Abstract: A method of predictive determination of a power needed by a rotary-wing aircraft to perform hovering flight in order to land. After measuring the current atmospheric pressure and the current temperature outside the aircraft, the method serves to estimate a predicted atmospheric pressure and a predicted temperature at the altitude where the hovering flight is to be performed. Thereafter, both a maximum power available from a power plant of the aircraft once the aircraft has reached the point where the hovering flight is to be performed, and also a power needed from the power plant in order to perform the hovering flight are calculated. Finally, the present invention serves to display the available maximum power and the power needed on an instrument of the aircraft and thereby warn the pilot of the aircraft when the power needed is close to or even greater than the maximum power available.

Abstract: A method and a device for estimating the instantaneous mass of a rotary wing aircraft, the aircraft having a power plant driving at least one main rotor and an anti-torque rotor in rotation, and also having a plurality of sensors. The method uses functional characteristics and flight characteristics of the aircraft, atmospheric characteristics, and performance curves for the aircraft to determine the measured instantaneous mass Mm of the aircraft. Furthermore, the method makes it possible to consolidate the measured instantaneous mass Mm of the aircraft by comparing it in arithmetical or statistical manner during a flight of the aircraft with a calculated instantaneous mass Mc determined from the variation in the quantity of fuel present in the aircraft.

Abstract: A method of predictive determination of a power needed by a rotary-wing aircraft to perform hovering flight in order to land. After measuring the current atmospheric pressure and the current temperature outside the aircraft, the method serves to estimate a predicted atmospheric pressure and a predicted temperature at the altitude where the hovering flight is to be performed. Thereafter, both a maximum power available from a power plant of the aircraft once the aircraft has reached the point where the hovering flight is to be performed, and also a power needed from the power plant in order to perform the hovering flight are calculated. Finally, the present invention serves to display the available maximum power and the power needed on an instrument of the aircraft and thereby warn the pilot of the aircraft when the power needed is close to or even greater than the maximum power available.

Abstract: A method and a device for estimating the instantaneous mass of a rotary wing aircraft, the aircraft having a power plant driving at least one main rotor and an anti-torque rotor in rotation, and also having a plurality of sensors. The method uses functional characteristics and flight characteristics of the aircraft, atmospheric characteristics, and performance curves for the aircraft to determine the measured instantaneous mass Mm of the aircraft. Furthermore, the method makes it possible to consolidate the measured instantaneous mass Mm of the aircraft by comparing it in arithmetical or statistical manner during a flight of the aircraft with a calculated instantaneous mass Mc determined from the variation in the quantity of fuel present in the aircraft.

Abstract: A method of determining parameters that are characteristics of the operation of a vehicle having a power plant with at least one engine and mechanical transmission means, sensors, and display means. During the method, various items of information about the aircraft, its state and/or its operation and/or its environment are measured, and then, for at least one parameter Pi relating to the state and the operation of the aircraft there are determined a first limit value Pi_lim for the parameter Pi, a second value Pi_X for each parameter Pi to enable the aircraft to perform a predetermined maneuver X, and an instantaneous third value Pi_inst for each parameter Pi. Thereafter, each first, second, and third values Pi_lim, Pi_X, Pi_inst is displayed simultaneously in order to show up the relative position of second and third values Pi_X and Pi_inst relative to a first value Pi_lim for each parameter Pi.

Abstract: A method of determining parameters that are characteristics of the operation of a vehicle having a power plant with at least one engine and mechanical transmission means, sensors, and display means. During the method, various items of information about the aircraft, its state and/or its operation and/or its environment are measured, and then, for at least one parameter Pi relating to the state and the operation of the aircraft there are determined a first limit value Pi_lim for the parameter Pi, a second value Pi_X for each parameter Pi to enable the aircraft to perform a predetermined maneuver X, and an instantaneous third value Pi_inst for each parameter Pi. Thereafter, each first, second, and third values Pi_lim, Pi_X, Pi_inst is displayed simultaneously in order to show up the relative position of second and third values Pi_X and Pi_inst relative to a first value Pi_lim for each parameter Pi.

Abstract: The present invention relates to a piloting assistance system for an aircraft that is provided with a controller for determining piloting information. The system comprises at least one display system wearable on the head of a pilot, the display system being provided with a left display suitable for being viewed by a left eye of the pilot and a right display suitable for being viewed by a right eye of the pilot, the controller including a first processor for processing the information and a second processor for processing the information, the processors including stored instructions for causing the information to be displayed simultaneously and respectively on the left display and on the right display.

Abstract: The present invention relates to a device (10) for varying blade pitch of a rotary wing aircraft (50) having a main rotor (11) with a plurality of blades (12), each blade (12) including at least one main flap (13) fastened to the trailing edge of the blade (12). The angle of inclination of each flap (13) is controlled via a swashplate (20). The device (10) makes provision for electric actuators controlled by a flight control system (54) to be mounted in a stationary frame of reference for the purpose of moving and varying the angle of inclination of the non-rotary plate of the swashplate (20). The electric actuators provide primary flight control and also multi-cyclic control for the purpose of attenuating noise and vibration as generated in particular by the blades (12) and the rotor (11).

