further comprising backup communication means enabling data exchanges between the first processing unit (1) and the second processing unit (2) in the case of a failure in the first link (3) and second link (4), said backup communication means comprising an array of sensors or actuators (13) and/or a secure onboard network for the avionics (14).
The present invention concerns a method for switching, by a local processing unit (1,2) of a flight control system of an aircraft, configured to control at least one local actuator, connected to at least one local sensor and connected via at least one link (3,4) to an opposite processing unit (2,1) configured to control at least one opposite actuator and be connected to at least one opposite sensor, said local processing unit (1,2) being further configured to be connected to backup communication means (13,14) enabling data exchanges between the local processing unit (1,2) and the opposite processing unit (2,1) in the case of failures of the links connecting same (3,4), said backup communication means comprising an array of sensors or actuators (13) and/or a secure onboard network for the avionics (14), comprising steps of: - sending, to the opposite processing unit (2,1), acquisition data relative to the at least one local sensor and actuator data relative to the at least one local actuator, - receiving, from the opposite processing unit (2,1), acquisition data relative to the at least one opposite sensor and actuator data relative to the at least one opposite actuator, - receiving an item of opposite health data and determining an item of local health data, - switching said local processing unit (1,2) from a first state to a second state chosen from an active state (15), a passive state (16) and a slave state (18), depending on the opposite health data received and the local health data determined.
The quick-assist device applies to a free-turbine turbomachine (11) of an aircraft comprising at least a first free-turbine turbomachine (11) provided with a gas generator, combined with an electric machine (12) that can operate as a starter and as a generator, the first turbomachine (11) being capable of being brought into idle mode or into unwanted switch-off mode, the electric machine being powered from an onboard power supply network (17). The device further comprises at least one electric power storage unit (14) suitable for being electrically connected to the electric machine (12) combined with the first turbomachine (11) in order to provide punctual assistance to the gas generator of said turbomachine (11). The electric power storage unit (14) constitutes a so-called "primary" power storage unit that is non-rechargeable and single-use. The device comprises a system for activating the electric power storage unit (14) and a device (15) for coupling the electric power storage unit with a power supply system (13, 16) of the electric machine (12).
B64D 31/00 - Power plant control; Arrangement thereof
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 41/00 - Power installations for auxiliary purposes
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 7/32 - Arrangement, mounting, or driving, of auxiliaries
H01M 10/46 - Accumulators structurally combined with charging apparatus
81.
PNEUMATIC DEVICE FOR RAPIDLY REACTIVATING A TURBINE ENGINE, ARCHITECTURE FOR A PROPULSION SYSTEM OF A MULTI-ENGINE HELICOPTER PROVIDED WITH SUCH A DEVICE, AND CORRESPONDING HELICOPTER
The invention relates to a device for rapidly reactivating a helicopter turbine engine (6). Said device is characterized in that it includes an air turbine (7) mechanically connected to said turbine engine (6) so as to be able to rotate it and ensure the reactivation thereof; - a pneumatic storage device (9) connected to said air turbine (7) by means of a pneumatic circuit (10) for supplying pressurized gas to said air turbine (7); and - a controlled, rapid-opening air valve (11) arranged on the pneumatic circuit (10) between said storage device (9) and said air turbine (7) and controllable at least in an open position, wherein the gas can be supplied to said air turbine (7), or in a closed position, wherein said air turbine (7) is no longer supplied with pressurized gas.
The invention relates to an overspeed protection device of an aircraft engine comprising a fuel supply system of said aircraft engine, the protection device comprising a voltage source configured to deliver voltage according to negative or positive polarization via a polarizer mounted in parallel with the voltage source, said polarizer being configured to polarize voltage coming from the voltage source; a logic control device connected in series with the voltage source, said logic control device being configured to open or close the supply system as a function of the polarization of the delivered voltage.
The present invention relates to an assembly for a turbomachine combustion chamber (20), comprising: -a boss (32) for turbomachine combustion chamber (20), comprising a tubular body (33) configured to be inserted in a receiving orifice (29) formed in an annular wall (25) of a flame tube (24) and to accommodate a shank (27) of a connecting pin (26) for connecting the flame tube to an external casing (21), a first end (34) of the boss comprising an annular flange (35) configured to prevent radial translational movement of the boss in a first direction (A), and a second end (37) of the boss being configured to be crimped to the wall (25) of the flame tube (25) so as to prevent radial translational movement of the boss in a second direction (B), and -an annular element (39) configured to fit coaxially with the receiving orifice (29) and in radial contact with the annular wall (25) of the flame tube (24) on the one hand and with the annular flange (35) of the boss (32) on the other.
