Abstract: A method of controlling and regulating a rotorcraft presenting a speed of advance that is high and stabilized, the rotorcraft including at least a main lift rotor (10), at least one variable pitch propulsion propeller (6), and at least one power plant for driving the main rotor(s) (10) and at least one propeller (6), said method consisting in using a first loop for regulating pitch or attitude, and a second loop for regulating speed by means of a control over the mean pitch of the propulsion propeller(s) (6), wherein the method further consists in controlling the deflection angle of a horizontal tailplane (30, 25, 35) by using a third loop for controlling and regulating said deflection angle of the horizontal tailplane (30, 25, 35) in order to minimize the total power consumed by the main rotor (10) and the propulsive propeller(s) (6), for a given speed and attitude.

Abstract: A transport wheel device for a helicopter having a runner defining an axial direction. The transport wheel includes a holding frame, a wheel axle, a slide, a wheel and a lifting device. The holding frame is detachably securable to the runner. The wheel axle is attached the holding frame by the slide. The slide is continuously adjustable axially with respect to the holding frame and securable in an axial position. A wheel is disposed on the wheel axle. The lifting device is coupled to the holding frame and the wheel axle. The lifting device is configured to vertically move the transport wheel.

Abstract: The present invention relates to a foldable step unit (10) for a rotorcraft, the step unit being provided with a bottom step (16) and a stationary support (11) that is secured to the structure of the rotorcraft (2). The step unit (10) comprises a left side beam (14) and a right beam (15) hinged on said stationary support (11), said bottom step (16) being arranged on a left free end (14?) and a right free end (15?) respectively of the left and right side beams (14, 15) via a pivot pin (16?), drive means (18) for said step unit (10) and connected via at least one control means (21, 22) to at least one side beam (14, 15) enable said step unit (10) to be retracted into and extended from a housing (2) formed in the rotorcraft (1).

Abstract: An assisted control system for controlling the attitude of a rotorcraft comprises a flight control device; at least one control member configured to act on an aerodynamic element of the rotorcraft; a mechanical connection connecting the flight control device to the at least one control member, the mechanical connection including at least one connecting rod moveable in translation, a crank device pivotably disposed about a stationary support shaft passing through the crank device, and a motor including a drive rotor and a stator configured to facilitate a pivoting movement of the crank device about the stationary support; a motor control device configured to control a rotation of the drive rotor relative to the stator; and at least one sensor configured to measure information representative of an operation performed by the flight control device under pilot control and to deliver a signal to the motor control device.

Abstract: The present invention relates to a bellcrank (40) for a flight-attitude-changing linkage (27, 25A, 28A) of the variable gain type. The linkage (27, 25A, 28A) connects a manual flight control device of a rotary wing aircraft (1), e.g. a hybrid helicopter, to at least one airfoil of said aircraft (1) suitable for causing a change in flight attitude. Said bellcrank (40) includes first and second sliders (50, 52). Said first slider (50) is connected to a movable portion (57, 61) of a setpoint flight-attitude-changing linkage (24) that delivers a setpoint, with movements of said linkage causing the gain of said variable-gain linkage (27, 25A, 28A) to be adjusted, which gain is proportional to the position of said movable portion (57, 61).

Abstract: A fairing (1) for a structural element of a rotorcraft, the fairing including a rear portion (2) that is substantially orthogonal to the longitudinal direction (L) of the aircraft, extending between two spaced-apart trailing edges (3, 4) and thus presenting a determined width, said rear portion (2) closing at least part of the internal volume defined by the fairing (1) and generating aerodynamic drag in forward flight. According to the invention, the rear portion presents, at least at the trailing edges (3, 4), a shape that is perturbed on the streamlines (5) of the air-flow.

