The invention relates to a manifold (20) comprising: - a chamber (21) defining an internal cavity (27); - at least one first inlet (22, 23); - a first outlet (24); - a ventilation pipe (29) comprising: - a ventilation air inlet (30) located outside the chamber (21); and - a ventilation air outlet (31) which is connected to the first liquid outlet (24); - a liquid discharge pipe (33) configured to discharge the liquid from the internal cavity (27) out of the chamber (21) when the liquid level in the chamber (21) is higher than or equal to a threshold level (S1) of liquid in the internal cavity (27), the discharge pipe (33) comprising: - a second inlet (34); and - a second outlet (35), the ventilation and liquid discharge pipes (29, 33) being independent of one another.
The invention relates to a turbine engine (1) comprising: - a nozzle (70) comprising a casing (71); - at least one drainage circuit (C1) for draining at least one liquid; - a manifold (20) comprising: - a chamber (21) defining an internal cavity (27); - a first inlet (22, 23) (27) connected to the first drainage circuit (C1); - a first outlet (24); - a suction duct (42) connected to the casing (71) and to the manifold (20), the suction duct (42) comprising: - a first open end (43) connected to the casing (71) and leading into the gas flow path (v1); and - a second end (44) connected to the first outlet (24); - an opening leading into the gas flow path (v1) and connected to the first end (43) of the suction duct (42), wherein the opening (50) is configured to create a negative pressure in the suction duct (42).
The invention relates to a manifold (20) comprising: - at least one first inlet (22, 23); - a first outlet (24); - a liquid discharge pipe (33) for discharging the liquid out of the enclosure (21) when the liquid level in the internal cavity (27) is greater than or equal to a first threshold level (S1) of liquid in the internal cavity (27); - a signalling pipe (37) located in the enclosure (21) and configured to collect liquid when the liquid level in the internal cavity (27) is greater than or equal to a second threshold level (S2) of liquid in the internal cavity (27), which is lower than the first threshold level (S1); - a device (40) for detecting and signalling that the liquid level in the internal cavity (27) has reached one and/or the other of the first and second threshold levels (S1, S2).
A hybrid electrical architecture for an aircraft includes a turbomachine and a reduction gearbox configured to rotate at least one propulsion member of the aircraft. The architecture also includes at least one low-voltage electrical network with at least one low-voltage electric machine mounted on the reduction gearbox or on the turbomachine and a high-voltage electrical network with high-voltage electric machines mounted on the turbomachine.
B64D 31/18 - Power plant control systemsArrangement of power plant control systems in aircraft for electric power plants for hybrid-electric power plants
B64D 31/14 - Transmitting means between initiating means and power plants
5.
TURBOMACHINE BLADE, TURBOMACHINE AND METHOD FOR MANUFACTURING THE BLADE
Turbomachine blade including an external shell including a pressure-side wall and a suction-side wall delimiting between them an interior volume, an insert arranged in the interior volume so as to form an air passage between the insert and the external shell; and at least one breakable joining portion, connected on the one hand to the insert and on the other hand to the external shell; wherein one at least of the mechanical breaking resistance of the breakable joining portion, the mechanical breaking resistance between the breakable joining portion and the external shell and the mechanical breaking resistance between the breakable joining portion and the insert is less than the mechanical breaking resistance of the external shell and the mechanical breaking resistance of the insert.
A collector for a drained liquid for an aircraft turbine engine includes an internal cavity having a first space for collecting the drained liquid and a second space for transferring the collected liquid to a recovery outlet; at least one inlet for the drained liquid, in fluid communication with the first space; and at least one recovery outlet, in fluid communication with the second space. The first space and the second space are separated from each other by a partition in which is arranged a means for restricting the passage of the drained liquid from the first space to the second space, the air in the first space being in communication with the air in the second space such that the air pressure in the first and second spaces is identical.
The invention relates to a method for assisting with propulsion by detecting a failure of a turboshaft engine of an aircraft operating in a nominal mode, the aircraft comprising a deactivated assistance motor, the method comprising: a step (2) of comparing operating parameters of the turboshaft engine with the equivalent parameters of a model representative of a healthy turboshaft engine; a step (4) of detecting a failure of the turboshaft engine by detecting an anomaly in at least one operating parameter of the turboshaft engine; a step (6) of selecting an activation mode of the assistance motor on the basis of the operating parameters of the turboshaft engine and/or of flight parameters of the aircraft; and a step (8) of activating the assistance motor with the selected activation mode.
The invention relates to a piloting assist method for an operator, for example a pilot, or for an automated system, for example an autopilot, the assist relating to at least one cumulative ageing process (CV), for example creep, in at least one portion of a gas turbine of an aircraft during a mission of the aircraft, wherein the method comprises a step of determining an ageing progression score that is a function of a plurality of terms comprising TERM 1, TERM 2, TERM 3, TERM 4 and TERM 5, and wherein the method comprises a step of supplying information on the at least one cumulative ageing process to the operator or to the automated system for the piloting assist, on the basis of the ageing progression score.
G07C 5/08 - Registering or indicating performance data other than driving, working, idle, or waiting time, with or without registering driving, working, idle, or waiting time
9.
PILOT ASSIST METHOD FOR MANAGING AT LEAST ONE CUMULATIVE AGEING PROCESS
The invention relates to a pilot assist method for an operator, for example a pilot, or for an automated system, for example an autopilot, relating to at least one cumulative ageing process, for example creep, in at least one portion of a gas turbine of an aircraft during a mission of the aircraft, wherein the method comprises a step of determining an ageing progression score (SCORE) that is a function of a plurality of terms, the plurality of terms comprising TERM1, TERM2 and TERM3, and wherein the method comprises a step of supplying information on the at least one cumulative ageing process to the operator or to the automated system for the pilot assist, on the basis of the ageing progression score (SCORE).
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE TOULOUSE III - PAUL SABATIER (France)
Inventor
Brunet, Clément, Pierre
Richard, Stéphane, Raphaël, Yves
Marragou, Sylvain, Mathieu, Julien
Schuller, Thierry
Magnes, Hervé
Abstract
The present disclosure relates to an injection system (20) for injecting a mixture of hydrogen and air along a first axis (X1) into an annular combustion chamber (4) of a turbine engine (1), the injection system (20) comprising: - a radially inner channel for the flow of hydrogen; - an annular radially outer channel for the flow of air; - an inner swirler being housed in the radially inner channel; - an outer swirler being housed in the annular radially outer channel, wherein a downstream end of the radially inner channel is arranged upstream along the first axis of a downstream end of the annular radially outer channel, the annular radially outer channel having an air flow section, the section being convergent along the first axis (X1) at least from a downstream end of the inner swirler or a downstream end of the outer swirler to the downstream end of the radially inner channel.
A hybrid propulsion system including a turbine engine including a high-pressure spool and a low-pressure spool, the low-pressure spool including reduction gear, the reduction gear forming part of a transmission gearbox which is positioned axially at a front end of the propulsion system; —first and second electric machines mechanically and respectively connected to the high-pressure and low-pressure spools, the electric machines being configured to operate in modes referred to as motor and generator, the first and second electric machines being fixed to the transmission gearbox; —a control system which is configured to allow the transfer of power between the high-pressure and low-pressure spools via the first and second electric machines.
B64D 35/022 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants for electric power plants of hybrid-electric type
09 - Scientific and electric apparatus and instruments
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
Computer programs and software providing information
regarding aircraft engines, operation, maintenance,
intervention, repair in relation to aircraft engines and/or
their modules and parts; content management software; data
banks on digital media; data banks on digital media
containing information concerning aircraft engines, their
operation, repair, overhaul, servicing, maintenance and
reconditioning of engines, modules and parts of aircraft
engines; all these goods used and/or intended for the
aeronautical field. Provision of information with respect to repair and
maintenance of aircraft engines and their components. Provision of technical documentation relating to aircraft
engines, their operation, repair, overhaul, servicing,
maintenance; provision of information online concerning
industrial analysis and research services; provision of
technical information in the aeronautical field; development
(creation) services for data banks; development (design),
installation, maintenance and updating of computer software
in the aeronautical field; rental of computer software in
the aeronautical field; scientific and technological
services and related research and design services, in the
aeronautical field, namely: surveying (engineering work);
analysis, expert reports and processing of the acquisition
for data recorded during the operation of aircraft engines,
modules and parts thereof; technical documentation provision
services online; online hosting of databases containing
technical documentation; electronic data storage; all these
services used and/or intended for the aeronautical field.
13.
AIRCRAFT TURBOMACHINE COMPRISING A BEARING SUPPORT HAVING AN IMPROVED DESIGN
An assembly for aircraft turbomachine, including a stator part, a first bearing, a second bearing and a support part in an oil enclosure delimited by an outer enclosure delimiting portion integral with the stator part, the support part including a first axial end portion forming an outer ring of the first bearing, or supporting such a ring; a second axial end portion forming an outer ring of the second bearing or supporting such a ring, a compression damping oil film being arranged between the second portion and the stator part; an intermediate ring, arranged axially between the first and second portions, and forming a flexible connection as well as an oil-spray protection element.
A propulsion assembly for a hybrid aircraft including a first and a second engine each having a gas generator and a free turbine, a main rotor coupled to the free turbine of the first and second engines, the first engine including a first electric machine and a second electric machine of lower power than the first electric machine, one of the first or of the second electric machine being able to be coupled to the gas generator and to set the gas generator in rotation during a start phase of the engine, and being further able to be coupled to the free turbine in order to generate electrical energy after the start phase, the other of the first or of the second electric machine being coupled to the gas generator only.