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: 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: A method of generating a terrain avoidance warning for a rotary wing aircraft including generating an avoidance trajectory including a proximal segment representative of a transfer time and an avoidance curve including at least one distal segment of a conic section curve following on from the proximal segment, wherein the proximal segment extends in continuation from a predicted trajectory over a distance representing an applicable reaction time, the applicable reaction time being minimized as a function of a route sheet for the aircraft, and wherein the generating includes calculating the at least one distal segment as a function of an instantaneous maneuverability of the aircraft.

Abstract: The present invention relates to a piloting assistance system (10) for an aircraft (1) that is provided with determination means (20) for determining piloting information. The system comprises at least one display system (40) wearable on the head of a pilot, the display system (40) being provided with left display means (41) suitable for being viewed by a left eye of said pilot and right display means (42) suitable for being viewed by a right eye of said pilot, said determination means (20) including first processor means (21) for processing said information and second processor means (22) for processing said information, the processor means including stored instructions for causing said information to be displayed simultaneously and respectively on the left display means (41) and on the right display means (42).

Abstract: A method of inspecting the integrity of an avionics system (1) installed in an aircraft, by inspecting the integrity in terms of electrical installation and data acquisition and transmission, wherein the method includes: using an Ethernet network (4) connecting the avionics system (1) to an inspection computer (3); using the computer (3) to generate first inspection signals transmitted to the component and/or to acquire second inspection signals coming from the component; using data read and write elements of a controller of the avionics system (1) to generate reference data relating to the component and transmitted to the inspection computer (3) via the Ethernet network (4); and displaying the reference data and the inspection signals and where appropriate recording the reference data and the inspection signals.

Abstract: A communications network for an on-board avionics system for conveying information between at least one on-board computer (1) and functional members such as sensors (2) and actuators (3), the network including analog links (6, 7) and digital bus links (8, 9a, 9b). At least one modular data concentrator (4) is configurable to digitize data from sensors (2) or to perform digital-to-analog conversions of data for actuators (3). A modular and configurable gateway (5) includes functions of receiving, acquiring, changing digital format, and switching information. Analog links (6, 7) are provided.

Abstract: A method of generating a terrain avoidance warning for a rotary wing aircraft including generating an avoidance trajectory including a proximal segment representative of a transfer time and an avoidance curve including at least one distal segment of a conic section curve following on from the proximal segment, wherein the proximal segment extends in continuation from a predicted trajectory over a distance representing an applicable reaction time, the applicable reaction time being minimized as a function of a route sheet for the aircraft, and wherein the generating includes calculating the at least one distal segment as a function of an instantaneous maneuverability of the aircraft.

Abstract: A method of inspecting the integrity of an avionics system (1) installed in an aircraft, by inspecting the integrity in terms of electrical installation and data acquisition and transmission, wherein the method includes: using an Ethernet network (4) connecting the avionics system (1) to an inspection computer (3); using the computer (3) to generate first inspection signals transmitted to the component and/or to acquire second inspection signals coming from the component; using data read and write elements of a controller of the avionics system (1) to generate reference data relating to the component and transmitted to the inspection computer (3) via the Ethernet network (4); and displaying the reference data and the inspection signals and where appropriate recording the reference data and the inspection signals.

Abstract: The maintenance device (2) comprises electronic equipments (4A, 4B, 4C) each comprising a test means (5A, 5B, 5C) and a memory (6A, 6B, 6C) for recording information relating to a fault, a central unit (8) comprising a test means (10), a memory (11), a computer (12) which receives information from, the electronic equipments (4A, 4B, 4C) indicating the test means that detected a fault, which determines, from this information, a faulty element and which indicates, on a screen (16, 17) while the complex system is in operation, the occurrence of a fault and the faulty element and, at the end of operation, a list of the faults that have occurred, with the faulty elements, a memory (13) for recording the information processed by the computer (12), and interfaces for accessing said memories.

Abstract: A method for harmonizing an equipment relative to a vehicle, such as a helicopter, based upon sensing of gravitational and magnetic fields in a number of reference systems. The equipment is installed fixedly on board the vehicle which, relative to a first absolute reference system (RM), has an orientation defined by a second reference system (R2) tied to the vehicle. The equipment is subjected to the earth's gravitational field represented by a gravity vector g oriented along its gradient, as well as to the earth's magnetic field represented by a vector H oriented along its gradient. The orientation of the equipment is defined, relative to the first absolute reference system (RM), by a third reference system (R3).

Abstract: A method for harmonizing an equipment relative to a vehicle, such as a helicopter, based upon sensing of the gravitational field in a number of reference systems. The equipment is installed fixedly on board the vehicle which, relative to a first absolute reference system (RM), has an orientation defined by a second reference system (R2) tied to the vehicle. The equipment is subjected to the earth's gravitational field represented by a gravity vector g oriented along its gradient. The orientation of the equipment is defined, relative to the first absolute reference system (RM), by a third reference system (R3). The equipment includes a device for measuring the components of the gravity vector g, and the vehicle carries a computer, a memory associated with the computer; and a link connecting the equipment to the computer and to the memory. The gravitational field is measured when the vehicle is in a number of different orientations, and the equipment is harmonized to the vehicle based thereon.