Abstract: A system and method for detecting yoke interference for an aircraft having an auto pitch trim function is provided. The system includes a source of elevator load data, a source of aircraft speed data, and a processing system. The processing system is coupled to receive the elevator load data, the aircraft speed data, and initial condition center-of-gravity (CG) data that is representative of at least an estimated initial position of the CG of the aircraft. The processing system is configured to process at least the speed data and the initial condition CG data to: (i) determine an expected elevator load on the elevator flight control surface, (ii) determine if the expected elevator load exceeds the sensed elevator load by a predetermined magnitude, and (iii) when the expected elevator load exceeds the sensed elevator load by a predetermined magnitude, generate a disconnect signal that will disable the auto pitch trim function.

Abstract: A flight control system having a plurality of dual mode operator control inputs is disclosed and includes a plurality of active parallel actuators, one or more processors, and memory coupled to the one or more processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the flight control system to receive a signal indicating an airspeed of the compound aircraft and select between rotary and fixed wing modes of operation based on the airspeed. In response to selecting a mode of operation, the flight control system sends either a rotary or a fixed wing force feel profile to the plurality of active parallel actuators, where the force feel profile defines the respective detent force gradient, where the fixed wing detent force gradient is at least about two times greater than a rotary wing detent force gradient.

Abstract: A linear electromechanical actuator with a main screw-nut assembly driven by a main motion device and having a hollow screw with an abutting surface; an anti-jamming piston arranged coaxially within the screw and shiftable between an engaged position in which locking dogs interfere with the abutting surface and a disengaged position in which the piston is free to slide within the screw; and actuating elements configured to shift the piston from the engaged to the disengaged position upon electrical or mechanical failure of the actuator. The actuating elements include a key axially movable between the engaged and disengaged positions and having a locking section, configured to bias the locking dogs into interference with the abutting surface in the engaged position, and an unlocking section, configured to allow free sliding of the piston within the screw in the disengaged position. An anti-jamming device for operating a critical flight control surface.

Abstract: A method for monitoring an aircraft piloting device including at least one piloting member (20, 30) and at least one fly-by-wire flight control system (40, 41). At least one monitoring module is integrated into this control system and is adapted to compute, on the basis of primary signals processed by sensors associated with at least one piloting member, at least one theoretical value of at least one monitored parameter of at least one piloting member, to compare each theoretical value with measurement signals of each monitored parameter and to select a monitoring action, particularly to generate monitoring signals (55, 56), as a function of the difference between each theoretical value and the measurement signals.

Abstract: A power supply and control device for a plurality of electric actuators, the device comprising a control circuit and a power circuit, the power circuit having an inverter module connected to a switch module provided with connection means for connection to the actuators, the control circuit being connected to the switch module to operate it so as to connect the actuators sequentially to the inverter module, wherein the control circuit is arranged to control the inverter module and the switch module sequentially, and wherein the control circuit includes a monitoring module arranged to detect a malfunction of the device. An actuator assembly comprising a plurality of actuators and such a power supply and control device, and an aircraft including a plurality of actuators and such a power supply and control device of the above-specified type.

Abstract: A multi-axis serially redundant, single channel, multi-path fly-by-wire control system comprising: serially redundant flight control computers in a single channel where only one “primary” flight control computer is active and controlling at any given time; a matrix of parallel flight control surface controllers including stabilizer motor control units (SMCU) and actuator electronics control modules (AECM) define multiple control paths within the single channel, each implemented with dissimilar hardware and which each control the movement of a distributed set of flight control surfaces on the aircraft in response to flight control surface commands of the primary flight control computer; and a set of (pilot and co-pilot) controls and aircraft surface/reference/navigation sensors and systems which provide input to a primary flight control computer and are used to generate the flight control surface commands to control the aircraft in flight in accordance with the control law algorithms implemented in the flight

Abstract: An actuator which can be installed in a system for the virtual simulation of aircraft, ships and vehicles to help the system with the controls. More specifically, the invention is an actuator used for a control loading system, which can be installed into a virtual simulation system, receive force delivered by the movement of a control stick laterally while measuring the applied force. It can connect a lever which delivers the control force equivalent to the applied force and reaction force by making use of a direct drive motor, including a potentiometer which senses the angle of rotation while sensing the initial reference position separately from an encoder, limiting the rotation range of the rotating lever, fixing the initial neutral position, and the control loading system for actuators manufactured compactly on the whole.