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
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 6/02 - Plural gas-turbine plants having a common power output
The invention relates to an assembly (100) for an aircraft turbine engine (1), the assembly comprising: - a first turbine engine part (20) and a second turbine engine part (22), wherein each of the parts is rotatable about a longitudinal central axis (X) of the assembly; - a main device (24) for mechanically coupling the first part (20) to the second part (22), wherein the main device is configured, in a normal operating configuration of the assembly, to mechanically couple the first part (20) to the second part (22) in translation along the longitudinal central axis (X) of the assembly; - an emergency device (124) for mechanically coupling the first part (20) to the second part (22), wherein the emergency device (124) comprises a shaft (126) that comprises: - a first coupling portion (128) fixed to or integrated into the first part (20); - a second coupling portion (132) comprising at least one emergency axial stop (134), and, in the normal operating configuration of the assembly, in which the emergency device is in an inactive axial coupling state, the emergency axial stop (134) is axially spaced from a complementary axial stop (136) provided on the second part (22); wherein the assembly is configured such that, in the event of a failure resulting in an undesired axial separation between the first and second parts (20, 22), the emergency device (124) switches to an active axial coupling state in which axial retention of one of the first and second parts relative to the other is ensured by the emergency axial stop (134) coming into contact with the complementary axial stop (136).
F01D 5/06 - Rotors for more than one axial stage, e.g. of drum or multiple-disc typeDetails thereof, e.g. shafts, shaft connections
F01D 21/04 - Shutting-down of machines or engines, e.g. in emergencyRegulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator, e.g. indicating such position
16.
Lubrication and cooling of equipment of an aircraft turbomachine
An aircraft turbomachine includes a gas generator having an output shaft and a first lubricating circuit. The turbomachine further includes equipment coupled to the output shaft and including a rotor rotationally guided by at least one rolling bearing and a second lubricating circuit that is independent of the first lubricating circuit and is configured to lubricate the rolling bearing. The equipment further includes a system for cooling the rolling bearing, the cooling system being configured to circulate oil in the region of at least one ring of the rolling bearing. The cooling system is independent of the second lubricating circuit and is connected to the first lubricating circuit.
The present invention relates to a computer-implemented method for controlling the flow rate of fuel injected into a combustion chamber (42) of a turbine engine (4) of a helicopter (1) by means of a fuel metering device (50), the method comprising the following steps: - calculating a fuel flow rate setpoint in accordance with a desired speed of the turbine engine (4); - calculating a virtual fuel flow rate corresponding to an increase or decrease in the temperature in the combustion chamber according to a thermal model (M, M') of the turbine engine comprising an intrinsic thermal model of the turbine engine influencing the temperature in the combustion chamber; - correcting the fuel flow rate setpoint of the virtual fuel flow rate; - controlling the fuel metering device so that it injects an amount of fuel that complies with the corrected fuel flow rate setpoint (C2) into the combustion chamber.
F02C 1/06 - Gas-turbine plants characterised by the use of hot gases or unheated pressurised gases, as the working fluid the working fluid being heated indirectly characterised by the type or source of heat, e.g. using nuclear or solar energy using reheated exhaust gas
F02C 7/14 - Cooling of plants of fluids in the plant
F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
F02C 7/08 - Heating air supply before combustion, e.g. by exhaust gases
18.
METHOD FOR QUANTIFYING BURNS GENERATED BY GRINDING A PART
The invention relates to a method (100) for quantifying burns generated by grinding a part (10), the method comprising at least: a step (110) of acquiring images Ii, i being an integer, of at least one zone (S1) of the part (10) to be observed on a background (12) by way of an apparatus (20) comprising an observation device (22) equipped with an image sensor (24) and a light source (26) able to illuminate the part (10), said acquisition apparatus (20) being configured to acquire images in N grayscale levels k, k = [0...N], N being an integer, each grayscale level being associated with a grayscale value gk, each image Ii comprising a first series of pixels associated with the part (10) and a second series of pixels associated with the background (12), a step (120) of processing the images acquired in the image acquisition step (110), and a step (130) of computing the level of surface burns ρ on the part (10).
The invention relates to an aircraft turbine engine (10) comprising: - a centrifugal compressor (14); - a combustion chamber (24); - a turbine (30) comprising a vaned nozzle diaphragm (31) mounted at the outlet of the chamber (24) and a vaned wheel (33) located downstream of this nozzle diaphragm (31), wherein the vaned wheel (33) is surrounded by a sealing ring (35) carried by an annular support (37) connected to the casing (29); - a system (32) for diffusing and straightening an air flow exiting the centrifugal compressor (14) in order to supply the combustion chamber (24); - a heat exchanger (38) connected by scroll arrangements (40) to the diffusion and straightening system (32); and - a circuit (50) for bleeding air and circulating the drawn-in air to the distributor (31) and/or to the sealing ring (35) for cooling.
The present invention relates to a method for realigning a digital twin with a part on which a visual marker is arranged, configured to allow angular positioning of the part, the method comprising: - an exposure of the part to at least one light source in the ultraviolet range, - at least one capture of the part on which the visual marker is arranged, and exposed under the light source, - a realignment implemented by computer, under the ultraviolet light, of the captured part with a digital twin of the part, on the basis of the visual marker present on the part, configured to be visible under the ultraviolet light, using a fluorescent penetrant fluid, and a digital marker present on the digital twin, the geometric shape of the visual marker positioned on the part allowing the realignment between the part and its digital twin.
Turbine engine (10) for an aircraft, comprising: - a centrifugal compressor (12), - a flow straightener-diffuser assembly (28), - an annular combustion chamber (14), - a turbine annular distributor (36), - a turbine wheel (22), and - an annular sealing system (48) connected to the distributor (36) by a first annular flange (50) and, on the other side, to the flow straightener-diffuser system (48) by a second annular flange (52), characterized in that the second flange (52) comprises an annular wall (76) which extends perpendicularly to the axis (A) and which gives this flange a certain flexibility.
A propulsive assembly including a first and a second gas turbine each having a gas generator and a free power turbine, a main rotor coupled to the free power turbine of a first and a second main coupling, a first and a second reversible electric machine each coupled to the gas generator, by way of a first deactivatable coupling, and each coupled to the main rotor by way of a second deactivatable coupling, the first deactivatable coupling being activated when the electric machines are rotating in a first direction of rotation, and the second deactivatable coupling being activated when the electric machines are rotating in a second direction of rotation opposite to the first direction of rotation.
B64D 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
B64D 35/023 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants for electric power plants of hybrid-electric type of series-parallel type
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
F02C 3/14 - 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 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
A turbomachine for hybrid aircraft, including a gas generator carried by a generator shaft, at least one free turbine carried by a turbine shaft and driven in rotation by a gas stream generated by the gas generator, a main rotor, and at least one reversible electric machine, the turbine shaft being a through shaft and extending axially between a first end engaged with the electric machine, and a second end engaged with the main rotor.
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
F02C 3/14 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
24.
VENTURI DRAINAGE ASSEMBLY AND METHOD FOR A TURBINE ENGINE
The invention relates to a drainage assembly (15) comprising a nozzle arm (11) contained in a nozzle (9) through which combustion gases flow, a drain collector (16) for collecting the fluids drained from the turbine engine (1), a Venturi ejector (17) having a primary channel (20) with a constriction (23) and a suction channel (26), and a duct (18) for connecting the drain collector to the inlet of the suction channel. The Venturi ejector is in the nozzle arm (11). The inlet and the outlet of the primary channel open into the nozzle such that a portion of the combustion gases passes through the primary channel and serves as the motive fluid for the ejector. The inlet of the suction channel is located on the side of the end of the nozzle arm that is directed towards the wall (13) of the nozzle. The inlet (21) of the primary channel (20) has a cross-section whose width corresponds to at least 70% of the maximum width (l) of the nozzle arm (11) that encloses the Venturi ejector (17).
A standard part (1, 100, 200) intended to be used as a standard for a non-destructive test or to measure the performance of the non-destructive test, the standard part being obtained by additive manufacturing, the standard part comprising a defect formed in a controlled manner.
B33Y 80/00 - Products made by additive manufacturing
B22F 10/00 - Additive manufacturing of workpieces or articles from metallic powder
B29C 64/00 - Additive manufacturing, i.e. manufacturing of three-dimensional [3D] objects by additive deposition, additive agglomeration or additive layering, e.g. by 3D printing, stereolithography or selective laser sintering
B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
26.
METHOD FOR ASSEMBLING METAL PIECES OF DIFFERENT MASSIVENESS AND CENTRIFUGAL DIFFUSER PRODUCED BY THIS METHOD
A method for assembling a first metal part to a second metal part, the first and second metal parts having different sizes, the method including the following operations a) producing a slot in a surface of the first metal part; b) positioning the second metal part in line with the slot of the first metal part; and c) welding the second metal part to the first metal part through the slot using a high-energy welding beam, the slot guiding the welding beam.
The invention relates to a device for characterizing at least one light beam (7), comprising: - a carrier board (1) comprising a plurality of measurement points, each measurement point being equipped with at least one sensor (4i) for measuring at least one energy-related and/or spatial and/or temporal parameter of a beam incident on the sensor and slave digital means (6i) for processing and/or storing data of the sensor; - master digital means (6) for storing data of the sensors of the various measurement points; - a plurality of digital buses (8i), each connecting one of the slave digital means (6i) to the master digital means (6), the various digital buses being parallel to one another, the slave digital means (6i) being able to, or programmed to, transmit measurement data of the sensors to the master digital means (6).
METHOD FOR COMPLETING THE FINISHING OF A COMPONENT PRODUCED BY DEPOSITING AND SOLIDIFYING SUCCESSIVE LAYERS OF A POWDER, AND COMPONENT OBTAINED BY THIS METHOD
A method for completing the finishing of a component produced by depositing and solidifying successive layers of a powder, characterized in that it comprises: - a rough-finishing step (E1) involving sandblasting a part of a surface of the component having an initial roughness (Ra0) using a first sand (S1) having a rough-finishing particle size (G1) and blasted at a rough-finishing sandblasting angle (A1) and a rough-finishing sandblasting pressure (P1) until the part of the surface of the component attains a rough-finished roughness (Ra1), and - a finishing step (E2, E2') of finishing the part of the surface of the component until the part of the surface of the component attains a final roughness (Ra2), where the final roughness (Ra2) is lower than the rough-finished roughness (Ra1).
A process for three-dimensional printing of a workpiece including a succession of steps for producing a layer of the workpiece by means of a nozzle which has an output cross-section along a path of the nozzle, wherein, during at least one step for producing a layer, use is made of at least one nozzle, an output portion of which has an output cross-section which is variable between a first maximum cross-section and a second minimum cross-section, and wherein the cross-section of at least one part of an output portion of the nozzle is varied along at least one portion of the path.
