Abstract: A coupling means (10) provided with a first member (20) suitable for being fastened to a first rotary part (2) and with a second member (30) suitable for being fastened to a second rotary part (3), the first member (20) being provided with a first diaphragm (22) and the second member (30) being provided with a second diaphragm (32) that is secured to the first diaphragm (22). An emergency torque transmission device comprises at least one bayonet system including a protuberance (50) secured to one member (30) and co-operating with an angled groove (60) secured to the other member (20) by being inserted in the angled groove (60) by a thrust-and-rotation movement, in the absence of said breakage said coupling means (10) including both axial clearances in compression (70?) and in translation (70?) and also circumferential clearance between each protuberance (50) and walls defining the corresponding angled groove (60).

Abstract: The present invention relates to a control and regulation method for a rotorcraft having at least one variable-pitch propulsive propeller driven by at least one power source, said method consisting in generating at least one mean pitch setpoint ?tcl* for the propeller(s) as a function of a thrust variation control order Tcl, wherein the method consists in defining a plurality of operating modes, including: a direct mode in which the value of the mean pitch value is a direct result of the control order Tcl; a forced mode in which the mean pitch is automatically forced to a calculated pitch value; a regulated mode in which the power of the propulsive propeller(s) is regulated as a function of a power setpoint from a pilot and of servo-controlling the mean pitch of the propeller(s); and a protected mode.

Abstract: A blade attachment (1, 20) for a bearingless main rotor of a helicopter with an airfoil blade (2), a flexbeam (3, 21) including a flexbeam body (16, 17) and a flexbeam head (13) at an end of said flexbeam body (16, 17). A control cuff (4, 22) encloses and extends along at least a predominant portion of said flexbeam (3, 21). 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. Said removable fasteners comprise a main bolt (7) and at least one supporting bolt (8).

Abstract: An airfoil blade (2) of a bearingless rotor of a helicopter. Said airfoil blade (2) has a pitch axis from a tip end to a root end and said root end is provided with junction means for a separable junction arrangement of said airfoil blade (2) with a flexbeam head (13), and a control cuff (4, 22). Said junction arrangement is mechanical with removable fasteners respectively removable connecting said root end of said airfoil blade (2) to said control cuff (4, 22) and said flexbeam head (13). Said root end comprises at least two openings (7, 8) being asymmetric with regard to said pitch axis.

Abstract: A standby instrument (10) for an aircraft, the instrument comprising at least one inertial sensor (1), at least one pressure sensor (2), calculation means (3) connected to said inertial and pressure sensors (1, 2), a display unit (4). Said calculation means (3) are suitable for determining critical flight information for said aircraft, and for displaying said critical flight information on the display unit (4) in the event of a main information system of said aircraft failing. In addition, said standby instrument (10) also incorporates stabilization relationships enabling said calculation means (3) to determine control relationships in order to control the actuators (15) of an autopilot of said aircraft in the event of said autopilot failing. Finally, said calculation means (3) are connected to at least one engine operation computer (5) enabling said instrument (10) to display information about a first limit of the engine on said display unit (4).

Abstract: An airfoil (10) having a main structure (15) including at least one spar (16) and a front structure (20) forming a leading edge (21), said airfoil (10) having an empty space (25) between said main structure (15) and said front structure (20). The airfoil (10) includes a rigid deflector member (30) in said empty space (25), the deflector member being fastened to the main structure (15), said deflector member (30) having a sharp edge (31) facing towards said front structure (20) in order to deflect an external obstacle impacting against said front structure (20).

Abstract: The present invention relates to a method of fabricating a painted substrate (30) using powder paint (21). The substrate is electrically insulating. The method includes the following steps performed in succession: a) a treatment coating (10) is placed on a mold (1), the treatment coating (10) includes at least one preparation layer having a release agent (15, 15?) in contact with the mold (1); b) the treatment coating (10) is electrostatically covered with powder paint (21) in such a manner that the powder paint (21) is attracted towards the mold (1) and is held against the treatment coating (10); c) the powder paint (21) is transformed at least in part in order to obtain a layer (20) of powder paint (21); and d) the substrate (30) is prepared on the layer (20) of powder paint to form the painted substrate.