The invention relates to a device intended for implementation of an anodization treatment of a part, the device comprising: - a treatment chamber comprising a part to be treated and a counter electrode situated opposite the part to be treated, the part to be treated constituting a first wall of the treatment chamber, - a generator, a first terminal of the generator being electrically connected to the part to be treated and a second terminal of the generator being electrically connected to the counter electrode, and - a system for storage and circulation of an electrolyte, the system comprising: a storage tank, different from the treatment chamber, intended to contain the electrolyte, and an electrolyte circulation circuit intended to allow the electrolyte to flow between the storage tank and the treatment chamber.
The invention relates to a turbine engine compressor (10'), comprising at least one annular row of variable-pitch stator blades (12), wherein said blades are substantially radial and comprise pivots (22, 24) at the radial ends thereof, the radially external pivots of the blades being inserted in first openings of a stator housing (20) and the radially internal pivots being inserted in second openings of a floating ring (60) that surrounds a rotor (18) of the compressor, characterized in that an annular stator part (62) is inserted between the floating ring and the compressor rotor, and in that first sealing means are mounted between the stator part and the compressor rotor, and second sealing means are mounted between the floating ring and the stator component.
F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
86.
DEVICE FOR ASSISTING A SOLID PROPELLANT PROPULSION SYSTEM OF A SINGLE-ENGINE HELICOPTER, SINGLE-ENGINE HELICOPTER COMPRISING SUCH A DEVICE AND CORRESPONDING METHOD
The invention concerns a device for assisting a propulsion system of a single-engine helicopter comprising an engine linked to a power transmission unit (15) suitable for rotating a rotor of the helicopter, characterised in that it comprises: a turbine (18) for driving the rotation of an output shaft (34) linked mechanically to said power transmission unit (15); and controlled means (16) for supplying said turbine (18) with pressurised fluid in order to allow said turbine (18) to convert the energy from said pressurised fluid into mechanical energy for rotating said output shaft (34).
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
87.
TURBINE ENGINE COMPRESSOR WITH VARIABLE-PITCH BLADES
Compresseur (10') de turbomachine, comportant au moins une rangée annulaire d'aubes de stator (12) à calage variable, ces aubes étant sensiblement radiales et comportant à leurs extrémités radiales des pivots (22, 24), les pivots radialement externes des aubes étant logés dans des premiers orifices d'un carter de stator (20) et les pivots radialement internes étant logés dans des seconds orifices d'un anneau flottant (60) qui entoure un rotor (18) du compresseur, caractérisé en ce qu'une pièce annulaire (62) de stator est intercalée entre l'anneau flottant et le rotor du compresseur, et en ce que des premiers moyens d'étanchéité sont montés entre la pièce de stator et le rotor du compresseur, et des seconds moyens d'étanchéité sont montés entre l'anneau flottant et la pièce de stator.
Assistance device (100) for an aircraft free turbine turbomachine (TAG1), the device comprising means (K11, CVS1) for electrically powering a coil (SU) of a rotary starter machine (GD1), referred to as first coil, in order to provide a first acceleration assistance to the gas generator of the turbomachine (TAG1). The device further comprises monitoring means (110) for monitoring the first assistance and means (K12, CVS2) for electrically powering a second coil (S12) of a rotary machine, so as to provide a second acceleration assistance to said gas generator if the monitoring means (110) observe the first assistance as insufficient.
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 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 9/42 - Control of fuel supply specially adapted for the control of two or more plants simultaneously
89.
METHOD FOR DETECTING A FAILURE OF A FIRST TURBINE ENGINE OF A TWIN-ENGINE HELICOPTER AND FOR OPERATING THE SECOND TURBINE ENGINE, AND CORRESPONDING DEVICE
The invention relates to a method for detecting a failure of a first turbine engine, referred to as faulty engine (4), of a twin-engine helicopter and for operating a second turbine engine, referred to as sound engine (5), each engine (4, 5) comprising protection end stops regulated by a regulating device which define a maximum power speed, characterized in that it comprises: a step (10) of detecting a fault index of said faulty engine (4); a step (11) of modifying said end stops for protecting said sound engine (5) to protection end stops corresponding to a maximum power single-engine speed, in the event of a fault index being detected; a step (12) of confirming a fault with said faulty engine (4); a step (13) of commanding an increase in the rate of supply of fuel to said sound engine (5) in the event of the fault being confirmed.
B64D 31/12 - Initiating means actuated automatically for equalising or synchronising power plants
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 9/56 - Control of fuel supply conjointly with another control of the plant with power transmission control
F02K 3/12 - Plants including a gas turbine driving a compressor or a ducted fan characterised by having more than one gas turbine
90.