B29C 64/118 - Processes of additive manufacturing using only liquids or viscous materials, e.g. depositing a continuous bead of viscous material using filamentary material being melted, e.g. fused deposition modelling [FDM]
A filtering device for a turbomachine, includes a filter support; a removable filter cooperating with the filter support; a cover; an indicator pin; a plate that is translatably movable relative to the cover and relative to the indicator pin, the plate being translatably movable between two positions: a first position such that the plate bears against the removable filter and the indicator pin is in a retracted position, a second position such that the plate is remote from the cover due to the absence of the removable filter and the indicator pin is in an extended position.
One aspect of the invention relates to a test specimen (100) for characterising a vibrational damping of an adhesive or sliding contact between two parts subjected to vibrational, oligocyclic and tribological stresses, wherein the specimen is provided with a male-female fastener (130) comprising a male part (131) that projects from the free end of a first longitudinal bar and a female part (132) that is indented at the free end of a second longitudinal bar. A second aspect of the invention relates to a device (300) for characterising a vibrational damping of an adhesive or sliding contact between two parts subjected to vibrational, oligocyclic and tribological stresses, wherein the device is provided with a test bench (300) on which the test specimen (100) is mounted, and wherein the device comprises a vibrator (310) that transmits vibrational excitations to the test specimen (100) and a static loading device (320) that transmits olygocyclic stresses to the test specimen. A third aspect of the invention relates to a method (400) for implementing the characterising device (300).
The invention relates to a method for measuring the crystal misorientation of a single-crystal blade (10) comprising a material formed of dendrites (30) having a main trunk (32) extending in a main direction (Z') and branches (34), the method comprising: defining a reference coordinate system (XYZ) for the blade comprising a radial direction (Z); mechanically and chemically treating the blade (10) in a base plane (P1) in order to acquire a first image (I1) by micrography; mechanically and chemically treating the blade (10) in a comparison plane (P2) radially offset from the base plane (P1) by a distance (D), in order to acquire a second image (I2) by micrography; and processing the first and the second image (I1, I2) so as to compare the position of the main trunk (32) of each dendrite (30) in the two images (I1, I2) and thus deduce the orientation of the main direction (Z') of the dendrites (30) relative to the radial direction (Z).
The system (100) includes a control system (124) configured to: - in response to receipt of a quick-start instruction, accelerate a first output shaft (116A) disengaged from a first turboshaft engine (102A); - during acceleration, not apply any limiting of the acceleration setpoint for the first output shaft (116A); - calculate and monitor an estimate of a time-to-clutch-engagement; - when the estimate of the time-to-clutch-engagement drops below a predefined threshold, reduce the acceleration of the first output shaft (116A); then - prior to clutch-engagement, apply limiting of the acceleration setpoint for the first output shaft (116A).
The invention relates to a diffuser (1) of an aircraft turbine engine (60) extending along a longitudinal axis (Z) oriented vertically in the opposite direction to gravity and comprising an upstream flange and a downstream flange that together define a flow duct for an air flow (F), wherein the flow duct is configured to be supplied with fluid by a centrifugal impeller (21) and to open into a peripheral chamber (30) surrounding a combustion chamber (31) of the aircraft turbine engine (60), wherein the downstream flange is configured to define a bottom of the peripheral chamber (30), and wherein the diffuser (1) comprises at least one drainage channel (10) passing through the downstream flange, the flow duct and the upstream flange, wherein the drainage channel (10) comprises a downstream end configured to open into the peripheral chamber (30) and an upstream end that leads out of the flow duct, thereby enabling a liquid (L) to be drained from the peripheral chamber (30) by gravity.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
UNIVERSITE DE LORRAINE (France)
Inventor
Barraco, Thomas Michel Andre Gerard
Klonowski, Thomas
Pierfederici, Serge Lionel
Weber, Mathieu Robert
Abstract
A DC-to-DC converter for an electrical aircraft propulsion system designed to be connected in series with an electrical energy storage unit of the electrical propulsion system. The DC-to-DC converter includes an inverter, a transformer and a rectifier, and further includes a current source that is connected to the rectifier and is configured to control the power passing through the DC-to-DC converter. The transformer includes a primary and two secondaries, the two secondaries sharing a common terminal designed to be connected to a high-voltage DC bus of the electrical propulsion system and two other terminals that are connected to the rectifier. The rectifier includes two arms including at least two transistors that are each in series and are connected, on the one hand, to the two other terminals of the transformer and, on the other hand, to the current source.
H02M 3/335 - Conversion of DC power input into DC power output with intermediate conversion into AC by static converters using discharge tubes with control electrode or semiconductor devices with control electrode to produce the intermediate AC using devices of a triode or a transistor type requiring continuous application of a control signal using semiconductor devices only
A turbine engine module, in particular an aircraft turbine engine, including an annular casing having an internal wall forming a channel wall; and a nozzle surrounded by the casing and including an annular external platform and an annular internal platform between which stator blades extend, the external platform having an external face that faces the internal wall of the casing and includes an annular groove oriented towards the outside and housing a sealing device, the sealing device coming into cylindrical contact with a track of the internal wall of the casing, wherein the internal wall of the casing includes a thermal barrier made of ceramic material directly above the track, the track being arranged between the thermal barrier and the sealing device.
Turbine engine subassembly (10) intended for a helicopter, comprising a primary outlet nozzle (12) for discharging an exhaust gas stream (F1), which comprises an outer casing (14) and an inner cone (16), these being arranged concentrically around a longitudinal axis (X) and connected to one another by radial arms (18), the casing (14) surrounding the cone (16) and defining between them an annular duct (C) for the flow of the stream (F1). The subassembly comprises a distributor (20) surrounding the casing (14) and configured so that at its inlet (20a) it receives a cooling-air stream (F2) having a temperature lower than the temperature of the stream (F1) and at its outlet it distributes the stream (F2) to inside the radial arms (18). The inner cone has an outer wall (16a; 16b') in which one or more openings (O) are arranged to allow the air of (F1).
A device for controlling a power-transfer system for the transfer of power between a high-pressure shaft and a low-pressure shaft of a turbomachine of an aircraft, including a fatigue analysis module analysing the fatigue of the turboma-chine and designed to determine, from between two indicators (D1, D2) respectively measuring two fatigues of the turbomachine, which is the one that is the most advanced, which is to say which is the one at risk of being first to reach a respective upper limit (D1max, D2max); and—a control module controlling the power transfer system and designed to slow the fatigue measured by the more advanced indicator (D1, D2).
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F02C 9/28 - Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
F02C 9/56 - Control of fuel supply conjointly with another control of the plant with power transmission control
39.
TRANSMISSION SYSTEM FOR AN AIRCRAFT PROPULSION UNIT HAVING THE CAPABILITY TO UNCOUPLE TWO ROTATING ELEMENTS BY MEANS OF THE FORCE PRODUCED BY HELICAL GEAR MESHING
A transmission system (2) with uncoupling capability for an aircraft propulsion unit (1) comprises: a first and a second rotating element; a pinion (6) mounted so as to rotate about an axis (14); coupling means (40) for coupling the pinion to one of the rotating elements (3); helical gear meshing means (42) for coupling the pinion with the other rotating element (7); an actuator (10). The actuator, on command, adopts: a coupling configuration in which it holds the pinion in a coupling position, in which the coupling means are engaged and the helical gear meshing means are engaged; and a disengagement configuration allowing an axial force (F1) induced by the helical gear meshing means to cause the pinion to move along its axis into a disengagement position, in which the coupling means or the helical gear meshing means are disengaged.
F16H 3/00 - Toothed gearings for conveying rotary motion with variable gear ratio or for reversing rotary motion
F16D 11/10 - Clutches in which the members have interengaging parts actuated by moving a non-rotating part axially with clutching members movable only axially
F16D 25/08 - Fluid-actuated clutches with fluid-actuated member not rotating with a clutching member
F16H 3/22 - Toothed gearings for conveying rotary motion with variable gear ratio or for reversing rotary motion without gears having orbital motion exclusively or essentially using gears that can be moved out of gear with gears shiftable only axially
A turboshaft engine for an air-craft including a gas generator including a compressor, a combustion chamber and an expansion turbine; a power turbine rotating a power take-off by a reduction gear; a heat exchanger including a first circuit and a second circuit. The compressor includes a first shaft rotated by a second shaft of the expansion turbine by a trans-mission mechanism, the transmission mechanism and the reduc-tion gear forming part of a gearbox which is arranged axially at a front end of the turboshaft engine, such that the compressor is arranged axially between the gearbox and the power turbine.
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 7/10 - Heating air supply before combustion, e.g. by exhaust gases by means of regenerative heat-exchangers
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
09 - Scientific and electric apparatus and instruments
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
(1) Logiciels et programmes informatiques fournissant des informations à propos des moteurs d'aéronefs, du fonctionnement, de la maintenance, de l'intervention, de la réparation en relation avec les moteurs d'aéronefs et/ou leurs modules et pièces; logiciels de gestion de contenus; banques de données sur supports numériques; banques de données sur supports numériques contenant des informations concernant les moteurs d'aéronefs, leur fonctionnement, la réparation, la révision, l'entretien, la maintenance et la remise en état de moteurs, modules et pièces de moteurs pour aéronefs; tous ces produits étant utilisés et/ou destinés au domaine aéronautique. (1) Mise à disposition d'informations en matière de réparation et de maintenance des moteurs d'aéronefs et de leurs pièces constitutives.
(2) Mise à disposition de documentation technique relative aux moteurs d'aéronefs, leur fonctionnement, la réparation, la révision, l'entretien, la maintenance; mise à disposition d'informations en ligne concernant des services d'analyse et de recherche industrielle; mise à disposition d'information technique dans le domaine aéronautique; services de développement (constitution) de banques de données; services d'élaboration (conception), installation, maintenance, mise à jour de logiciels d'ordinateurs dans le domaine aéronautique; location de logiciels d'ordinateurs dans le domaine aéronautique; services scientifiques et technologiques et services de recherche et conception y relatifs, dans le domaine aéronautique à savoir: expertises (travaux) d'ingénieurs; service d'analyse, d'expertise et de traitement de l'acquisition de données enregistrées lors du fonctionnement de moteurs d'aéronefs, de leurs modules et pièces; services de mise à disposition en ligne de documentation technique; hébergement de bases de données en ligne contenant de la documentation technique; stockage électronique de données; tous ces services étant utilisés et/ou destinés au domaine aéronautique.