Abstract: The present invention relates to a method of draining a tank (2) having a plate, the plate (12) including a canister (19) having an outer casing (20) provided with a base (19?) that is pressed against the plate (12), the base (19?) presenting at least one through orifice (21) suitable for being closed by a movable member (22) of the canister (19), the plate (12) carrying a pump (30) that co-operates with the canister (19) by actuating the movable member (22) to open the through orifice (21), during which method the pump (30) is removed from the canister (19) and then a fluid connector (50) is inserted to take the place of the pump (30), the fluid connector (50) co-operating with the canister (19) by actuating the movable member (22) to release the through orifice (21) in order to drain the tank (2) via the fluid connector (50).

Abstract: An energy absorber element comprises an absorber portion having first and second distal ends and having a first elongate element and a second elongate element facing the first elongate element and connected to the second element solely at the first and second distal ends, a first anchor point secured to the first distal end of the absorber portion, and a second anchor point secured to the second distal end of the absorber portion, wherein each of the first and second elongate elements includes a succession of straight-line segments parallel to one another and at least one circularly-arcuate segment such that at least two successive straight-line segments are connected together by one of the at least one circularly-arcuate segment.

Abstract: A method of controlling a buoyancy system for an aircraft includes: determining the roll angle ? and the pitching angle ? of the aircraft; verifying whether ??R <+R and whether ??R <?<+?R, where ?R and ?R are predefined limit angles; if at least one of the angles ? and ? is no longer in its above-defined respective range, activating an automatic trigger of the buoyancy system; if the angles ? and ? are in their above-defined respective ranges, determining the altitude A of the aircraft; inhibiting the automatic trigger if A >AR, where AR is a predefined limit altitude; and if AR ?A, and if at least partial immersion of the aircraft has been detected, activating the automatic trigger.

Abstract: The present invention relates to a piloting assistance method for an aircraft, the method consisting in using data from at least one active telemeter sensor (A) in order to construct a sensor safety cordon (B) for avoiding the terrain and obstacles that are overflown. The method; defines and calculates angular sectors (w) over the field of regard facing the pilot; constructs a terrain safety cordon (D) using at least one terrain database (C); for at least some of the angular sectors (w), constructs a hybrid safety cordon (E) that, in each of the angular sectors (w) in question, makes use of the higher of the sensor and terrain safety cordons (B, D); and displays one of the cordons selected from: the hybrid safety cordon (E), the terrain safety cordon (D), and the sensor safety cordon (B).

Abstract: The invention relates to a skid-type landing gear (1) for a helicopter (2) comprising bow shaped crosstubes (10-13) directed symmetrically towards a left- and a right-hand side of a longitudinal mid axis (9) of the helicopter (2) with respective skids (14, 15) at outboard ends thereof and main beams (5-8) in a fuselage (4), the crosstubes (10-13) being connected to the main beams (5-8) via supports (16-23) fixed to the main beams (5-8). The skid-type landing gear (1) is retractable. A method for operating such a skid-type landing gear (1) for a helicopter (2) is provided.

Abstract: A method of transferring and processing flight data of an aircraft including recording the aircraft flight data in an onboard avionics system in a digital format so as to obtain digital data, the flight data including safety impact flight data and flight data having no recognized effect on safety. The method further includes recording at least the safety impact flight data in the onboard avionics system in an additional format dissimilar to the digital format so as to obtain additional flight data, storing the digital data and the additional flight data separately, transferring the digital data and the additional flight data to a monitoring computer on the ground, displaying the digital data on a screen in the form of display digital data, and transferring the digital data and the additional flight data to a complementary interpretation unit so as to generate one of display data and print data relating to the additional flight data.

Abstract: A landing gear vibration absorber (1) of a helicopter (2) with a landing gear (3) comprising a pair of skids (4, 5) and at least one cross tube (6, 7) for mounting the skids (4, 5) to a helicopter's fuselage (8). At least one spring-mass system (10, 20, 24) is mounted to the landing gear (3). Said at least one spring-mass system (10, 20, 24) is tuned to the helicopter's main excitation frequency and said at least one spring-mass system (10, 20, 24) being located at or near at least one antinode of the landing gear (3). The invention is related as well to a method of operating a landing gear vibration absorber (1) of a helicopter (2).

