Abstract: A structure (10) for aircrafts, includes a part (20) including a metal leading edge (30), the leading edge (30) being covered by a coating (40) having a thickness which is less than or equal to ten micrometers and having a hardness higher than six hundred in the Vickers hardness test (HV). According to a mode of embodiment, the coating (40) is a multilayer stainless steel coating consisting of a superposition of layers with a low nitrogen gradient and layers with a high nitrogen gradient, the layers having a thickness essentially equal to a micrometer. An aircraft including such a structure is also described.

Abstract: The invention relates to a device for monitoring the integrity and soundness of a mechanical structure, such as an aircraft. The device comprises a control unit, a radio frequency transmission means, and an electric battery. The control unit recovers data from a set of digital and/or analog sensors. The radio frequency transmission means enables the control unit to transmit the data received from the sensors to a man/machine interface. The electric battery powers the device and is rechargeable. The device further comprises a module for recovering electromagnetic energy capable of converting the recovered electromagnetic energy into electric power so as to recharge the battery and/or directly power the device.

Abstract: An aircraft (1) having a rotary wing (2) and turboshaft engines (11, 12, 13) for driving said rotary wing (2). The aircraft then includes two main engines (11, 12) that are identical, each capable of operating at at least one specific rating associated with a main power (maxTOP, OEIcont), and a secondary engine (13) capable of operating at at least one specific rating by delivering secondary power (maxTOP?, OEIcont?) proportional to the corresponding main power (maxTOP, OEIcont) in application of a coefficient of proportionality (k) less than or equal to 0.5, said aircraft having a control system (20) for driving the rotary wing by causing each main engine (11, 12) to operate continuously throughout a flight, and by using the secondary engine (13) as a supplement during at least one predetermined specific stage of flight.

Abstract: A reversible electrical machine (1) comprising: a first electrical device (10) having a first stator (11) and a first rotor (12); a second electrical device (20) including a second rotor (22) and a second stator (21) together with an outlet shaft (50) and first disengageable coupling means (30) enabling said first and second rotors (12, 22) to be associated and dissociated in rotation. Said reversible electrical machine (1) also includes second disengageable coupling means (40) that are disengageable under a predetermined force and that mechanically connect said second rotor (22) to said outlet shaft (50). Said first electrical device (10) is a motor for transmitting high mechanical power to said outlet shaft (50), while said second electrical device (20) is a motor-generator for operating in motor mode to transmit additional mechanical power to said outlet shaft (50), and in generator mode for receiving mechanical power from said outlet shaft (50).

Abstract: A method of automatically triggering an emergency buoyancy system (10) for a hybrid helicopter (20) having a fuselage (21), two half-wings (23, 23?), and two propulsive propellers (24, 24?). During the method, said emergency buoyancy system (10) is primed, and then if a risk of said hybrid helicopter (20) ditching is detected, two retractable wing undercarriages (28, 28?) are deployed, each wing undercarriage (28, 28?) being fastened under a respective half-wing (23, 23?) and being provided with at least one immersion sensor (16). Finally, if the beginning of said hybrid helicopter (20) ditching is detected, at least one main inflatable bag (11, 11?) 7B suitable for being arranged under such fuselage (21) and at least one secondary inflatable bag (12, 12?) suitable for being arranged under each half-wing (23, 23?) are inflated so as to ensure that said hybrid helicopter (20) floats in stable manner.

Abstract: A method of driving a rotorcraft rotor (1, 2) at variable controlled speeds. Starting from a required speed (Nr), a setpoint for total power (Pt) is calculated while taking account of an anticipated power (Pf) to be delivered by the power plant (3) for driving the main rotor (1) at the required speed (Nr), and while taking account of a power surplus (S) relating to progressive power needs to be delivered from a current power (Pc) for driving the main rotor (1) at a current speed (V) of rotation to an anticipated power (Pf) for driving the main rotor (1) at the required speed (Nr). By way of example, account may be taken at least of the acceleration or conversely of the deceleration of the main rotor (1) between the current speed (V) and the required speed (Nr), or indeed account may be taken of the flight circumstance of the rotorcraft determining the calculated required speed (Nr).