Abstract: An inceptor is provided including a control stick mounted to a seat. The control stick is movable about at least a first axis. A fly-by-wire inceptor control system is housed within a portion of the seat. The inceptor control system is configured to provide input signal indicative of movement of the control stick to a flight control system.

Abstract: A control system for an aircraft including a passive feedback arrangement, a stick and a positioning arrangement is provided that provides for user adjustability of the resistance applied to the stick, such as in a fly-by-wire system. The passive feedback arrangement is movable relative to a mechanical ground. The stick is moveable relative to the mechanical ground and the passive feedback arrangement. The passive feedback arrangement acts on the stick to resist movement of the stick relative to the passive feedback arrangement. This resistance provides passive feedback to the pilot for the fly-by-wire system. The positioning arrangement is coupled to the passive feedback arrangement for adjusting the position of the passive feedback arrangement relative to the mechanical ground in response to movement of the stick relative to the mechanical ground.

Abstract: A monitoring and control device for an aircraft actuator includes a control module, delivering control signals for the actuator and position signals for the actuator determined according to control messages received from a piloting management system of the aircraft, and at least one first position sensor supplying information concerning the position of the actuator, and a monitoring module, delivering position signals for the actuator and receiving the control signals received from the piloting management system of the aircraft, and information relating to the position of the actuator supplied by at least one second position sensor. The control and monitoring modules are capable of assessing the consistency of the signals processed therein and of controlling accordingly a power supply and disabling module of the actuator.

Abstract: A device for piloting includes a piloting member connected to at least one member for driving the craft, an electromechanical support box including, for each pivot connection, at least one actuating motor of the piloting member constituted of an electromagnetic actuator including a movable armature integrally formed with the piloting member and equipped with at least one permanent magnet having a magnetic moment parallel to the axis of the pivot connection, and an electromagnetic circuit including a fixed armature, arranged to allow the movable armature to move in the air gap zone, the magnetic moment of the movable armature sweeping each radial surface portion, and at least one coil winding which is dependent in position on the fixed armature and is capable of generating electromagnetic torque on the movable armature.

Abstract: A device for determining a position of a throttle lever includes position sensors each connected to a primary flight control computer, position sensors each connected to an engine management computer, the position sensors being split into at least three groups of sensors with no common simple failure mode, and at least one interface computer, that includes at least one input to receive measurement information emitted by the computers connected to the position sensors and outputs leading to the engine management computer.

Abstract: A power supply and control device for a plurality of electric actuators, the device comprising a control circuit and a power circuit, the power circuit having an inverter module connected to a switch module provided with connection means for connection to the actuators, the control circuit being connected to the switch module to operate it so as to connect the actuators sequentially to the inverter module, wherein the control circuit is arranged to control the inverter module and the switch module sequentially, and wherein the control circuit includes a monitoring module arranged to detect a malfunction of the device. An actuator assembly comprising a plurality of actuators and such a power supply and control device, and an aircraft including a plurality of actuators and such a power supply and control device of the above-specified type.

Abstract: A piloting method and device for avoiding pilot induced oscillations filters a piloting order with a non-linear filtering function that is controlled by speed comparison information.

Abstract: A system includes a plurality of fly-by-wire control units to control flight control surfaces of an airplane, wherein each control unit is adapted to operate independently of other control units and directly control an actuator to control a control surface as a function of direct pilot control device input.