Abstract: Method for fabricating a reinforcing coating for a part made of composite material. Pre-impregnated reinforcing fibers are used in the form of slivers and said slivers are laid on the part in two layers of longitudinal slivers by being placed thereon. This fabrication makes provision for using an angular orientation for the slivers of each layer respectively of 0° and of some other angle relative to said determined direction of the solid body for covering using an interlaced pattern type that gives layer crossing lines parallel to a longitudinal axis of the solid body for covering, which pattern type is selected to form interlacing patterns that limit the number of layer crossing lines along the longitudinal axis of the solid body for covering to a single line.

Abstract: An aircraft (1) provided with a tail boom (4) including a tail plane (100) provided with an airfoil surface (5), the airfoil surface (5) being provided with at least one stabilizer portion (11, 12) extending laterally outside the tail boom (4) and with a central portion (13) extending inside the tail boom (4), said stabilizer portion (11, 12) being implanted in an orifice (6, 7) of the tail boom (4) that is defined by a peripheral wall (6?, 7?), clearance (21, 22) existing between said stabilizer portion (11, 12) and said peripheral wall (6, 7). The aircraft includes combating means (30) for combating the flutter phenomenon, which means are provided with stiffener means (31, 32) exerting predetermined compressive stress on said stabilizer portion (11, 12), said stiffener means (31, 32) being arranged at the root of the stabilizer portion (11, 12).

Abstract: A connection means (1) for mounting a rotor blade (2) to a rotor hub, particularly for mounting a rotor blade (2) to a rotor hub of a helicopter, with at least two bolts (4, 5) inserted vertically relative to a main plane of said rotor blade (2) into an end (7) of said rotor blade (2) that is generally directed towards the rotor hub. At least two bolts (4, 5) are arranged distant from each other along a longitudinal axis (6) of said rotor blade (2) to said rotor hub.

Abstract: A buoyancy system (2) comprising at least one float (3) and deployment means (30) for deploying said float (3), said buoyancy system (2) having engagement means (5) for activating the deployment means (30) of said float (3), said buoyancy system (2) including at least two immersion sensors (20) for issuing an order for automatic deployment of said float (3) to the deployment means (30). Said engagement means (5) are activated only manually, said deployment means (30) are provided with a memory (31) containing a pre-established list of events, and said deployment means (30) deploy each float (3) when a predetermined event occurs.

Abstract: A method of optimizing the operation of left and right propellers disposed on either side of the fuselage of a rotorcraft including a main rotor. Left and right aerodynamic surfaces include respective left and right flaps suitable for being deflected, yaw stabilization of the rotorcraft being achieved via first and second pitches respectively of the left and right propellers, and the deflection angles of the left and right flaps are adjusted solely during predetermined stages of flight in order to minimize a differential pitch of the left and right propellers so as to optimize the operation of the left and right propellers, the predetermined stages of flight including stages of flight at low speed performed at an indicated air speed (IAS) of the rotorcraft that is below a predetermined threshold, and stages of yaw-stabilized flight at high speed performed at an indicated air speed of the rotorcraft greater than the predetermined threshold.

Abstract: An automatic configuration-tracking apparatus for tracking the configuration of a vehicle having a multitude of components for which it is desired to consult and record various items of data specific thereto, and their identification/authentication parameters, history, and operating state includes transponders. The transponders are coupled respectively to components of the vehicle. The transponders communicate this specific data via a meshed wireless network in order to enable the data to be transmitted by secure wireless transmission to an external installation. A secrecy perimeter coincides substantially with an outline of the vehicle.

Abstract: A rotary actuator incorporating a force relationship includes a twistably-deformable structure having a first and a second end plate and a helical spring with a first and a second terminal end secured on the first and the second end plate, respectively, and wherein the first and the second end plate have a first and a second notch, respectively, and wherein the first and the second notch have a first and a second contact zone, respectively. The rotary actuator further includes an outlet shaft and a motor assembly having an outlet gear configured to drive the twistably-deformable structure so as to drive the outlet shaft in rotation, wherein the motor assembly, the twistably-deformable structure and the outlet shaft are disposed in succession. The rotary actuator has a first finger constrained to rotate with the outlet gear and a second finger constrained to rotate with the outlet shaft.

