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 automatic method of regulating a power plant (3?) of an aircraft (1), the power plant having at least one turbine engine (3), each engine (3) being capable of operating in an idling mode of operation. A calculation system (15) executes stored instructions in order to implement the idling mode of operation as a function of operational and ordered conditions either via a first regulation mode by regulating a first speed of rotation (Ng) of said gas generator (4), or via a second mode of regulation by regulating a second speed of rotation (NTL) of said free turbine (7).

Abstract: A fluid temperature control installation for a rotorcraft, the installation comprising a system for cooling a main transmission (3) and a system for temperature control of the ambient air (32) of a cabin (16). A heat exchanger (8) is fitted with a bladeless ventilator (9). A source air stream (11) generated by an accessory compressor (5) driven by the main transmission (3) passes through the heat exchanger (8) where it takes heat. The source air stream (11) is distributed selectively to the bladeless ventilator (9) to generate a cooling air stream (13) passing through the heat exchanger (8), and/or is distributed to the cabin (16) after passing through the heat exchanger (8) in order to heat the ambient air (32) of the cabin.

Abstract: A system and a method for controlling a rotorcraft (1) in yaw (Y). The control system comprises a plurality of drive channels that operate an anti-torque rotor (4) of the rotorcraft (1). A first drive channel makes use of a rudder bar (13) including a set of pedals operated by the human pilot of the rotorcraft in order to control the anti-torque rotor (4). A second drive channel includes an autopilot. A third drive channel includes objective-type flight control means comprising a movable control member (12) that is operated by the human pilot and that continuously issues control signals (11) relating to a state of progression that the rotorcraft (1) is to achieve. Use of the second drive channel and of the third drive channel depends on inhibit means (19) for inhibiting operation thereof, which means are activatable by detector means (37) for detecting operation of the rudder bar (13).

Abstract: A blade (20) of a rotor (5), the blade (20) having a suction side (21) and a pressure side (22) extending transversely from a leading edge (23) towards a trailing edge (24) and extending spanwise from a root section (31) towards a free end section (41). The blade (40) comprises, going from the root section (31) towards said free end section (41): a root zone (30); followed by rounded zone (35); said rounded zone (35) presenting rounded pressure and suction sides (22?, 21?) departing from said main plane (P1) in a direction (Z) parallel to the axis of rotation (AXROT) of said blade.

Abstract: A pitching stabilization means (10) having at least one stationary stabilization surface (20) extending in a thickness direction from a bottom face (21) to a top face (22) and in a transverse direction from a leading edge (23) towards a trailing edge (24). The stabilization surface (20) has at least one slot (30) passing through said thickness of the stabilization surface (20) from said top face (22) to said bottom face (21), said slot (30) being arranged within the stabilization surface (20) between said leading edge (23) and said trailing edge (24) so as to allow a flow of air coming from a rotor to pass from said top face (22) towards said bottom face (21).

Abstract: A regulator device (10) for regulating a turbine engine (3). The regulator device (10) includes mechanical power take-off means (100) for taking off power mechanically from a gas generator (4), and an engine computer (20) controlling said engine (3) to comply with at least a first limitation (LimTET, LimT45) of a temperature (TET, T45) of the gas within the engine, and with a second limitation (LimNg) of a speed of rotation (Ng) of the gas generator (4). The engine computer (20) determines whether the speed of rotation (Ng) of the gas generator has reached said second limitation (LimNg), and whether said temperature (TET, T45) has reached said first limitation (LimTET, LimT45). An avionics computer (30) causes the mechanical power take-off means (100) to operate if the speed of rotation (Ng) of the gas generator (4) has reached said second limitation (LimNg), and if said temperature (TET, T45) has not reached said first limitation (LimTET, LimT45).

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: 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 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: 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).

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 method of providing a volume-mass law for determination of a fuel flow rate of an engine, particularly providing a fuel flow rate to a helicopter turbine, comprising the steps of: determining a sample type of fuel and a start density ?0 of said sample type of fuel in said fuel tank using an equation ?0=aT+b0, with a and b0 being known for said sample type of fuel and calculating real time offset parameters bn from an algorithm to determine real time densities ? of the fuel.