Abstract: An aircraft flight surface control system and method simultaneously provides the benefits of both an active/active system architecture and in active/standby system architecture. The system is preferably implemented using hydraulic actuator assemblies and electromechanical actuator assemblies coupled to the same flight control surface. During normal system operations the electromechanical actuator assemblies are energized to supply a relatively minimal force to associated flight control surfaces. In effect, the electromechanical actuators, although energized, may be pulled along by the associated hydraulic actuator assemblies, until needed. Thus, the electromechanical actuator assemblies are controlled in a manner that closely resembles the active/standby architecture.

Abstract: A system and method of controlling a rotary-wing aircraft includes transforming a non-rate limited servo position command into a rate limited servo position command.

Abstract: An electronic operational control device for an aircraft piloting device with two connected piloting members, includes electronic circuits for main monitoring from signals delivered by sensors associated with one of the piloting members, and at least one electronic circuit (52 to 55) for cross-monitoring, for digital processing of signals delivered by sensors (83, 93) associated with the other piloting member, adapted to detect any deviation of these signals corresponding to a fault and to generate a signal representing such a fault. A piloting device and an aircraft including such an electronic operational control device with cross-monitoring are also disclosed.

Abstract: Embodiments of the invention relate to a flight control system for controlling an aircraft in flight having a backup control system integrated into an active control stick. The actuated control stick may include a processing unit that includes independent and separate hardware and/or software dedicated to the primary control system and the backup control system. For the primary control system, the processing unit may receive a sensed primary control stick signal and communicate with a primary processor, which may be configured to generate a primary control signal. For the backup control system, the processing unit may receive a sensed backup control stick signal and generate a backup control signal. The processing unit may also generate tactile signal for use by the actuated control stick to adjust the feel of a pilot's control stick.

Abstract: A fly-by-wire control system for an aircraft including detection of pilot induced oscillations and a mobile control unit for such a system, According to the invention, provision is made for the pivotings of the mobile control unit to be damped by a controllable damper controlled by a detector that detects oscillations corresponding to pilot induced oscillations.

Abstract: The invention relates to an electric flight control system for aircraft elevators. According to the invention, the flight control system can be controlled in terms of load factor or rate of pitch. The inventive system comprises built-in protections in relation to load factor, incidence and pitch attitude.

Abstract: Electromechanical actuation systems and methods are provided. The electromechanical actuation system includes first, second, and third linear actuators having respective first, second, and third ranges of motion and an output member coupled to the first, second, and third linear actuators such that a position of a selected portion of the output member is based on actuation of the first, second, and third linear actuators.

Abstract: A pedal system particularly tailored to a Fly-By-Wire (FBW) flight control system includes a double gradient linkage assembly. The double gradient linkage assembly includes a damper system and a spring system that improves yaw axis (azimuth) control of the aircraft. Control is only required when a change in the yaw axis state is demanded. Since only minimal displacement inputs to such a FBW flight control system is required, the travel of the pedals are exceeding compact, such that pilot workload is significantly reduced through the reduction in the frequency and magnitude of aircraft control inputs.

Abstract: An apparatus and method for rotary code-based control for command systems. A flight deck module comprises a knob connected to a shaft; a plurality of sensors affixed to a surface of a printed circuit board, the printed circuit board connected to the shaft; and a carrier platter. The carrier platter is connected to the shaft and a plurality of sources is affixed onto a first side of the carrier platter in an arcuate orientation (such as in a circular fashion along the circumference of the carrier platter or spanning a sector of a circle).

Abstract: The invention concerns a system (1) comprising a steering vane (2) which can turn within a rudder deflection range which is limited by first and second rudder deflection limits and means (12) for asymmetrically varying said first and second rudder deflection limits based on the current values of the aircraft flight parameters.

Abstract: An apparatus is described comprising: an aircraft having one or more control surfaces; and one or more micro-electro mechanical systems (“MEM S”) actuators to control the movement of the one or more control surfaces.