09 - Scientific and electric apparatus and instruments
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
Computer programs and software providing information regarding aircraft engines, operation, maintenance, intervention, repair in relation to aircraft engines and/or their modules and parts; content management software; data banks on digital media; data banks on digital media containing information concerning aircraft engines, their operation, repair, overhaul, servicing, maintenance and reconditioning of engines, modules and parts of aircraft engines; all these goods used and/or intended for the aeronautical field. Provision of information with respect to repair and maintenance of aircraft engines and their components. Provision of technical documentation relating to aircraft engines, their operation, repair, overhaul, servicing, maintenance; provision of information online concerning industrial analysis and research services; provision of technical information in the aeronautical field; development (creation) services for data banks; development (design), installation, maintenance and updating of computer software in the aeronautical field; rental of computer software in the aeronautical field; scientific and technological services and related research and design services, in the aeronautical field, namely: surveying (engineering work); analysis, expert reports and processing of the acquisition for data recorded during the operation of aircraft engines, modules and parts thereof; technical documentation provision services online; online hosting of databases containing technical documentation; electronic data storage; all these services used and/or intended for the aeronautical field.
43.
AIRCRAFT TURBOMACHINE COMPRISING A DEVICE FOR INHIBITING THE ACCUMULATION OF COKE IN A DUCT
An aircraft turbomachine has a gas generator that includes, along a longitudinal axis (X), at least one compressor, a combustion chamber, and at least one turbine. The turbomachine further includes at least one duct for supplying liquid to at least one member chosen from an oil jet and a fuel injector. The duct (20) has rectilinear portions and bent portions and includes at least one region in which the liquid is liable to coke. The turbomachine also includes at least one turbulence element in the at least one region in the duct.
F02C 7/30 - Preventing corrosion in gas-swept spaces
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
44.
METHOD FOR REGULATING THE SPEED OF ROTATION OF A PROPULSION DEVICE OF A HYBRID PROPULSION UNIT FOR AN AIRCRAFT, IN THE EVENT OF A FAILURE OF THE MAIN REGULATION SYSTEM OF THE HEAT ENGINE OF THE HYBRID PROPULSION UNIT
A method for regulating the speed of a propulsion device of an aircraft including: the propulsion device and a gearbox MGB; the heat engine and at least one electric motor, mounted in parallel on the MGB, the heat engine having a fuel circuit; main and backup regulation systems, and a regulation system, each capable of regulating the speed of the heat engine or the electric motor, respectively; a control system of the aircraft, capable of sending a speed or power setpoint to each of the regulation of the heat engine and the electric motor. The method includes: sending a speed setpoint NM2ref to the regulation system of the electric motor, the regulation system sending a power command PM2*, to obtain an instantaneous power PM2m; simultaneously, sending a speed or power command to the backup regulation system of the heat engine, the backup regulation system sending a selected fuel flow command QCarbAux* to the fuel circuit of the heat engine.
B64D 31/18 - Power plant control systemsArrangement of power plant control systems in aircraft for electric power plants for hybrid-electric power plants
B64D 31/14 - Transmitting means between initiating means and power plants
45.
METHOD FOR CHECKING THE MAXIMUM POWER AVAILABLE TO DIFFERENT MEMBERS OF A PROPULSION CHAIN OF AN AIRCRAFT
A method for checking the maximum power available to members of a propulsion system of an aircraft includes first members that are sized to compensate for the failure of second members of the propulsion system by delivering a maximum power to keep the aircraft in a safe operating range. The method includes the following steps for each of the first members: placing the first member in a state that is substantially equal to a maximum power state; adjusting the power delivered by the second member working in synergy with the first member so that the first member and the second member contribute to delivering the power required for the aircraft in the flight phase; determining the power delivered by the first member placed in the maximum power state; from the determined power, deducing information relating to the maximum power available to the first member.
The present invention relates to a metal powder for a powder bed additive manufacturing process, the metal powder comprising a nickel-based alloy comprising at least 0.05% carbon, at least 14.25% cobalt, at least 14% chromium, at least 4% aluminium, at least 3.9% molybdenum, at least 3% titanium, at most 0.5% iron, at least 0.012% boron, at most 0.060% zirconium, at most 0.150% manganese, at most 0.2% silicon, at most 0.1% copper, at most 0.5 ppm bismuth, at most 5 ppm silver, at most 5 ppm lead, at most 25 ppm sulphur, at most 200 ppm oxygen, and at most 60 ppm nitrogen.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
Le Pottier, Nathalie
Cotinot, Jérémie
Frayret, Jérôme
Castetbon, Alain
Pettier, Sophie
Gurt Santanach, Julien
Pommiers Belin, Sébastien
Thielleux, Delphine
Potin Gautier, Martine
Vialas, Nadia
Zoccali, Sandra
Abstract
The invention relates to a method for anticorrosion treatment of a magnesium alloy part, in which the part is immersed for a given duration in an anticorrosion aqueous solution having a given temperature and containing permanganate ions MnO4- and dihydrogen phosphate ions H2PO4-. According to the invention, the anticorrosion aqueous solution contains, prior to immersion of the part, a molar ion concentrate of permanganate ions [MnO4-] of greater than or equal to 0.18 mol/L and less than or equal to 0.32 mol/L, the pH of the anticorrosion aqueous solution is maintained for the given duration at a value of greater than or equal to 3.2 and less than or equal to 4.2, preferably greater than or equal to 3.4 and less than or equal to 4.0, by addition of phosphoric acid, and, after the given duration, the part is removed from the solution and rinsed with water.
C23C 22/73 - Chemical surface treatment of metallic material by reaction of the surface with a reactive liquid, leaving reaction products of surface material in the coating, e.g. conversion coatings, passivation of metals characterised by the process
48.
CONTROL DEVICE FOR CONTROLLING AN AIRFLOW GUIDING SYSTEM, IN PARTICULAR IN AN AIRCRAFT TURBINE ENGINE
A device for controlling an airflow guiding system comprising:
at least one vane movable in rotation about an axis of rotation between a first angle and a second angle,
an actuator comprising a body inside which a piston is mounted in translation secured to an actuation rod, and
a control rod comprising a downstream end connected to the axis of the vane, the actuator being configured to drive the control rod in movement between a first end position and a second end position of a nominal operating range and, in the event of breakdown of said device, to perform an over-stroke of the actuation rod into a safety position.
A device for controlling an airflow guiding system comprising:
at least one vane movable in rotation about an axis of rotation between a first angle and a second angle,
an actuator comprising a body inside which a piston is mounted in translation secured to an actuation rod, and
a control rod comprising a downstream end connected to the axis of the vane, the actuator being configured to drive the control rod in movement between a first end position and a second end position of a nominal operating range and, in the event of breakdown of said device, to perform an over-stroke of the actuation rod into a safety position.
The device comprises a drive mechanism linking an upstream end of the actuation rod to an upstream end of the control rod, opposite to the downstream end.
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 25/24 - CasingsCasing parts, e.g. diaphragms, casing fastenings
49.
METHOD FOR CONTROLLING AN AIRCRAFT PROPULSION SYSTEM HAVING TURBOSHAFT ENGINES OPERATING IN PARALLEL AND CAPABLE OF BEING PLACED ON STANDBY, AND CORRESPONDING AIRCRAFT
122), and at least one electric starter coupled to each turboshaft engine; the propulsion system being designed to have a nominal mode in which the two turboshaft engines drive the rotor and air is bled/electricity is tapped from the two turboshaft engines, and an eco mode in which only the first turboshaft engine drives the rotor, the second turboshaft engine being in standby mode, and the air being bled/electricity being tapped only from the first turboshaft engine, the rotor having a setpoint speed in the two modes, the method comprising: - with the propulsion system in nominal mode, verifying compatibility between aircraft operating parameters and the eco mode; - if compatible and if a command to switch to an eco mode is received, stopping the bleeding of air/tapping of electricity from the second turboshaft engine and verifying compatibility between operating parameters of the first turboshaft engine and the eco mode; - if compatible, idling the second turboshaft engine so that its output shaft has a speed lower than that of the rotor; - verifying correct operation of the starter of the second turboshaft engine; - if correct, placing the second turboshaft engine in standby mode by interrupting a supply of fuel to the second turboshaft engine in order to switch to the eco mode.
The present disclosure relates to a propulsion assembly (4) for an aircraft (1) and to a method for the thermal management of such a propulsion assembly (4). The propulsion assembly (4) comprises a first engine (5a) with a heat exchanger (16a) through which an air duct (15a) passes, a second engine (5b), which is a heat engine, and an air interconnection duct (21) connecting the air duct (15a), downstream of the heat exchanger (16a) of the first engine (5a), to the second engine (5b). The thermal management method comprises the steps of heating, in the heat exchanger (16a) of the first engine (5a), an air flow circulating in the air duct (15a) passing through the heat exchanger (16a) of the first engine (5a), and diverting the heated air flow downstream of the heat exchanger (16a) of the first engine (5a), through the air interconnection duct (21), to the second engine (5b).