TURBINE ENGINE RAPID REACTIVATION METHOD AND SYSTEM
The aircraft turbine engine rapid reactivation system includes an electrical machine (60) supplied with DC power by an onboard electrical power supply network (10). The system also includes: - a switch (50) positioned between the onboard network (10) and the electrical machine (60); - an additional assembly (30) including a plurality N of electrical power storage elements (30a,,30n); and - a control unit (20) for controlling a device (40) for discharging the storage elements. Said control unit is suitable for placing the onboard network (10) in parallel with a series circuit including at least a portion of the N electrical power storage elements (30a,,30n) such that, when the rapid reactivation system is in operation, the electrical machine (60) is supplied with power at a voltage level above that of the nominal characteristics thereof.
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 6/20 - Adaptations of gas-turbine plants for driving vehicles
Assistance device (100) for a free turbine turbomachine (TAG1) of an aircraft comprising at least two free turbine turbomachines (TAG1, TAG2), the device comprising an electric starter machine (D1) for providing prolonged assistance to the gas generator of a first turbomachine (TAG1) using energy produced by an electric generator machine (G2) driven by the second turbomachine (TAG2), the device additionally comprising at least one electrical storage member (S1) electrically connected to said electric starter machine (D1) to provide short-term assistance to said gas generator, in which the electric starter machine (D1) is powered by a first power converter (CVS1) that allows it to exchange energy with the storage member (S1) for the short-term assistance, and that transmits to it energy supplied by a second power converter (CVS2) for the prolonged assistance.
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 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
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
92.
HYDRAULIC DEVICE FOR EMERGENCY STARTING A TURBINE ENGINE, PROPULSION SYSTEM OF A MULTI-ENGINE HELICOPTER PROVIDED WITH ONE SUCH DEVICE, AND CORRESPONDING HELICOPTER
The invention relates to an emergency start device for a helicopter turbine engine (6). Said device is characterized in that it includes: - a hydraulic engine (7) mechanically connected to said turbine engine (6); - a hydropneumatic storage device (9) connected to said hydraulic engine (7) via a hydraulic system (10) for supplying pressurized liquid to said hydraulic engine (7); and - a controlled quick-opening hydraulic valve (11) arranged on the hydraulic system (10), between said storage device (9) and said hydraulic engine (7), and suitable for being placed, on command, at least in an open position, wherein the liquid can be supplied to said hydraulic engine (7), or in a closed position, wherein said hydraulic engine (7) is no longer supplied with pressurized liquid.
F01D 19/00 - Starting of machines or engines; Regulating, controlling, or safety means in connection therewith
F02C 6/14 - Gas-turbine plants having means for storing energy, e.g. for meeting peak loads
93.
METHOD FOR ASSISTING A TURBOSHAFT ENGINE IN STANDBY OF A MULTI-ENGINE HELICOPTER AND ARCHITECTURE OF A PROPULSION SYSTEM OF A HELICOPTER COMPRISING AT LEAST ONE TURBOSHAFT ENGINE THAT CAN BE IN STANDBY
The invention concerns an architecture of a propulsion system of a multi-engine helicopter comprising turboshaft engines (5, 6), characterised in that it comprises: at least one hybrid turboshaft engine (5), capable of operating in at least one standby mode during a stabilised flight of the helicopter, the other turboshaft engines (6) operating alone during said stabilised flight; an air turbine (30) linked mechanically to the gas generator (17) of the hybrid turboshaft engine (5) and suitable for rotating said gas generator (17); means for collecting pressurised air on the gas generator (27) of a turboshaft engine (6) that is running; and a conduit (31) for conveying said collected air to said air turbine (30).
F02C 6/02 - Plural gas-turbine plants having a common power output
F02C 6/08 - Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas the gas being bled from the gas-turbine compressor
F02C 7/277 - Mechanical drives the starter being a turbine
94.
METHOD OF OPTIMIZED GLOBAL MANAGEMENT OF AN ENERGY NETWORK OF AN AIRCRAFT AND CORRESPONDING DEVICE
The invention relates to a method and a device for optimized global management of an energy network of an aircraft comprising a plurality of energy-related equipment, characterized in that it comprises a module 40 for selecting at least one optimization objective (19) from among a plurality of predetermined objectives, a module (42) for receiving data from equipment, a module (41) for receiving aircraft data, and a module (43) for determining operating settings (22) of the energy-related equipment on the basis of the data (21) of equipment and of aircraft data (20) that are adapted to achieve at least one selected optimization objective (19).
G05B 13/02 - Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
B64C 19/00 - Aircraft control not otherwise provided for
95.
ARCHITECTURE OF A MULTIPLE-ENGINE HELICOPTER PROPULSION SYSTEM, AND CORRESPONDING HELICOPTER
The invention relates to an architecture of a multiple-engine helicopter propulsion system comprising turbine engines (1, 2) which are connected to a power transmission gearbox (3), comprising: a hybrid turbine engine (1) able to operate in at least one standby state during steady stabilized flight of the helicopter; a pack (5, 6) with a rapid restarting of said hybrid turbine engine (1) to cause it to exit said standby state and attain a nominal operating state; an auxiliary power unit (11) connected to said electrotechnical restart pack (5, 6) via a first (10) AC/DC converter designed, on demand, to supply said restart pack (5, 6) with the power required to cause said hybrid turbine engine (1) to leave said standby state.