Abstract: An apparatus for operating a controlled member on a rotary-wing aircraft is provided with at least one control intended to be subject to the action of a pilot of the rotary-wing aircraft and capable of moving a linkage, a device connected to the linkage for operating the controlled member as a function of movement of the linkage, a ram incorporated into the linkage which is capable of influencing the movement of the linkage as a function of control commands received, a trim device capable of acting on the control and on the linkage as a function of control commands received, a computer determining the control commands which are transmitted to the ram and to the trim device, and at least one sensor which is capable of measuring the values of a parameter that represents an action exerted by a pilot of the rotary-wing aircraft on an auxiliary control.

Abstract: The subject tiltrotor aircraft has three modes of operation: airplane mode, helicopter mode, and transition mode. A tilting mast, which transitions the aircraft between airplane mode and helicopter mode, is controlled by systems that allow selective movement of the rotor blades between the flight modes. A hub couples the rotor blades to the tilting mast such that torque and thrust are transferred, while allowing rotor thrust vector tilting. A main swash plate controls rotor thrust vector direction. Pitch horns are coupled to the rotor blades and the main swash plate via pitch links such that swash plate inputs are communicated to the rotor blades. The pitch links are coupled at “delta-3” values that are not optimum. A feedback swash plate and feedback links receive disk tilting inputs from the rotor blades, and supply inputs to the main swash plate, which compensates for the less than optimum delta-3 coupling.

Abstract: An electrical fly-by-wire system for operating an aircraft rudder includes a low-pass filter, arranged between a rudder bar and an actuator of a rudder. The low-pass filter receives a control command from the rudder bar corresponding to the degree of travel the rudder bar has experienced from a neutral position. Based on the amplitude of the control command, the filter generates an operating command for the actuator. Additionally, the filter operates such that the higher the fraction of the rudder bar's travel away from the neutral position, with respect to its maximum value of travel, the higher the filter's time constant is set.

Abstract: A control apparatus comprises a control stick, a sensor for detecting the angular displacement amount of the stick and outputting a position signal, a spring for generating an operation force in accordance with the angular displacement amount based on a neutral point of the stick, a back drive unit coupled to the stick via the spring, for driving the stick to move the neutral point; and a controlling section for controlling the unit, the unit including a differential gear section having input shafts and an output shaft, a back drive motor coupled to the input shaft, an electromagnetic brake coupled to the input shaft, and an output arm coupled to the output shaft to drive the stick, and when the stick is being moved, the electromagnetic brake is put into disconnection to release the input shaft, the motor is rotated in the reverse direction, thereby preventing the occurrence of a transient.

Abstract: A position servo-control device comprising at least one servo-control loop and a servo-controlled system placed in said loop, said loop having compensation means which comprise means for filtering at one or more resonant frequencies concerning displacement for the mechanical part(s) of position controlled by the servo-controlled system, the device including compensation means implementing filtering at one or more frequencies constituted by resonant frequencies for the forces exerted on the servo-controlled system and/or on the mechanical parts it controls.

Abstract: The present invention relates to a flight control device for an aircraft comprising a control (2) and a means for actuating a controlled member (RP, RQ) to which a command is applied. According to the invention, said device (1) additionally includes a sensor (E) for determining a second value that is representative of the control executed by the aircraft (He) with respect to the control axis, and second means (M2) which determine: as long as the second value is lower than or equal to a reference value, a first trim command that is proportional to the actuation of the control (2); and when the second value is higher than the reference value, in addition to a second trim command, a speed command that is proportional to the additional actuation over and above said second value.

Abstract: System for controlling an aerodynamic surface (1) (flap, aileron, control surface, etc.) of an aircraft. According to the invention, this system comprises: an electric machine (8) which can operate either as a motor or as a generator, the output of which is connected mechanically to said aerodynamic surface (1); and a logic device (17) receiving information about the operating status of the electrical power supply (11) of the machine (8) and of the hydraulic servocontrol (2, 5) and controlling the actuation of said surface (1) either by said machine (8) or by said servocontrol (2, 5).

Abstract: An airplane high lift surface drive system having a first actuator connected to a first high lift surface and a second actuator connected to a second high lift surface, wherein the first actuator is electrically connected to the second actuator in series.