Abstract: A method of automatically regulating a power plant (3?) of an aircraft (1), said power plant comprising at least one turbine engine (3), said aircraft (1) having at least one rotary wing (300) provided with a plurality of blades (301) having variable pitch and driven in rotation by said power plant (3?), it being possible for each engine (3) to operate in an idling mode of operation and in a flight mode of operation. During a selection step (STP0), a two-position selector (60) is operated either to stop each engine (3) or to set each engine (3) into operation. During a regulation step (STP1), each engine (3) is controlled automatically so as to implement the idling mode of operation if the collective pitch (CLP) of said blades is less than a threshold and if the aircraft (1) is standing on the ground.

Abstract: A method of failure simulation for an aircraft (7) having a power plant (10) with at least two turbine engines (11, 12), together the two engines develop an overall power, each engine (11, 12) being capable of delivering at least one contingency power in order to compensate for a total failure of other engines (11, 12). The device serves during a failure simulation to modify the overall power delivered by the power plant, with this modification being performed with the help of first adjustment means (20). Second adjustment means (30) serve to modify also the difference between the minimum power obtained during the simulated failure and the stabilized overall power, and also the time between said failure stabilizing on said overall power.

Abstract: A regulator device (10) for reducing the risk of surging in a turbine engine (3) that includes a gas generator (4), air extraction means (8), and mechanical power take-off means (100). An engine computer (11) includes storage means (16) that store a plurality of acceleration regulation relationships, each acceleration regulation relationship corresponding to air extraction in a first range, and to mechanical power take-off in a second range, said regulator device (10) including first measurement means (20) for measuring current air extraction, and second measurement means (30) for measuring current mechanical power take-off, said engine computer (11) controlling acceleration of the engine (3) by implementing the acceleration regulation relationship corresponding to the current air extraction and to the current mechanical power take-off.

Abstract: A method of automatically performing an engine health check for checking the health of at least one turbine engine of an aircraft. During flight, the stability of at least one monitoring parameter is verified by acquiring a measurement signal, by performing first filtering of each signal by a high-pass filter over a long first duration (TPS1) and by verifying that a first amplitude (A1) of the signal filtered in this way does not exceed a first threshold defined by the manufacturer, by performing second filtering of each signal by a high-pass filter over a short second duration (TPS2) in parallel with said first filtering by a high-pass filter, and by verifying that a second amplitude (A2) of the signal filtered in this way does not exceed a second threshold defined by the manufacturer, the second duration (TPS2) being less than the first duration (TPS1), and the second amplitude (A2) being less than the first amplitude (A1).

Abstract: An interface and a method for mounting a piece of equipment (5) with a rotary drive source (3). The mounting interface has a first frame (1) fastened to the drive source (3) and a second frame (2) fastened to the piece of equipment (5), the frames (1, 2) being fitted with assembly means (19) for assembling them together. Rails (22, 23) for guiding the piece of equipment (5) are removably mounted on the first frame (1) and co-operate with windows (20, 21) of the second frame (2). The windows (20, 21) are laterally open providing transverse passages for the rails (22, 23) into the windows (20, 21). The piece of equipment (5) can be moved transversely by an operator towards the rails (22, 23) in order to support it and guide it axially towards the drive source (5).

Abstract: The invention relates to a drainage circuit (4) for a flow of a liquid (5), the circuit being fitted with a monitoring appliance (13) for monitoring an excess flow rate of the liquid (5). The monitoring appliance (13) comprises a monitoring duct (14) interposed between an upstream duct (7) for collecting the liquid (5) and a downstream duct (9) for discharging the liquid (5). The monitoring duct (14) has retaining means (15) for retaining part of the liquid (5) inside a chamber (16) arranged to receive a spillway (22) of the captured liquid (19) in a reserve (20) fitted with indicator means (23) with a visual scale for providing warning information relating to the quantity of liquid (24) contained in the reserve (20).

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.