Abstract: A proximity warning system for a helicopter (22) comprising at least two radar units (1-3), preferably three radar units (1-3) arranged to transmit microwaves and receive reflections of said microwaves from obstacles (10). The at least two radar units (1-3) are fixed next to a main rotor head(s) (20) of the helicopter (22) for horizontally scanning an entire environment of 360° around the helicopter (22), all of said at least two radar units (1-3) operating essentially at the same frequency.

Abstract: A method of adapting a man-machine interface (20) of an aircraft (50) depending on the functional level of a pilot, the method comprising a plurality of successive steps. Firstly, before starting a mission, the characteristics of said mission and the characteristics and the physiological state of said pilot are determined. Thereafter, during said mission, the state of the mission and the current state of said aircraft together with a current behavior of said pilot are determined and a current functional level of said pilot is estimated. Thereafter, said current functional level of said pilot is compared with reference functional levels, and said man-machine interface (20) is then adapted in order to assist said pilot automatically and in optimum manner in making the pilot aware of the situation, in the pilot's decision-making, or in the actions taken by the pilot depending on the pilot's stress state or work load state.

Abstract: A rotary wing aircraft (1) provided with a main lift rotor (2), a tail rotor (5), and a power plant (4) driving a main gearbox (3) that co-operates with said main rotor (2), said tail rotor (5) being provided with a plurality of blades (10) of variable pitch (I) and with a pitch modification device (20), and said aircraft (1) having control means (30) for controlling said pitch modification device (20). The aircraft (1) includes an electric motor (9) for rotating said tail rotor (5) and regulator means (TRCU) connected to the control means (30) and also to the electric motor (9) and to the pitch modification device (20). The regulator means (TRCU) generate a first setpoint concerning pitch that is transmitted to the pitch modification device (20), and a second setpoint for controlling a motor parameter that is transmitted to the electric motor (9).

Abstract: A detection system and method for detection of damages on at least one rotating component (10, 20, 30. . . N) of an air vehicle caused by at least one solid, comprising measuring means (S1, S2, S3 . . . Sn) for measuring vibration and/or sound of said rotating components (10, 20, 30 . . . N), to acquire the time signal of vibration and/or sound of any of said rotating components (10, 20, 30 . . . N), and processing means (PRO) for comparing said measured vibration and/or sound with a maximum value or an envelope of standard vibration and/or sound stored in a memory (EM1, EM2, . . . EMn) and warning means (WS) for issuing a warning signal when the measured vibration and/or sound exceeds said standard vibration and/or sound. The air vehicle is an airplane or a helicopter, the solid is a projectile, stones, ice-crystals or bird strike, and the measuring means are vibration sensors (S1, S2, S3 . . . Sn) and/or acoustic sensors (S1, S2, S3 . . . Sn).

Abstract: A shock and impact resistant multilayered composite comprising an outer shatter protection body, and at least one active body along side the shatter protection. At least one highly elastic and tearproof absorption/dissipation body is provided along side the at least one active body, said outer shatter protection body being cured to said at least one active body and said at least one active body being cured to said at least one absorption/dissipation body. The present invention relates as well to a method for fabrication of such a shock and impact resistant multilayered composite.

Abstract: An air vehicle with a slip protecting and gas sealing composite floor (1, 10) inside a fuselage, particularly a helicopter with a slip protecting and gas sealing composite floor (1, 10) inside a fuselage. Said slip protecting and gas sealing composite floor (1, 10) is built up of a profiled elastomer layer (2) provided at its bottom side with a cross-linking agent, n-layers (3) of a further component or a variety of different components and partially thermosetting resin covered by a profiled elastomer layer (2), a honey comb layer (4) and lower n-layers (5) of said further component or variety of different components and partially thermosetting resin, said honey comb layer (4) being sandwiched between said n-layers (3) and said lower n-layers (5).