Abstract: A device for the control of a control surface of an aircraft has at least two actuators (110, 112, 114), each having at least one electric control input (111a, 113a, 115a). At least one of the actuators (115), called the mixed actuator, also has a mechanical control input (115b), and an electric control system for the actuators is able to occupy a state corresponding to a fault of an engine in which at least two of the actuators simultaneously operate the control surface and a state corresponding to an electric fault, in which the mixed actuator (115) operates the control surface from the mechanical control input.

Abstract: An actuator for a primary control surface of an air-craft, the actuator comprising at least one motor of the vibration type mounted in series with at least a second motor which is a piezoelectric, electrostrictive, or magnetostrictive actuator of the direct type, and further including control means causing one or both of the motors to operate in such a manner that the second motor drives movements of the control surface of the type used for stabilizing the aircraft, while the vibration drives movements of greater amplitude.

Abstract: A flying aid device for a fly-by-wire aircraft which has at least two mechanically independent control columns (M1, M2). The device (1) includes: a system (SO) for detecting multiple control corresponding to simultaneous operation of at least two of the control columns (M1, M2); and informing systems (S1, S2, S3, S4), each of which is associated with one of the control columns (M1, M2) and is intended to generate within the control column (M1, M2), when the detection system (SO) detects multiple control, a tactile sensation that informs a pilot of the aircraft who is operating the control column (M1, M2) of the multiple control.

Abstract: An electrical mechanical blow-down mechanism for use with an aircraft spoiler avoids possible damage due to high inertial loading when a power failure occurs as the spoiler (10) is moving toward or is at an extended position by means of a structure including a ball screw mechanism (18) adapted to be connected to the spoiler (10) and including an input gear (34) adapted for rotation about an axis (36). The mechanism is ultimately driven by an electrical motor (20) having a rotary output (22) with at least one gear (26, 28) interconnecting the input gear (34) and the rotary output (22). A pawl (36) is mounted for movement between positions engaged with and disengaged from the input gear (34) and a spring (78) is utilized to normally bias the pawl (36) toward the engaged position. A solenoid (50) is operable to move the pawl (36) against the spring (78) toward the disengaged position and a motion sensor (90) is associated with the input gear 34 to determine the direction of rotation thereof.

Abstract: A system for controlling the tab (6) of an aircraft control surface (3) having a pair of position sensors (12 and 15) that supply information regarding the turning of the control surface (3) and the position of a jack (13), respectively, on the basis of signals generated by the sensors (12 and 15) as well as possibly on the basis of at least one parameter (p) originating from the aircraft, such as airspeed, positions of lift-augmenting devices, forces exerted by the pilot on a control, etc., and which formulates a command for the jack (13).

Abstract: The present invention relates to a system for coupling at least two control columns (2A, 2B) of a control device (DC) of a machine, especially an aircraft. The coupling system includes for each of the columns (2A, 2B), at least one operating member (13A, 13B, 14A, 14B) capable of making the column (2A, 2B) tilt about an axis, and a calculation unit (CAL) receiving information relating to the position of tilt of the columns (2A, 2B) and generating, when at least one of the columns (2A, 2B) is operated, tilt commands which are transmitted to the operating member (13A, 13B, 14A, 14B) of the other column (2A, 2B).

Abstract: A hand controller system which provides the proper feel of the hand controller during aircraft operation. In both non-redundant and redundant embodiments, torque, and position measurements are made from the hand controller movements and processed to provide a feedback to a control motor in mechanical connection with the hand controller. The system includes self-monitoring of the motor, as well as the signals provided by the position, and torque sensors. Connections are provided from the hand controller system so that cross-coupling may be provided between the pilot's and copilot's hand controllers.

Abstract: The motor provided includes a unique laminated stator core and a means for changing the continuity of the magnetic circuit. The stator core structure is split into two part-cylindrical elements which define spaces between their opposing ends. Flux gates fill the spaces and provide magnetic circuit continuity. The flux gates are connected to solenoids which can displace the flux gates in and out of the spaces, thereby changing the continuity of the magnetic circuit of the motor. The motor rotor aligns its magnetic axis with the spaces, and will resist substantial imposed torque loads.

