Abstract: A mobile power plant having an engine together with a main gear box (5) and a free wheel assembly (4) suitable for driving a rotary wing (7) by means of the main gear box (5) and further having an electric motor-generator (10), wherein the power plant includes an accessory gear box (12), an additional freewheel assembly (11) and a pump (8), said electric motor-generator (10) being mechanically connected in series via said additional freewheel assembly (12) to said main gear box (5) in order to be capable of being driven by said main gear box (5) and said electric motor-generator (10) being connected to said pump (8) in order to be capable of driving said pump (8). The invention relates as well to an application of such a mobile power plant and to a method of operating the mobile power plant.

Abstract: A heat control system for an adaptive consolidation roller (8) for the production of consolidated layers (2) made of fiber reinforced polymers comprising the adaptive consolidation roller (8) for applying, on an application surface (3), a band (4) formed of at least one resin pre-impregnated flat tape (5), and a heating system (6, 7) adapted to emit heat radiation directed towards the band (4). Said adaptive consolidation roller (8) is rotatable on a spindle (9) and comprises a surface made of an elastic deformable, flexible material. At least one additional roller (10) with thermal control means is provided and said at least one additional roller (10) is pressed from radial outside against the surface of the adaptive consolidation roller (8).

Abstract: An electric circuit with one power amplifier (V1) for at least two piezoelectric actuators (C1, C2), said piezoelectric actuators (C1, C2) being in preferably antagonist arrangement connected mechanically in series and being connected electrically with each other and operated or clocked by a control circuit, such as a pulse modulator circuit, for time-variant energizing of the piezoelectric actuators (C1, C2) in push-pull fashion, said power amplifier (V1) being connected to a junction point between the two piezoelectric actuators (C1, C2), characterized in that said power amplifier (V1) is connected to a one directional diode (D1) and in that two buffer capacitors (C3, C4) and two static resistors (R3, R4) are provided downstream the one directional diode (D1) with regard to the power amplifier (V1).

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 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 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 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: The invention relates to a separable blade attachment (1) for a bearingless main rotor of a helicopter with an airfoil blade (2), a flexbeam (3) and a control cuff (4) enclosing at least a predominant portion of said flexbeam (3). A separable junction arrangement between said flexbeam head (13), said control cuff (4, 22) and said root end of said airfoil blade (2) is mechanical with removable fasteners. A stiffened link (17) is provided between respective trailing edges (9) of the control cuff (4) and said airfoil blade (2).

Abstract: Aircraft landing gear (5) having a first skid (6) and a second skid (7), together with two cross-members, each having a first branch (16) secured to the first skid (6) and a second branch (17) secured to the second skid (7), and a central portion (13) secured to the downward first branch (16) and to the second branch (17). The landing gear (5) has at least one stiffener (20) with at least one connecting rod (30) and at least one limiter means (40) for limiting roll deformation of the central portion (13) of a cross-member, each limiter means (40) being secured to said central portion (13) of the cross-member, at least one main hinge (33) hinging each connecting rod to said limiter means and a secondary hinge (34) connecting each connecting rod to an outside point that is outside the central portion (13) in order to limit the deformation of the central portion (13) as a result of a roll movement of an aircraft (1).

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).

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 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 each 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 the collective pitch of said blades (31) of said main rotor (3) is equal to a setpoint collective pitch, so that the longitudinal cyclic pitch of said blades (31) of said main rotor (3) is equal to a setpoint longitudinal cyclic pitch, and so that the lateral cyclic pitch of said blades (31) of said main rotor (3) is equal to a setpoint lateral cyclic pitch during a stabilized stage of flight.

Abstract: A pitch stabilizer [10] equipped with an upper stabilizer surface [15] and a lower stabilizer surface [20], with the upper stabilizer surface [15] being located in an upper plane [P1] and with the lower stabilizer surface [20] being located in a lower plane [P2], with the upper plane [P1] overhanging the lower plane [P2]. The upper stabilizer surface [15] is rotatable about an upper axis of rotation [AX1], with the said lower stabilizer surface [20] being rotatable about a lower axis of rotation [AX2], with the stabilizer means [10] having linking means [30] whereby a rotary movement of one stabilizer surface causes an equal rotary movement of the other stabilizer surface.