B64D 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
09 - Scientific and electric apparatus and instruments
35 - Advertising and business services
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
Generators for aircraft, compressors; machine coupling and
transmission and propulsion components (other than for land
vehicles); systems using propulsive and non-propulsive power
(machines) for aircraft and component parts thereof included
in this class, including turbines, motors, engines,
thrusters, nacelles, thrust reversers; auxiliary power units
for air vehicles (machines); lubrication systems for motors,
engines and turbines of air vehicles (machines); test
benches for motors, engines, turbines and other thrusters
(machines) for aircraft. Electric and electronic apparatus and instruments namely
generators and/or starters for fixed or mobile installations
for aircraft; electric, electronic and magnetic pressure,
speed, displacement, temperature, position and vibration
sensors; electronic on-board or ground systems, apparatus
and equipment for data and parameter acquisition and
processing; electric and electronic maintenance and control
equipment and hardware for generators, starters and
integrated sets for generating propulsive and non-propulsive
power. Sales services for propulsion and non-propulsion power
systems (machines) for aircraft; administrative and
commercial management of parts, replacement equipment for
users of engines, systems, equipment and parts of aircraft;
commercial advice on propulsion systems for airplanes,
turbines; processing services for acquisition of data
recorded during the operation of aircraft engines, systems,
equipment and parts [office work]. Repair, overhaul, servicing and maintenance services for
propulsive and non-propulsive power systems (machines) for
aircraft and their component parts, including turbines,
engines, thrusters, nacelles, thrust reversers; advisory
services for definition and selection of tools for repair,
overhaul, servicing, standardization and maintenance of
aircraft systems, equipment and parts. Technical, scientific and industrial research; engineering;
research and development (engineering work) in the
aeronautical field; analysis of technical data; services
provided by engineers relating to evaluations, estimates and
research in connection with the technologies used in
systems, equipment and parts of aeronautical vehicles;
testing of machines and materials; design and development of
software and computer programming; analysis and expertise
services for aircraft equipment and parts; analysis and
expertise of data recorded during the operation of aircraft
engines, systems, equipment and parts; technical project
studies in connection with aeronautical vehicles and their
components, including motors and engines, aircraft engine
pods, reactors, thruster units or reversers, aeronautical
vehicles.
52.
LUBRICATION/COOLING SYSTEM FOR AN AIRCRAFT, AND HYDRAULIC ENCLOSURE
Disclosed is a lubrication/cooling system (100) for an aircraft, comprising: a first circuit (20), a second circuit (30), a third circuit (40) for circulating a lubricating and/or cooling fluid (13), a first pump (50) and a second pump (60) for circulating the lubricating and/or cooling fluid (13), a first tank (70) and a second tank (80), and a first heat exchanger (90), wherein the first tank (70) and the second tank (80) are fluidically connected. Also disclosed are a hydraulic enclosure for such a system (100) as well as a turbine engine and an aircraft comprising such a system and/or such an enclosure.
Disclosed is a hybrid turbomachine (10), in particular for a rotary-wing aircraft, comprising a gas generator (13) having a first shaft (16), at least one electric machine (11) having a second shaft (17), and a rotating equipment (15) coupled to a third mechanical shaft (18). The hybrid turbomachine (10) further comprises a switching coupling means (20) configured to couple the third mechanical shaft (18) to the first mechanical shaft (16) or the second mechanical shaft (17), depending on the operating phases of the gas generator (13) and the electric machine (11).
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
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
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
A rotor of an aircraft electric motor includes a shaft made of a first material, and a conductive assembly made of a second material different from the first material. The shaft includes a shoulder portion, the shoulder portion includes longitudinal notches. The notches include two contiguous notches radially superimposed in the shoulder portion, a first opening on a radially outer face of the shoulder portion, and a second opening connecting the two contiguous notches. The conductive assembly is a one-piece structure including a conductive bar that is positioned in one notch of the notches, and a skin that is fixed on the shoulder portion.
H02K 17/16 - Asynchronous induction motors having rotors with internally short-circuited windings, e.g. cage rotors
B64D 27/30 - Aircraft characterised by electric power plants
H02K 15/00 - Processes or apparatus specially adapted for manufacturing, assembling, maintaining or repairing of dynamo-electric machines
55.
AIRCRAFT PROPULSION ASSEMBLY COMPRISING A PROPULSION POWER TRANSMISSION SYSTEM WITH THE CAPACITY TO DECOUPLE TWO ROTATING ELEMENTS ACTUATED FROM THE INSIDE OF ONE OF THE ROTATING ELEMENTS
A system (2) for the transmission of propulsion power of a propulsion assembly (1) for an aircraft comprises means (12) for coupling a first rotating element (4) with a second rotating element (8), and an actuator (10) for moving the second rotating element (8) axially between a coupling position in which the coupling means (12) are engaged with one another, and a disengaging position in which the coupling means (12) are disengaged from one another. The actuator (10) comprises an actuating member (40) extending in a bore (8A) of the second rotating element (8) and cooperating with a drive device (42). A rolling bearing (44) is interposed between the actuating member (40) and an inner surface (8B) of the second rotating element (8) so as to secure the actuating member (40) and the second rotating element (8) to one another axially while allowing the second rotating element (8) to rotate freely relative to the actuating member (40). The rotor is kinematically connected to the rotor of an electric machine configured to receive propulsion power from the first rotating element (4).
F16D 11/10 - Clutches in which the members have interengaging parts actuated by moving a non-rotating part axially with clutching members movable only axially
56.
TURBINE ENGINE ELEMENT COMPRISING AT LEAST ONE BLADE OBTAINED BY ADDITIVE MANUFACTURING
The present invention relates to a turbomachine element (1), comprising at least one blade (2) obtained by additive manufacturing, the blade (2) having a skin (4) and an internal lattice (6) allowing air circulation in the blade (2) and having an additive manufacturing support function for the skin (4).
A method for training a pilot to cope with a fault affecting one powertrain of a hybrid propulsion system for an aircraft. The aircraft includes, connected in parallel to a transmission unit, n powertrains (where n≥2), including a first and a second powertrain that are heterogeneous in nature. It involves, during a flight of the aircraft, simulating a fault affecting the first powertrain while, at the same time as performing the simulation, checking the status of the n powertrains of the propulsion system. If a fault affecting one of the n powertrains is detected, the simulation is halted and the instantaneous power delivered by at least one of either the first or the second powertrain is increased so that the sum of the instantaneous powers delivered by the n powertrains is ≥ a minimum total instantaneous power required for the aircraft to continue its flight.
G09B 9/44 - Simulators for teaching or training purposes for teaching control of vehicles or other craft for teaching control of aircraft, e.g. Link trainer providing simulation in a real aircraft flying through the atmosphere without restriction of its path
G09B 9/46 - Simulators for teaching or training purposes for teaching control of vehicles or other craft for teaching control of aircraft, e.g. Link trainer the aircraft being a helicopter
58.
IMPROVED PROPULSION ASSEMBLY FOR A MULTI-ENGINE HYBRID AIRCRAFT
The invention relates to a propulsion assembly (100) comprising a first and second engine (1, 2) each having a gas generator (12, 22) and a free turbine (11, 21), a main rotor (62) coupled to the free turbine (11, 21), the engines (1, 2) each comprising a first electric machine (30, 40) coupled to the gas generator (12, 22) only, a second electric machine (32, 42) coupled to the gas generator (12, 22) via a first coupling means (34, 44) when said machine rotates in a first direction, and coupled to the free turbine (11, 21) via a second coupling means (36, 46) when said machine rotates in a second direction, the first electric machine (30, 40) operating selectively in a motor or generator mode, the second electric machine (32, 42) operating in the motor mode when it rotates in the first direction, and selectively in the motor mode or the generator mode when it rotates in the second direction.
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 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F02C 6/20 - Adaptations of gas-turbine plants for driving vehicles
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 6/02 - Plural gas-turbine plants having a common power output
A rotor of an aircraft electric motor includes
a shaft made of a first material, and
a skin made of a second material different from the first material. The skin includes two half-shells welded together,
each half-shell of the two half-shells including a chamfer,
and the chamfers assembling the two half-shells together. The shaft includes a shoulder portion the skin being fixed on the shoulder portion. The rotor further includes
an interpenetration layer of the first material and of the second material, the interpenetration layer including an alloy of the first material and an alloy of the second material,
the interpenetration layer being between the shaft and the skin.
The invention relates to a sensor (1) for determining a liquid level (NE) for an aircraft tank (100), the determining sensor (1) comprising a closure device (2) for closing a port (101) of the tank (100) and a measuring device (3), removably mounted on the closure device (2), comprising a liquid line (20) configured to convey liquid from the port (100) of the tank (100), and a member (21) for automatically sealing the liquid line (20) if the measuring device (3) is not mounted on the closure device (2), the measuring device (3) comprising at least one pressure measuring member (30) configured to measure a pressure difference between the liquid pressure (P1) in the liquid line (20) and a reference pressure (P2) in order to deduce the liquid level (NE) thereof.
G01F 23/16 - Indicating, recording, or alarm devices being actuated by mechanical or fluid means, e.g. using gas, mercury, or a diaphragm as transmitting element, or by a column of liquid
B64D 1/16 - Dropping or releasing powdered, liquid or gaseous matter, e.g. for fire-fighting
G01F 22/02 - Methods or apparatus for measuring volume of fluids or fluent solid material, not otherwise provided for involving measurement of pressure
G01F 23/18 - Indicating, recording, or alarm devices actuated electrically
61.
TURBOPROP CAPABLE OF PROVIDING A RAM AIR TURBINE FUNCTION AND AIRCRAFT COMPRISING SUCH A TURBOPROP
A turboprop includes a propeller, a propeller shaft carrying the propeller, the propeller being a variable-pitch propeller having a propeller pitch, a rotating electric machine having at least a first configuration in which it is mechanically coupled to the propeller shaft and at least one oil pump configured to supply a hydraulic circuit for adjusting the pitch of the propeller. The oil pump is configured to be electrically operated. An aircraft can include such a turboprop and methods can control such a turboprop and such an aircraft.
B64D 35/021 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants for electric power plants
B64D 27/30 - Aircraft characterised by electric power plants
B64D 31/06 - Initiating means actuated automatically
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
SAFRAN HELICOPTER ENGINES (France)
UNIVERSITE TOULOUSE III - PAUL SABATIER (France)
Inventor
Richard, Stéphane Raphaël Yves
Viguier, Christophe Nicolas Henri
Marragou, Sylvain
Schuller, Thierry
Abstract
A longitudinal-axis (X) dihydrogen injection device is configured to be mounted on an annular bottom of an annular combustion chamber of a turbomachine. The injection device includes an inner channel for dihydrogen circulation and an outer annular channel for circulation of a mixture of at least air. The inner channel and the outer annular channel are coaxial. An inner swirler is housed in the inner channel and an outer swirler is housed in the outer annular channel. A downstream end of the inner channel is arranged upstream, at a distance r, from a downstream end of the outer annular channel. With this dihydrogen combustion, polluting carbon emissions such as carbon monoxide, unburned hydrocarbons or even fine and smoke particles can be eliminated.