Abstract: The present invention provides a method of driving a main rotor (2) of a rotorcraft (1) in rotation. A regulation setpoint (C) for a power plant (3) used for driving the main rotor (2) at a variable speed of rotation is generated by a control unit (4) and is transmitted to a regulator unit (5) for regulating the operation of the power plant (3). The value of an initial setpoint (NRini) is generated progressively and continuously depending on variation in the current value of the density (D) of the ambient air outside the rotorcraft (1). The value of the initial setpoint (NRini) is potentially corrected depending on predefined flight conditions of the rotorcraft (1). Prior to being transmitted to the regulator unit (5), the value of the regulation setpoint (C) is preferably limited to a range of acceptable speeds for driving the main rotor (2) in rotation.

Abstract: A method of optimizing the performance of a rotary wing aircraft having at least one turbine engine with a gas generator and a turbine assembly comprising at least one turbine. In a definition step (STP1), first and second performance levels are defined for the aircraft. During a health check step (STP2), a first power margin is determined as a function of a speed of rotation of said gas generator and a second power margin is determined as a function of a temperature in said turbine assembly. During an evaluation step (STP3), each power margin is compared with a first threshold in order to determine whether a target performance level is equal to the first performance level or to the second performance level. During a display step (STP4), the target performance level is displayed.

Abstract: The present invention relates to an electric machine (1) that includes a stator (10) and a rotor (20), with the said stator (10) being equipped with at least one annular exciter unit (11) that includes a coil (12) and at least two annular yokes (13a, 13b), with the said rotor being equipped with a structure (21) and at least one annular receiver unit (22). Each receiver unit (22) includes at least two rows (24) of magnets (23). Two sides (131, 132) of each yoke (13) include teeth (14) distributed angularly in a regular manner, and the said teeth (14) of the two adjacent yokes (13) fit onto a face (121) of the said exciter unit (11), alternately forming north poles and south poles.

Abstract: The present invention relates to an electric machine (1) consisting of a stator (10) that is provided with at least one exciter unit (11) consisting of a coil (12), at least two annular yokes (13a,13b) and at least one row of intermediate pieces (15), and a rotor (20) having a structure (21) and at least one receiver unit (22) consisting of at least two series (25) of at least two rows (24) of magnets (23). Two sides (131,132) of each yoke (13) consist of the first teeth (14), fitting with the said intermediate pieces (15) on a face (121) of the said exciter unit (11) and alternatingly forming the second north poles and the second south poles. Each series (23) is positioned opposite one face (121), forming an air gap (30) with the said exciter unit (11), with the electric machine (1) thus including at least two air gaps (30), with a flux F thus circulating inside the said electric machine (1), dividing and regrouping itself in the vicinity of the said magnets (23) and of the said yokes (13).

Abstract: A buoyancy system (10) for an aircraft (1), said buoyancy system (10) comprises an inflatable float (15) and a cover (20). Said buoyancy system (10) includes at least one extensible connection means (25) fastened to the cover (20) to connect the cover (20) to a compartment (3) of an aircraft (1), while enabling the cover (20) to be retained and to be moved transversely during inflation of the float (15).

Abstract: A tank (10) having a casing (20) with a bottom (21) together with a side wall (23) and a top wall 22. The tank (10) has at least one sloping wall (30) for providing at least one slope (30) for directing said fuel towards a predetermined zone (100) of the tank (10) by gravity, said sloping wall (30) subdividing the inside volume of the tank into a main volume (V1) and an under-slope volume (V2) both of which are to receive fuel. Said tank (10) is provided with a fuel transfer system (40) for transferring fuel from the under-slope volume (V2) to the main volume (V1), and from the main volume (V1) to the under-slope volume (V2).

Abstract: An aircraft (1) having a rotary wing (2) and at least one main gearbox (5) for driving rotation of said rotary wing (2), said aircraft (1) having a first main engine (11) and a second main engine (12) driving said main gearbox (5), said aircraft (1) being provided with a main regulation system (15) regulating the first main engine (11) and the second main engine (12) in application of a setpoint that is variable. A secondary engine (21) is also capable of driving said main gearbox (5), said aircraft (1) having a secondary regulation system (25) that regulates the secondary engine (21) in application of a setpoint that is constant and that is independent of said main regulation system (15).

Abstract: A fuel storage device (10) having at least one tank (15), the tank (15) having a casing (20) extending upwards from a bottom (21) to a top wall (22). The storage device (10) is provided with at least one inflatable bag (25) arranged inside said casing, said device including inflation/deflation means (35) for inflating each inflatable bag (25) at least in part so as to reach an inflated volume prior to filling the tank, and for deflating said inflatable bag after filling in order to guarantee the presence of an expansion volume inside said tank.