Abstract: A method of assisting a pilot in order to minimize the risks of an aircraft rolling over on the ground, the aircraft having a main rotor and a yaw movement control rotor (4), together with lateral cyclic pitch control means for the main rotor and yaw movement control means for the yaw movement control rotor. At least one measurement is taken relating to left forces exerted on a left side undercarriage and to right forces exerted on a right side undercarriage in order to evaluate asymmetry, if any, between the left and right forces, and then a recommended position (26) is determined for at least one of the yaw and lateral cyclic pitch control means in order to make the left and right forces more symmetrical. Each recommended position is displayed on a display unit.

Abstract: An empennage (1.1) of a helicopter comprising a ducted counter-torque device with a multi-blade rotor (4) of rotor blades (3) and optional vertical tail fins (1.2). Flow-straightening stators (5) of stationary vanes are disposed substantially in a star configuration parallel to the rotor plane downstream from the rotor (4). A shroud (2.1) of the ducted counter-torque device is sheathed with a composite structure of an outer erosion protecting surface layer (7.1, 8.1) made of a hard plastic or a plastic composite material, and at least one succeeding layer (7.2, 8.2) of an elastomeric damping material.

Abstract: The present invention is related to a compound helicopter (1) with a main rotor (13) providing lift, a pair of additional propulsion devices (14, 15) providing thrust and anti-torque and a fixed wing structure (16, 17, 26) on each side providing additional lift during horizontal cruise flight. The propulsion devices (14, 15) are arranged at each side of a fuselage body (2) with two tail booms (7, 8), one arranged at each side of said fuselage body (2), each of both tail booms (7, 8) having one of both propulsion devices (14, 15) arranged at its boom rear end (9). Each of the two tail booms (7, 8) houses an engine (21) and a drive shaft (24) driving the corresponding propulsion device (14, 15).

Abstract: The invention relates to a compound helicopter (1) comprising a fuselage (2), at least one engine (14) and a main rotor (11) driven by said at least one engine (14). At least one pair of fixed wings are mounted in an essentially horizontal plane on a left hand and a right hand side of the fuselage (2) and horizontally oriented propulsion devices (12, 13) are mounted to each of said fixed wings, said fixed wings encompassing each a drive shaft (16) from said at least one engine (14). Each fixed wing comprises a lower main wing (18) and an upper secondary wing (19) being connected to each other within an interconnection region (22). The propulsion devices (12, 13) are arranged at said interconnection region (22) and said upper secondary wing (19) houses the drive shaft (16) from said at least one engine (14).

Abstract: A system comprising a rotor blade (1), in particular of the tail rotor of a rotary wing aircraft, in a fiber-reinforced composite design, with a blade section (5) and with a coupling section (9) for attaching the rotor blade (1) to the hub of a drive device, and comprising a separate sleeve-shaped control tube (40) with an essentially hollow-cylindrical shaft (41), with an also essentially tubular tie section (56) to tie the control tube (40) to the rotor blade (1) by sliding it onto its coupling section (9), is improved in that the coupling section (9) of the rotor blade (1) and the tie section (56) of the control tube (40) comprise a cross-sectional shape for positive-locking interconnection of the rotor blade (1) and the control tube (40).

Abstract: A method of performing a health check of at least one turbine engine (3). During a development step (STP0), the installation losses (1) are quantified for a plurality of test values for a reduced speed of rotation (Ng?) of a gas generator (4) of the engine. During an acquisition step (STP1), the speed of rotation of said gas generator (4) is increased until said engine develops a maximum power, and then the speed of rotation of the gas generator (4) is decreased until the reduced speed of rotation (Ng?) reaches a test value. The aircraft is stabilized and at least one monitoring value is acquired. During an evaluation step (STP2) of evaluating the health check, at least one operating margin is determined by using a monitoring value and the effects of mounting the engine in an airplane.

Abstract: A method of assisting a pilot of a single-engined rotary wing aircraft (1) during a stage of flight in autorotation, said aircraft (1) including a hybrid power plant having an engine (13), an electric machine (12), and a main gearbox (11). Said aircraft (1) also includes electrical energy storage means (14) and a main rotor (2) mechanically connected to said hybrid power plant (5). According to said method, while in flight, the operation of said engine (13) is monitored in order to detect a failure thereof, in particular by monitoring for a drop of power on said main rotor (2), and then in the event of a failure of said engine (13) being detected, said electric machine (12) is controlled to deliver auxiliary power We to said main rotor (2), thereby enabling the pilot of said aircraft (1) to be assisted in maneuvering said aircraft (1) during said stage of flight in autorotation following said failure.