F23D 14/24 - Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other at least one of the fluids being submitted to a swirling motion
F23R 3/28 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
The invention relates to a metal powder for an additive manufacturing process, the metal powder comprising an alloy comprising by weight between 20% and 24% chromium, between 20% and 24% nickel, between 13.00% and 16.00% tungsten, between 0.02% and 0.12% lanthanum, between 0.05% and 0.15% carbon, between 0.20% and 0.50% silicon, at most 1.25% manganese, at most 3.00% iron, at most 0.015% sulfur, at most 0.020% phosphorus, at most 0.0001% bismuth, at most 0.0010% silver, at most 0.0010% lead, at most 0.015% boron, at most 0.0250% oxygen, at most 0.0200% nitrogen and less than 0.050% other elements in total, the balance being cobalt.
C22C 1/04 - Making non-ferrous alloys by powder metallurgy
B22F 1/052 - Metallic powder characterised by the size or surface area of the particles characterised by a mixture of particles of different sizes or by the particle size distribution
B22F 9/08 - Making metallic powder or suspensions thereofApparatus or devices specially adapted therefor using physical processes starting from liquid material by casting, e.g. through sieves or in water, by atomising or spraying
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B33Y 70/00 - Materials specially adapted for additive manufacturing
B22F 10/366 - Scanning parameters, e.g. hatch distance or scanning strategy
B22F 10/64 - Treatment of workpieces or articles after build-up by thermal means
The invention relates to a metal powder for an additive manufacturing method, the metal powder comprising an alloy comprising, by weight, between 23% and 24.5% of chromium, between 9% and 11% of nickel, between 6.5% and 7.5% of tungsten, between 3% and 4% of tantalum, between 0.55% and 0.65% of carbon, between 0.3% and 0.5% of zirconium, between 0.15% and 0.25% of titanium, at most 2% of iron, at most 0.3% of silicon, at most 0.1% of manganese, at most 0.1% of copper, at most 0.015% of sulfur, at most 0.015% of phosphorus, at most 0.01% of boron, at most 0.025% of oxygen, at most 0.020% of nitrogen and at most 0.010% of hydrogen and less than 0.050% of other elements in total, the remainder being cobalt.
C22C 1/04 - Making non-ferrous alloys by powder metallurgy
B22F 1/05 - Metallic powder characterised by the size or surface area of the particles
B22F 9/08 - Making metallic powder or suspensions thereofApparatus or devices specially adapted therefor using physical processes starting from liquid material by casting, e.g. through sieves or in water, by atomising or spraying
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
B33Y 70/00 - Materials specially adapted for additive manufacturing
C22C 19/07 - Alloys based on nickel or cobalt based on cobalt
B22F 10/28 - Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
B22F 10/36 - Process control of energy beam parameters
B22F 10/64 - Treatment of workpieces or articles after build-up by thermal means
B33Y 50/02 - Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
B33Y 80/00 - Products made by additive manufacturing
C22F 1/10 - Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of nickel or cobalt or alloys based thereon
B22F 10/366 - Scanning parameters, e.g. hatch distance or scanning strategy
B22F 5/00 - Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
65.
Fire safety system for a turbomachine comprising means for maintaining a cooling air speed and corresponding turbomachine
An assembly for a turbomachine includes at least one turbine having a turbine disc with an internal bore and an annular cavity which is arranged upstream of the disc. The assembly further includes a fire safety system with a cooling device that supplies the cavity with cooling air via injection means. The fire safety system includes means that divide the annular cavity into first and second cavities. A cooling air speed is maintained at the outlet of the injection means and the cooling air in the first cavity is guided to the internal bore of the turbine disc. A diffuser co-operates with the injection means and an annular cover co-operates with the diffuser and covers first attachment members arranged in the cavity. A radially outer surface of the cover at least partially guides the cooling air at the outlet of the diffuser.
A fuel supply circuit of an aircraft engine includes a centrifugal pump mechanically coupled with an engine shaft delivering mechanical power. The circuit further includes at least one electromagnetic pump including at least one stator delimiting an annular internal volume in which is present a rotor able to drive a fluid, a plurality of magnets annularly distributed on the rotor and at least a plurality of coils annularly distributed inside the stator face-to-face with the magnets. The rotor is connected to the engine shaft by a one-way clutching element.
The invention relates to a method for starting an aeronautical turbine engine having a free turbine and a single-spool gas generator, wherein, in order to ensure, under the control of a computer for controlling the turbine engine, the turbine engine is started only from a battery delivering a nominal DC voltage of 28 V, two electric machines coupled to the same accessory drive mechanically linked to the gas generator of the turbine engine and connected in parallel to the battery are actuated sequentially, the first electric machine being started with a starting torque enabling an increase in the speed of the gas generator with a given minimum acceleration and the second electric machine being started only after the ignition of the combustion chamber of the gas generator has been detected and only before a critical speed corresponding to the maximum resisting torque of the gas generator has been reached, the sum of the starting torques (40) produced by the two electric machines being sufficient to ensure, at all times, that the torque margin M at the maximum drag point B corresponding to the maximum resisting torque (42) of the generator is positive.
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 method for starting an aeronautical turbine engine of a twin-engine aircraft, the aeronautical turbine engine having a free turbine and a single-spool gas generator and comprising two independent electrical networks each being provided with a 28 V battery (BAT1, BAT2) selectively powering a starter-generator (S/G 1, S/G 2), wherein, in order to ensure, under the control of a computer for controlling the turbine engine (EECU), the turbine engine is started first with a voltage of 28 V by connecting the two batteries in parallel and then with a voltage of 56 V by connecting them in series, while the gas generator is prevented from accelerating at an excessive speed, the two batteries are connected in series only when the combustion chamber of the gas generator has been ignited and the speed of the gas generator is higher than a predetermined speed threshold NI, which makes it possible to provide, by the batteries being connected in series, a positive acceleration margin at the maximum drag point of the gas generator.
The invention relates to a device (12) for centring and rotationally guiding a shaft line (14) of a turbomachine; said device (12) comprising: - a first inner ring (16) fixed to the shaft line, a first outer ring (18), a first ball bearing (20) arranged between the first inner ring (16) and the first outer ring (18), a first support (22) supporting the first outer ring, - a second inner ring (24) fixed to the shaft line, a second outer ring (26), a second bearing (28) arranged between the second inner ring (24) and the second outer ring (26), and a flexible cage (30) supporting the second outer ring. The first support (22) has a stiffness greater than the stiffness of the flexible cage (30), and the second bearing (28) is spaced apart from the first ball bearing (20) by a predefined axial spacing (E).
A rotor wheel for an aircraft turbine engine has a disc with a main axis and cells at its outer periphery. The cells extend along the axis, and each has a bottom and two side flanks. Vanes are mounted in the cells of the disc, each vane including a blade connected by a platform to a root mounted in one of the cells. Each root includes, at its radially inner end, a lobe with a first axial end having a circumferential notch and a second axial end having a radially inward facing stop configured to axially bear on a first face of the disc. An annular ring engages the notches of the vanes and is axially clamped against a second face of the disc. Each lobe has a radially inward facing projecting bulb configured to radially bear on the surface of the bottom of the corresponding cell.
A propulsion system includes a gas turbine designed so that a combustion chamber can be ignited in a first ignition range of rotational speeds of a compressor shaft. The system further includes a control device designed to control an electric starter to accelerate the compressor shaft and, when the compressor shaft is accelerated, to control an attempt to ignite the combustion chamber. The gas turbine is designed so that the combustion chamber can be ignited in a second ignition range which is higher than the first ignition range, but not between these two ignition ranges, and the ignition attempt is carried out in the second ignition range.
A disengageable coupling assembly (2) for an aircraft propulsion assembly (1) comprises a first shaft (4) and a second shaft (6); a coupling sleeve (8) movable between positions of coupling and of uncoupling the shafts; fluid chambers defined between the coupling sleeve and the shafts; a tube (10) extending into a bore of the coupling sleeve; and means (12) for selectively supplying a first selection of the chamber(s) (39, 52, 62) and a second selection of the chamber(s) (64) with a fluid (F), via the tube (10). When the coupling sleeve is in the coupling position, a pressure of the fluid in the first selection of chamber(s) applies to the coupling sleeve a first axial force (F1) that maintains the coupling. A pressure of the fluid (F) in the second selection of chamber(s) applies to the coupling sleeve a second axial force (F2) that brings the sleeve into the uncoupling position.
F16D 25/08 - Fluid-actuated clutches with fluid-actuated member not rotating with a clutching member
F01D 3/02 - Machines or engines with axial-thrust balancing effected by working fluid characterised by having one fluid flow in one axial direction and another fluid flow in the opposite direction
B64D 35/00 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
Method for assisting the piloting of a rotary wing aircraft, the aircraft comprising at least two engines, a first engine capable of being placed on standby to ensure an operation of a second engine in the fuel-economy mode, called ECO mode, in which method, in order to reach a determined time between overhauls by limiting the wear of the second engine in the ECO mode, an engine control temperature (TC_PME) associated with the second engine is calculated in a flight computer as a function of a predefined maximum power (PME) of the second engine in ECO mode, the engine control temperature being representative of a current state of damage of the second engine and displayed on a flight screen for the attention of a pilot of the aircraft, so as to allow said pilot to keep below this engine control temperature.
The invention relates to a system for assisting with the piloting of a rotary wing aircraft (10), including two engines, a first engine of which is put on standby to ensure an operation in a fuel economy mode referred to as ECO, the second engine remaining active in the ECO mode, in which system, in order to enable the activation of the ECO mode by a pilot, a flight computer (20) is configured to check in real time the fulfilment of the following conditions to authorise the entry into ECO mode: the sum of the powers supplied by the engines is less than a continuous maximum power, the speed of rotation N2 of a free turbine of the second engine is greater than a predetermined speed threshold that is comprised between 80% and 90% of the maximum speed of the engine, the altitude of the aircraft is greater than a minimum value allowing a transient autorotation phase during the reactivation of the engine in standby in the event of failure of the active engine, and there is no critical failure.