Abstract: A hybrid helicopter (1) includes firstly an airframe provided with a fuselage (2) and a lift-producing surface (3), together with stabilizer surfaces (30, 35, 40), and secondly with a drive system constituted by: a mechanical interconnection system (15) between firstly a rotor (10) of radius (R) with collective pitch and cyclic pitch control of the blades (11) of the rotor (10), and secondly at least one propeller (6) with collective pitch control of the blades of the propeller (6); and at least one turbine engine (5) driving the mechanical interconnection system (15). The device is remarkable in that the outlet speeds of rotation of the at least one turbine engine (5), of the at least one propeller (6), of the rotor (10), and of the mechanical interconnection system (15) are mutually proportional, the proportionality ratio being constant.

Abstract: A helicopter is fitted with a main rotor (1) having at least two blades (10, 20), each blade (10, 20) being provided with attachment elements (11, 21) for attaching it to a hub (2) of the rotor (1). The rotor is also fitted with one lift element (12, 22) per blade (10, 20), each lift element (12, 22) being mechanically linked to a single blade (10, 20) to vary the pitch of the single blade (10, 20) with which the lift element (12, 22) is linked.

Abstract: A method of controlling an overspeed safety system for an aircraft having at least two engines. The method includes setting said overspeed safety system for the engines, monitoring the speeds of rotation of the engines, detecting overspeed on one of the engines, shutting down the engine in question in the event of detecting overspeed, and inhibiting the operation of the overspeed safety system for the other engine(s) still in operation. The method also comprises the steps of monitoring safety parameters associated with the operation and/or the shutting down of the engine that has been shut down; and maintaining the inhibition or resetting the overspeed safety system for the other engine(s) still in operation, as a function of one or more safety parameters.

Abstract: A weighing system for detecting total weight, including optional external loads, and monitoring of center of gravity of a helicopter (2), comprising a fuselage (1), a landing gear (4), mounted to the fuselage (1) by flanges (10-13) and weighing cells (41-44). The weighing cells (41-44) are integral with the flanges (10-13) between the fuselage (1) and the landing gear (4). Attachment means are provided at the landing gear (4) for external loads (39, 40).

Abstract: The present invention relates to a method and to a device for controlling and adjusting the ground clearance (h) of an aircraft (1) provided with at least one undercarriage (30), said undercarriage (30) being provided with a damper-actuator (2) for performing a damping function and a function of retracting said undercarriage (30). The method is remarkable in that control means (10) automatically activate adjustment means (11, 12) for adjusting the length of said damper-actuator (2) so as to maintain said ground clearance (h) at a predetermined value.

Abstract: A power plant (2) having at least one engine (3, 4) and control means (5) for controlling said engine (3, 4). The control means (5) include a memory (6), said memory (6) containing information for operating said engine (3, 4) in accordance with at least two distinct utilization envelopes during a maximum number of flying hours, the two envelopes comprising an envelope enabling takeoff from a platform and another envelope enabling takeoff from takeoff zones not including platforms, each utilization envelope comprising at least two distinct utilization ratings each defined by a developed power and by a utilization duration for said developed power.

Abstract: An aircraft (1) having an airframe (2) extending along an anteroposterior plane of symmetry (P1) in elevation transversely separating a first side (11) of said aircraft (1) from a second side (12) thereof, said aircraft (1) being provided with hoist means (20) including a winch (21) provided with a drum for winding a hoist cable (22), said hoist means (20) including a beam (30) for guiding said cable, said beam (30) extending from a first end (31) constituting an inlet end for said cable (22) to a second end (32) constituting an outlet end for said cable (22). The hoist means (20) have a counterweight (23), said counterweight including said winch (21) that is fastened to said first end (31), said second end (32) being located on said second side (12) while said winch (21) and said first end (31) are located on said first side (11).