Abstract: The present invention relates to a locking device provided with a latch (10) and a retaining pin (100), said latch (10) having a handle (11) and a locking arm (20) hinged to said handle (11) by a pivot peg (13). The retaining pin (100) comprising a first transverse branch (101) and a second transverse branch (102) separated by a space (103), a second end (22) of said locking arm (20) has a contact finger (23) provided with a first surface and with a second surface (25) opposite from said first surface (24), said first surface (24) comprising closure means for co-operating with said first transverse branch (101), said second surface (25) comprising extractor means of said latch (10) for co-operating with said second transverse branch.

Abstract: A method of managing an engine failure on a rotary wing aircraft (1) having a hybrid power plant (5) with at least two fuel-burning engines (13, 13?), at least one electric machine (12), and a main gearbox (11). Said aircraft (1) also has electrical energy storage means (14) and a main rotor (2) mechanically connected to said hybrid power plant (5). In said method, during each flight, the operation of said engines (13, 13?) is monitored in order to detect a failure of any one of them, and then once a failure of one of said engines (13, 13?) has been detected, said electric machine (12) is controlled, if necessary, to deliver auxiliary power We in order to avoid a deficit appearing in the total power WT of said hybrid power plant (5), thereby enabling the pilot of said aircraft (1) to fly said aircraft (1) safely without degrading said hybrid power plant (5).

Abstract: A rotorcraft fitted with means for mounting at least one front avionics rack (3, 3?) and man-machine interface instruments (2) on board a fuselage structure (4). A mounting structure (1) comprises a slotted body forming compartments for receiving interface instruments (2), the avionics rack (3, 3?), and a unitary cabling assembly (12) suitable for incorporating as a block with the mounting structure (1). The mounting structure (1) is installed as a block on the fuselage structure (4) after its functioning has been verified. The unitary cabling assembly (12) may also include separation connectors (14) segregating communications of the front avionics rack (3, 3?) respectively with remote computers (5) by means of a primary communications bus of the multiplexed, unidirectional, or bidirectional type, and with ancillary equipment (6) and/or with on-board instruments (7) by means of a secondary fieldbus.

Abstract: A noise treatment device (D) comprising at least one local noise sound sensor (30) and at least one sound system having a support (11, 12) and at least one sound actuator (21, 22, 21?, 22?). The device also includes a position sensor for determining the position of a person's head, at least one treatment unit (40, 40?) connected to the local noise sound sensor to receive a local noise signal and configured to deliver a control signal to each sound actuator (21, 22, 21?, 22?), the control signal being a function of said local noise signal and of at least one transfer function per ear, and active matching means (70) co-operating with said position sensor in order to keep each transfer function used in preparing each control signal representative of the path to be traveled by said anti-noise.

Abstract: A method of controlling a high-speed rotary wing aircraft (1) comprising a fuselage (2), at least one main rotor (3), at least one variable-pitch propulsive propeller (4), at least two half-wings (11, 11?) positioned on either side of said fuselage (2), at least one horizontal stabilizer (20) provided with a movable surface (21, 21?), and at least one power plant driving said main rotor (3) and said propulsive propeller (4) in rotation. Said method serves to adjust the lift of said half-wings (11, 11?) and the lift of the horizontal stabilizer (20) so that said lift of said half-wings (11, 11?) represents a predetermined percentage of the total lift of said aircraft (1) and so that the power consumed by said main rotor (3) is equal to a setpoint power during a stage of stabilized flight.

Abstract: A chained information exchange system (10) comprising a chain of modules (1, 2, 3, 4, 5), each module (1, 2, 3, 4, 5) being connected to one or two other modules (1, 2, 3, 4, 5) via digital buses (11, 12, 13, 14, 15), thereby forming a chain that is open or else a continuous loop that is closed. Each digital bus (11, 12, 13, 14, 15) is a hardened digital bus, capable of withstanding external electromagnetic disturbances, and it is unidirectional. A signal travels in said information exchange system (10) and consequently through each module (1, 2, 3, 4, 5), and after passing through a module (1, 2, 3, 4, 5), said signal contains information that the module (1, 2, 3, 4, 5) through which it has passed does not modify and that is addressed to at least one other module (1, 2, 3, 4, 5), together with specific information that has been added by said module (1, 2, 3, 4, 5) through which it has passed and that is addressed to at least one other module (1, 2, 3, 4, 5).