B64D 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
B64D 27/14 - Aircraft characterised by the type or position of power plants of gas-turbine type within, or attached to, fuselages
A device for guiding a main air flow (F1) for an aircraft turbine engine, the device including a first air flowing pipe of a main air flow, the first pipe having a main axis, a plurality of ejectors of a secondary air flow located within the first pipe and configured to eject a secondary air flow and force the flow of the main air flow into the first pipe, the ejectors being distributed around the main axis, and a second air flowing pipe located at the outlet of the ejectors and including one end which is connected to one end of the first pipe, wherein the second pipe includes a narrow end.
F04F 5/16 - Jet pumps, i.e. devices in which fluid flow is induced by pressure drop caused by velocity of another fluid flow the inducing fluid being elastic fluid displacing elastic fluids
A device for a hybrid aircraft including a turboshaft engine having a gas generator, a free turbine, and a main rotor. The device includes a first reversible electric machine coupled to a shaft of the free turbine by way of a first free wheel, and to the main rotor. The device further includes a second reversible electric machine coupled to a shaft of the gas generator by way of a second free wheel, and coupled to the main rotor by way of a third free wheel. The second free wheel activates when the second electric machine rotates in a first direction of rotation, and the third free wheel activates when the second electric machine rotates in a second direction of rotation opposite to the first direction of rotation.
B64D 35/022 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants for electric power plants of hybrid-electric type
B64D 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
F01D 15/10 - Adaptations for driving, or combinations with, electric generators
F02C 7/32 - Arrangement, mounting, or driving, of auxiliaries
77.
PROPELLER FOR AN AIRCRAFT PROPULSION ASSEMBLY, PROPULSION ASSEMBLY, AND METHOD FOR THE USE OF SUCH A PROPULSION ASSEMBLY
A propeller for an aircraft propulsion assembly extending longitudinally along an axis X. The propeller comprising a propeller cone, blades, a guide member extending longitudinally along the axis X and rotating as one with the propeller cone, the guide member being mounted outside the propeller cone in such a way as to form between them a guide path, the guide member having an upstream opening configured to convey a flow of air in the guide path and a downstream opening in such a way as to remove the flow of air downstream, the guide member having through-orifices through which extend the blades of the propeller and compressor vanes, which rotate as one with the propeller cone and which are positioned in the guide path in such a way as to generate an accelerated air flow.
The invention relates to a method for the automated control of the turbine engines of a rotary-wing aircraft during a failure on a turbine engine. Following detection of a failure on a first turbine engine of the aircraft, the automated control method comprises: - determining (102) the flight phase, and then, when the aircraft is in a phase other than a takeoff phase, - activating (110) operation of the turbine engines in a misaligned mode via a progressive decrease in the power of the first turbine engine and a progressive increase in the power of at least one second turbine engine, the progressive decrease in power of the first turbine engine and the progressive increase in power of the at least one second turbine engine being controlled in a complementary manner so as to maintain the rotational speed of the rotary wing at the speed of the rotational speed setpoint of the rotary wing, and - activating (120) an indicator of an engine anomaly, and - arming (130) a power limiter.
The invention relates to an electrical supply circuit of a turbine engine comprising a high-voltage DC circuit powered by a high-voltage DC source, connected to at least one DC/AC converter (3), the at least one DC/AC converter being respectively connected to at least one rotary electrical machine (2), the at least one rotary electrical machine being respectively coupled to at least one propeller of the turbine engine so as to rotate the at least one propeller or to generate electricity by the rotation of the at least one propeller, the circuit comprising at least one voltage step-up stage (3') connected between the high-voltage source and the at least one DC/AC converter (3), the at least one voltage step-up stage being capable of raising the voltage supplied to the at least one rotary electrical machine when the rotary electrical machine rotates the at least one propeller.
H02M 3/00 - Conversion of DC power input into DC power output
H02M 1/32 - Means for protecting converters other than by automatic disconnection
H02M 3/335 - Conversion of DC power input into DC power output with intermediate conversion into AC by static converters using discharge tubes with control electrode or semiconductor devices with control electrode to produce the intermediate AC using devices of a triode or a transistor type requiring continuous application of a control signal using semiconductor devices only
H02M 3/158 - Conversion of DC power input into DC power output without intermediate conversion into AC by static converters using discharge tubes with control electrode or semiconductor devices with control electrode using devices of a triode or transistor type requiring continuous application of a control signal using semiconductor devices only with automatic control of output voltage or current, e.g. switching regulators including plural semiconductor devices as final control devices for a single load
B64D 27/02 - Aircraft characterised by the type or position of power plants
80.
METHOD FOR MEASURING THE THERMAL HISTORY OF A PART
OFFICE NATIONAL D'ETUDES ET DE RECHERCHES AEROSPATIALES (France)
Inventor
Brevet, Philippe, Jean, Christian
Roblet, Anais
Lempereur, Christine
Abstract
The invention relates to a method for analysing the thermal history of a part (100) by means of an optical device, the part having a coating that includes a first marker, the optical device comprising a camera (20) and a light source (10), the method comprising: a calibration step, wherein a reference part having the same coating as the part undergoes a predetermined thermal cycle, a step of acquiring images of the part and a step of acquiring images of the reference part, a step of analysing the images of the part and a step of analysing the images of the reference part, a step of determining a transfer function, wherein a transfer function linking the temperature, the processed optical information and the spatial dimensions is determined, an interpretation step, wherein an image of the thermal history is calculated and the image of the thermal history represents the thermal history of the part according to the predetermined angle.
G01K 11/12 - Measuring temperature based on physical or chemical changes not covered by group , , , or using changes in colour, translucency or reflectance
G01K 11/14 - Measuring temperature based on physical or chemical changes not covered by group , , , or using changes in colour, translucency or reflectance of inorganic materials
G01K 11/20 - Measuring temperature based on physical or chemical changes not covered by group , , , or using thermoluminescent materials
G01K 15/00 - Testing or calibrating of thermometers
G01K 1/143 - SupportsFastening devicesArrangements for mounting thermometers in particular locations for measuring surface temperatures
09 - Scientific and electric apparatus and instruments
37 - Construction and mining; installation and repair services
42 - Scientific, technological and industrial services, research and design
Goods & Services
Computer software and programs providing information relating to aircraft engines and to the operation, maintenance and repair of aircraft engines and/or parts and fittings therefor; Content management software; Databases on digital media; Databases on digital media featuring information relating to aircraft engines, the operation thereof, the repair, servicing, upkeep, maintenance and reconditioning of engines, modules and fittings for aircraft engines; All of the aforesaid goods being for use in and/or designed for aeronautics. Providing of information in relation to the following fields: Repair and maintenance of aircraft engines and components thereof. Providing of technical documentation relating to aircraft engines, the operation thereof, repair, servicing, upkeep, maintenance; Providing online information about industrial analysis and research services; Provision of technical information, in the field of aeronautics; Development (setting up) of databases; Design, installation, maintenance, updating of computer software relating to the aeronautical sector; Rental of computer software in the aeronautical sector; Scientific and technological services and research and design relating thereto, relating to the aeronautical sector: namely engineering; Analysis, surveying and processing of data recorded during the operation of aircraft engines and parts and fittings therefor; Providing of online technical documentation; Hosting of online databases featuring technical documentation; Storage of data; All the aforesaid services being used and/or for use in the aeronautical sector.
82.
AIRCRAFT ELECTRIC OR HYBRID PROPULSION ARCHITECTURE
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
Inventor
Darfeuil, Pierre, Alain, Jean-Marie
Barraco, Thomas, Michel, André, Gérard
Lelong, François, Joseph, Paul
Pierfederici, Serge, Lionel
Klonowski, Thomas
Abstract
Disclosed is an aircraft electric or hybrid propulsion architecture comprising at least two propulsion trains, each propulsion train comprising at least one electric motor (Moteur1, Moteur2) powered by at least two power sources (Source1, Source2, Source3) of the propulsion architecture through at least two power-supply paths that each comprise at least one power converter, and one electrical protection delivering a DC voltage to an HVDC bus (HVDC bus1, HVDC bus2), each HVDC bus distributing this DC voltage to at least one electric motor through at least one power converter and one electrical protection, the propulsion architecture comprising power-supply paths that are at least partially dissimilar and preferably completely dissimilar.
The invention relates to an installation for transferring power (P) between a high-pressure body and a low-pressure body of an aircraft turbine engine, comprising: - an electrical network (PDS) designed to have a DC voltage; - a first electromechanical system (104) connected to the electrical network (PDS) and coupled to the high-pressure body; and - a second electromechanical system (106) connected to the electrical network (PDS) and coupled to the low-pressure body. The installation also comprises a control system (108) designed to control at least one of the first and second electromechanical systems (104, 106) so as to control the transferred power and to control at least the other of the first and second electromechanical systems (104, 106) so as to regulate the DC voltage.
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 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
84.
METHOD AND SYSTEM FOR CONTROLLING THE PERMEABILITY OF A LUBRICATION CIRCUIT OF AN AIRCRAFT TURBOMACHINE
The invention relates to a method for controlling the permeability of a lubrication circuit for an aircraft turbomachine, the turbomachine comprising at least one guide bearing, the lubrication circuit being configured to circulate an oil flow from upstream to downstream between an oil inlet and an oil outlet for lubricating the guide bearing, the method comprising the steps of: measuring a first temperature (T1) of the oil upstream of the guide bearing; measuring a second temperature (T2) of the oil downstream of the guide bearing; calculating a temperature difference between the second temperature (T2) and the first temperature (T1); comparing the temperature difference (ΔT) with a predetermined expected temperature difference (ΔT0); and, when the temperature difference (ΔT) is greater than the expected temperature difference (ΔT0), signalling a permeability fault in the lubrication circuit.