Abstract: A helicopter fly-by-wire flight control system (21) includes a model following control system architecture which provides the pilot with maneuvering feel. Maneuvering feel is generated by varying the gain of a rate model (52) in the control system to schedule a constant sidearm controller force per g of load factor currently on the helicopter main rotor. This gain variation works in conjunction with control logic which requires the pilot to apply a nose up pitch command during steep banked turns in order to offset a control system induced downward (pitch) bias of the aircraft nose. The control system induced bias is generated only during steep banked turns, thus allowing the pilot to fly shallow banked coordinated turns without having to apply a nose up pitch command.

Abstract: A hand-held joystick has spaced-apart groups of accelerometers and is free to move or rotate on multiple axes. Each group has, for example, three mutually perpendicular accelerometers, with each one producing an electrical signal responsive to linear acceleration along one axis. The signals are combined to derive a definition of acceleration of the joystick along the three mutually perpendicular linear axes and/or rotation about these axes. The combined signal is used by an electrical control system which provides input to a device being controlled. The signal to the device under control is responsive to the motion of the joystick.

Abstract: A helicopter programmable collective control system includes a motorized collective control stick 2 used in conjunction with a multi-axis control stick 30, with the multi-axis control stick 30 as the default collective controller. A collective control signal processor reads collective input the pilot is giving through one of the two control sticks 2, 30 and provides a motor drive signal 44 (which is dampened to prevent over control) to position the collective control stick 3. The collective control stick position is used to drive the rotor collective pitch actuators 16.

Abstract: A digital power controller for an aircraft having a computer controlled propulsion system is disclosed. A movable handle is in electronic communication with a computer of the propulsion system. The handle is positionable is a neutral location for maintaining the aircraft at a substantially constant velocity and for resetting power command functions. The handle is also positionable in locations forward the neutral location for commanding increasing acceleration rates of the aircraft. The handle is positionable in locations aft the neutral location for commanding increasing deceleration rates of the aircraft. A switch associated with the movable handle is engagable at the forward end aft positions for commanding velocity hold and for permitting reposition of the handle to the neutral location for resetting power command functions.

Abstract: A force control throttle suitable for military or civil aircraft which is stowable in a console when not required for use. Fore and aft forces applied to the throttle handle by the pilot are sensed by pressure transducers which generate, in response, signals for controlling engine thrust. Tactile information relating to engine thrust demand is fed back to the pilot by means of a drive motor which tilts the throttle handle as fore or aft pressure is exerted. The angle of tilt is a function of the degree of pressure exerted and the demanded engine thrust.

Abstract: System for controlling roll and yaw of an aircraft. The system includes a device capable of elaborating, from electric signals respectively representative of the position of a first voluntary actuation member, the rolling speed, the attitude, the yaw speed, the sideslip and of the position of a second voluntary actuation member, a single electrical order for roll control formed by a linear combination of the electric signals. There is a device capable of elaborating, from the same electric signals, an electrical order for yaw control formed by a linear combination of the electric signals. There is also a device which makes it possible to combine the electrical order for yaw control and a mechanical order coming directly from the second voluntary actuation member.

Abstract: Control surfaces in an aircraft are biparted, and one portion is acted upon by a mechanical signal transmission system in a redundant manner, the other one by an electrical signal transmission system normally operating in parallel; different construction versions and embodiments are described.

Abstract: Mechanical impedance determining apparatus for the control stick of an aircraft, an aircraft simulator, or non-aircraft apparatus such as a video game. Impedance determination is accomplished by a feedback path connected from output to input of the control stick with feedback transfer function coefficients determined by operator election in order to achieve positive, negative, or zero mechanical impedance at the control stick. Electrical network and computer realization of the feedback signal transfer function are contemplated using s plane pole descriptions of overall stick and feedback path characterizing transfer functions.