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 9/42 - Control of fuel supply specially adapted for the control of two or more plants simultaneously
96.
TURBOSHAFT ENGINE, TWIN-ENGINE HELICOPTER EQUIPPED WITH SUCH A TURBOSHAFT ENGINE, AND METHOD FOR OPTIMISING THE ZERO-POWER SUPER-IDLE SPEED OF SUCH A TWIN-ENGINE HELICOPTER
The invention relates to a turboshaft engine comprising a gas generator (5) designed to be rotated and a free turbine (6) rotated by the gases from the gas generator (5). The turboshaft engine is characterised in that it comprises a device (20) for the spontaneous mechanical coupling of the gas generator (5) and the free turbine (6), said device being adapted to connect the gas generator (5) and the free turbine (6) in a spontaneous and mechanical manner once the ratio of the rotation speed (NGG) of the gas generator to the rotation speed (NTL) of the free turbine reaches a pre-determined threshold value.
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 3/113 - 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
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
97.
ARCHITECTURE OF A MULTI-ENGINE HELICOPTER PROPULSION SYSTEM AND CORRESPONDING HELICOPTER
The invention relates to an architecture of a multi-engine helicopter propulsion system comprising turbine engines (1, 2) connected to a power transmission gearbox (3), and a low-voltage on-board DC network (7) intended to power helicopter equipment during flight, characterized in that the architecture comprises: a hybrid turbine engine (1) able to operate in at least one standby state during steady stabilized flight of the helicopter; an electrotechnical pack (20) for the rapid restart of said hybrid turbine engine in order to cause it to exit said standby state and attain a state in which it supplies mechanical power, said restart pack (20) being connected to said onboard network (7); and at least two sources (4, 16, 18) for electrically powering said on-board network (7).
B64D 35/08 - Transmitting power from power plant to propellers or rotors; Arrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 9/42 - Control of fuel supply specially adapted for the control of two or more plants simultaneously
98.
TURBOSHAFT ENGINE COMPRISING A CONTROLLED MECHANICAL COUPLING DEVICE, HELICOPTER EQUIPPED WITH SUCH A TURBOSHAFT ENGINE, AND METHOD FOR OPTIMISING THE ZERO-POWER SUPER-IDLE SPEED OF SUCH A HELICOPTER
The invention relates to a turboshaft engine comprising a gas generator (5) designed to be rotated and a free turbine (6) rotated by the gases from the gas generator. The turboshaft engine is characterised in that it comprises a device (40) for the controlled mechanical coupling of the gas generator (5) and the free turbine (6), said device being designed to connect the gas generator (5) and the free turbine (6) mechanically, on command, once the rotation speed of the gas generator (5) reaches a pre-determined threshold speed.
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
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 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
99.
TRANSMISSION ASSEMBLY FOR AN AIRCRAFT AND A HELICOPTER
A transmission assembly for an aircraft and a helicopter comprising such an assembly for independently adjusting the speed of rotation of an engine and a torque receiver. According to the invention, the assembly comprises a first input shaft (10a), configured to receive mechanical torque from a first engine (10), an output shaft (60a), configured to transmit mechanical torque to a torque receiver (60), a first transmission member (20) having at least two degrees of mobility comprising first, second and third movable parts, a first regulating reversible electric machine (30), and a first balancing reversible electric machine (40), in which the input shaft (10a) is coupled to the first movable part, the output shaft (60a) is coupled to the second movable part, the first regulating electric machine (30) is coupled to the third movable part, and the first balancing electric machine (40) is coupled in series to the input shaft or the output shaft (60a).
F16H 3/72 - Toothed gearings for conveying rotary motion with variable gear ratio or for reversing rotary motion using gears having orbital motion with a secondary drive, e.g. regulating motor, in order to vary speed continuously
The invention relates to an annular turbine engine combustion chamber having an axial direction (X), a radial direction, and an azimuthal direction. Said chamber includes a first annular wall (12) and a second annular wall (14). Each wall defines at least part of the enclosure of the annular combustion chamber. The first and second walls (12, 14) have complementary interlocking elements (12d, 14d). The first wall (12) has at least one first through-hole (12f), while the second wall (14) has at least one second through-hole (14f). The combustion chamber also includes at least one pin inserted into a pair of holes that include a first hole (12f) and a second hole (14f). Said pin (18) is formed by an injector and securely interlocks the first and second walls (12, 14).
patents