F16N 29/04 - Special means in lubricating arrangements or systems providing for the indication or detection of undesired conditionsUse of devices responsive to conditions in lubricating arrangements or systems enabling a warning to be givenSpecial means in lubricating arrangements or systems providing for the indication or detection of undesired conditionsUse of devices responsive to conditions in lubricating arrangements or systems enabling moving parts to be stopped
G01K 3/14 - Thermometers giving results other than momentary value of temperature giving differences of valuesThermometers giving results other than momentary value of temperature giving differentiated values in respect of space
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
85.
TURBINE ENGINE COMPRISING AN IMMOBILISING MAGNETIC COUPLING DEVICE
The invention relates to a turbine engine (10) for an aircraft, comprising: - a rotating element (14) which is rotatably mounted in a structural element (12) and is intended to generate thrust during the rotation thereof; and - controlled means for immobilising the rotating element (14) with respect to the structural element (12); characterised in that the controlled immobilisation means are formed by a magnetic coupling device (42) that comprises: - a rotor (44) which is coupled with the rotating element (14) and comprises first magnetic elements (46); - a stator (52) which is stationary with respect to the structural element (12) and comprises second magnetic elements (54); the magnetic coupling device (42) being controlled between an inoperative state, in which the rotor (44) is free to rotate with respect to the stator (52), and an operative state, in which the rotor (44) is rotatably immobilised by an immobilisation resisting torque.
The present invention relates to a device (1) for controlling an electric aircraft-propelling assembly, said propelling assembly comprising a propeller (3) and at least one electric motor (4) that is powered by an electric supply voltage and that delivers a torque and a rotation speed to drive the propeller (3). The control device (1) comprises at least a unit (11) for measuring an electric supply voltage, and a control unit (12) suitable for making a signal delivered to the electric motor vary as a function of said electric supply voltage, with a view to making the rotation speed of the propeller vary.
A liquid fuel supply system for an aircraft engine includes a fuel tank, a suction duct connected to the fuel tank and located higher than the fuel tank, an electric pump, a supply pump that is mechanically driven by an accessory gear box and an outlet of the supply pump being connected to a fuel supply circuit of the engine, and an air expulsion drain. The electric pump is in communication with the suction duct independently of the supply pump, the electric pump is in communication with the air expulsion drain, and the supply pump is in communication with the suction duct independently of the electric pump.
The invention relates to a hybrid turboprop engine (100) for an aircraft, comprising: a gas generator (12) which is supported by a generator shaft (14); a free turbine (11) which is supported by a turbine shaft (13) and rotated by a gas flow generated by the gas generator (12), the turbine shaft (13) meshing with a main rotor (60) via a transmission system (50) comprising a first overrunning clutch (51) which is oriented such that the main rotor (60) cannot drive the free turbine (11); and a reversible electric machine (30) meshing with the main rotor via the transmission system (50) in order to drive the main rotor (60) during electric or hybrid operation, the turboprop engine comprising a single oil pump (40) meshing with the transmission system (50) in order to be driven selectively by the turbine shaft (13) or by the electric machine (30) according to the operating mode.
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
F02K 5/00 - Plants including an engine, other than a gas turbine, driving a compressor or a ducted fan
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
89.
Free-turbine turbomachine comprising equipment driven by the free turbine
A turbomachine including a gas generator endowed with a first shaft, a gear box, at least one reversible electric machine coupled to the gear box, and a free turbine endowed with a second shaft and rotationally driven by a gas stream of the gas generator and at least one accessory coupled to the accessory gear box, wherein the turbomachine includes a first mechanical coupling means configured to mechanically couple said first mechanical shaft to the accessory gear box in a first configuration and to mechanically uncouple said first mechanical shaft from the accessory gear box in a second configuration, and a second mechanical coupling means configured to mechanically couple said second mechanical shaft to the accessory gear box in a first configuration and to mechanically uncouple said second mechanical shaft from the accessory gear box in a second configuration.
The invention relates to a method capable of providing information to an operator, for example a pilot, or to an automated system, for example an autopilot, to assist in managing, flight after flight, at least one cumulative ageing process, for example creep, affecting at least part of a gas turbine of an aircraft of versatile use, for example a helicopter, in order to make the best use of such a turbine over the long term.
The invention relates to an electromagnetic signal-based data collection system, comprising, on the one hand, a collector (100) provided with a first antenna (105) having a first bandwidth and, on the other hand, at least one remote electronic device (200) comprising a second antenna (205) tuned so as to exchange signals with the first antenna, characterized in that the device comprises a thermally protective block (209) in which the second antenna (205) is embedded, and in that the second antenna (205) is tuned such that the second antenna (205) and the thermally protective block (209) form a radiating assembly having a second bandwidth that coincides with the first bandwidth. The invention also relates to a vehicle (A) comprising at least one propulsion engine (M) arranged in an engine compartment (G), and at least one remote device (200) of such a system.
The invention relates to an electromagnetic signal-based data collection system, comprising, on the one hand, a collector (100) provided with a first antenna (105) having a predetermined bandwidth and, on the other hand, at least one remote electronic device (200) comprising a second antenna (205) having a second bandwidth that coincides with the first bandwidth for exchanging signals with the first antenna. The device comprises a thermally protective casing (207; 210) surrounding at least the second antenna (205) while defining a cavity around it, such that the second antenna radiates signals outside the protective casing substantially in the second bandwidth. The invention also relates to an aircraft equipped with such a system.
A system for pumping and metering a fluid for a turbine engine, which system includes at least one pump for the fluid and an electronic computer configured to determine the flow rate of the fluid to be delivered to the turbine engine, the pumping and metering system being wherein it further includes a first electric motor and a second electric motor, which are each configured to drive the at least one pump, and in that the electronic computer includes a first control loop for controlling at least the first electric motor and a second control loop for controlling at least the second electric motor.
A fixed-wing combat aircraft comprising an electrical power source, a propulsion system, a low-power non-propulsion assembly comprising a flight control system, a high-power non-propulsion assembly comprising an electrical weapon system, and a management unit configured to selectively establish on command multiple operating modes comprising: a flight mode, in which the management unit distributes the electrical power supplied by the electrical power source to the propulsion system and to the low-power non-propulsion assembly, and an attack mode, in which the management unit limits the electrical power supplied by the electrical power source to the propulsion system and to the low-power non-propulsion assembly to the power required to allow the aircraft to glide, and reserves a majority of the available electrical power for the high-power non-propulsion assembly.
A turbine for a turbomachine of longitudinal axis including an alternating arrangement of annular rows of movable blades and of fixed blades and a radially inner annular cavity formed radially inside the movable and fixed blades, and a supply circuit for supplying cooling air to the inner annular cavity, the downstream end of the supply circuit comprising an inner annular row of orifices and an outer annular row of orifices opening into the radially inner annular cavity. The turbine may also include means for controlling the flow rate of supply air to the orifices of the inner and outer annular rows of orifices.
A turbomachine including a rotary body including a motor shaft supplying mechanical power, and at least one magnetic drive pump including at least: one stator delimiting an annular inner space and including a first and a second flange, a rotor arranged in the inner space between the first and second flanges and capable of driving fluid, the rotor being able to rotate about an axis of rotation, a pair of magnets having opposite polarities coaxially arranged on the rotor with the axis of rotation, a magnet arranged on the first flange in order to co-operate with one of the magnets of the pair of magnets of the rotor, a magnetic rotator for rotating the rotor arranged on the second flange, the second flange being non-magnetic.
The invention relates to an assembly comprising a turbomachine engine structure (10) extending along a first axis (Z), at least one equipment item (30) and a suspension device (20) for suspending the equipment item (30) on the engine structure (10) while being offset in the direction of at least one second axis (X) perpendicular to the first axis, the suspension device (20) comprising: at least six support links (21; 22; 23; 24; 25; 26) connecting the equipment item (30) to the engine structure (10) in a statically-determinant manner; at least one safety link (27) connecting the equipment item (30) to the engine structure (10), the safety link (27) being configured to have a mechanical loading lower than a mechanical loading of each of the support links (21; 22; 23; 24; 25; 26).
The present invention relates to an aircraft turbomachine (10) with a recuperation cycle, comprising: - a heat exchanger (6) comprising a first circuit (62) with an inlet (622) connected to an outlet (444) of a flow path (44) of a turbine (4), and a second circuit (64) with an inlet (642) connected to an air bleed system (20), and an air outlet (644), - at least one duct (7) for passing services (S) extending from a turbine casing (40) to a bearing housing (5), and - an air circulation device (8) comprising a first channel (82) with a first upstream end (822) connected to the system (20) and a first downstream end (824) connected to the inlet (642) and a second channel (84) having a second downstream end (844) connected to the outlet (644), and wherein the duct (7) extends radially outwards as far as the device (8).
Device (210) for measuring a rotational speed of an aircraft propeller (211), comprising: - an optical speed sensor (220); - an optical fiber (230) connected to the optical speed sensor and intended to be connected to a control unit of the aircraft; and - at least one target (251, 252) configured to follow the rotation of the propeller, the optical speed sensor being configured to detect the target.
G01P 3/481 - Devices characterised by the use of electric or magnetic means for measuring angular speed by measuring frequency of generated current or voltage of pulse signals
B64D 35/00 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions
B64C 11/50 - Phase synchronisation between multiple propellers
B64F 5/60 - Testing or inspecting aircraft components or systems
B64D 31/12 - Initiating means actuated automatically for equalising or synchronising power plants
B64D 45/00 - Aircraft indicators or protectors not otherwise provided for
B64D 27/24 - Aircraft characterised by the type or position of power plants using steam or spring force
Propulsion system for a helicopter comprising a main engine, a main rotor, a main gearbox including an output mechanically connected to the main rotor, a reduction gearbox mechanically coupled between the main engine and a first input of the main gearbox, and an assistance device. The assistance device comprises a first electric machine mechanically coupled to the reduction gearbox and configured to operate as an electric generator to take off energy produced by the main engine, and a second electric machine mechanically coupled to a second input of the main gearbox, the second electric machine being supplied with electrical power by the first electric machine and configured to operate as an electric motor to deliver additional mechanical power to the main gearbox.
B64D 35/023 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions specially adapted for specific power plants for electric power plants of hybrid-electric type of series-parallel type
B64D 31/18 - Power plant control systemsArrangement of power plant control systems in aircraft for electric power plants for hybrid-electric power plants
