Abstract: An aircraft empennage, comprising a rear fuselage section, a trimmable horizontal tail plane comprising a first and a second lateral torsion box), each comprising a front spar and a rear spar, a front fitting and a rear fitting, an actuator acting on the front fitting for a rotation of the trimmable horizontal tail plane around a hinge axis passing through the rear fitting, the front fitting comprising a first and a second front fitting units being joined to the front spars of the lateral torsion boxes, and the actuator comprising first and a second actuator units, each acting on the first front fitting unit and the second front fitting unit respectively.

Abstract: Aircraft pitch control systems and methods are disclosed. An aircraft pitch control system (28) comprises a movable horizontal stabilizer (24) and an elevator (26) movably coupled to the horizontal stabilizer. The elevator is electronically geared to the horizontal stabilizer.

Abstract: An electromechanical device for actuating a movable flight control surface having a skin hinged to a structure of an aircraft about a pivot axis, the device including: at least two electric motors for causing the skin to pivot about the pivot axis; a power circuit and a control circuit for powering and controlling each of the motors; and a device for limiting opposing forces exerted by the two electric motors.

Abstract: A thrust reverser actuation system for a jet engine having a turbine engine surrounded by a nacelle to define an annular air flow path between the turbine engine and the nacelle, with a thrust reverser having a movable element to reverse the direction of at least a portion of the air flow along the air flow path, where the thrust reverser actuation system includes a hydraulic actuator configured to be operably coupled to the movable element, and a remote actuator that controls the inhibit function of the hydraulic system.

Abstract: The invention relates to a flap actuator, in particular to a servo/spoiler actuator, comprising at least one electrohydraulic servo valve and at least one actuator which are hydraulically connected to one another, wherein the actuator comprises at least one actuator housing in and/or at which the at least one channel is arranged by means of which one or more components, preferably likewise arranged in and/or at the actuator housing, are hydraulically connected and wherein no valve block is arranged between the actuator and the servo valve.

Abstract: An aircraft comprising rotors and/or propellers rotated by a shaft and including a power unit. The distributed electric power unit is formed by stacked electric power elements, each conferring, on the shaft, a fraction of the total power necessary; the distributed electric power unit is in direct engagement with the shaft, with no movement transmission mechanism inserted between the unit and the shaft; each electric power element is directly connected to the rotation shaft and comprises a fixed stator, a moving rotor and a mechanical or electromagnetic free wheel in direct engagement with the shaft, said moving rotor co-operating with the free wheel to be coupled to the rotation shaft during normal operation of the electric power unit and to be disconnected from the shaft in the event of the failure of the electric element.

Abstract: In some embodiments, a method of provided boosted actuation to an aircraft flight control device includes receiving an input from a pilot input device via a mechanical input member, providing mechanical energy to a driving member of a controlled-slippage actuator, and varying the strength of a magnetic field applied to a magnetorheological (MR) fluid disposed between the driving member and a driven member of the controlled-slippage actuator based on the relative positions of the mechanical input member and a mechanical output member that is in mechanical communication with the driven member and the aircraft flight control device.

Abstract: An electric flight control system onboard an aircraft provided with flight control surfaces and controllers for controlling these surfaces, that include at least one local electro-hydraulic generator to supply hydraulic servocontrols connected to flight control surfaces.

Abstract: A method to maneuver an aircraft in flight, in which the center of gravity of the aircraft is shifted by transferring fuel from at least one first fuel tank to at least one second fuel tank of the aircraft. A system implementing this method, the system including: at least one first fuel tank and at least one second fuel tank, a flight control unit capable of sending out a maneuver command upon being handled. A computer capable, as a function of this command, of determining a quantity of fuel to be transferred from the first tank to the second tank, at least one means of transfer connecting the first and second tanks and being controlled by the computer, to transfer the fuel from the first tank to the second tank.

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 hybrid power distribution system for an aircraft generates hydraulic power from one of a plurality of power sources based on which power source provides energy most efficiently. Power sources includes an electric power distribution bus that distributes electrical energy onboard the aircraft, a pneumatic distribution channel that distributes pneumatic energy onboard the aircraft, and mechanical power provided by one or more engines associated with the aircraft.

Abstract: Primary flight controls (10) for main and/or tail rotors (4) of helicopters with electro-mechanical interface between any of fly-by-wire and/or fly-by-light controls and hydraulic servo actuators (8) for control force amplification towards said main and/or tail rotor controls (4). For each of the main and/or tail rotor controls (4) there is provided but one of the hydraulic servo actuators (8), connected by one mechanical linkage (7) to one electro motor (5), said one hydraulic servo actuator (8) being of the type having the one mechanical linkage (7) connected to its input (29) and its output (30) and the one electro motor (5) being of the direct drive type, the position of said electro motor (5) having a reference to said one mechanical linkage (7) and the torque delivered by said electro motor (5) to the hydraulic servo actuator (8) being related to the power consumption of said electro motor (5).

Abstract: A system and method for a controlling an aircraft with flight control surfaces that are controlled both manually and by a computing device is disclosed. The present invention improves overall flight control operation by reducing the mechanical flight control surface components while providing sufficient back-up control capability in the event of either a mechanical or power-related failure. Through the present invention, natural feedback is provided to the operator from the mechanical flight control surface which operates independent of computer-aided flight control surfaces.

Abstract: A system for single wire secondary distribution comprising a spacecraft platform; a central bus interface unit coupled to the spacecraft platform; a payload unit coupled to the central bus interface unit; and a centralized power supply for powering the central bus interface unit and the payload unit; wherein the spacecraft platform provides a command to the central bus interface unit; wherein the central bus interface unit interrupts the power to the payload unit in a manner corresponding to the commands received by the central bus interface unit; wherein the payload unit decodes the interruption to the power and executes the command from the spacecraft platform.

Abstract: An aircraft having at least one in-flight attitude control system, in turn having at least one actuator; and a hydraulic circuit connected to the actuator and having at least one pump designed to deliver a first flow when the pressure of the hydraulic circuit is above a presettable threshold value. The pump is designed to deliver a second flow greater than the first flow, and the aircraft has a sensor for detecting a quantity associated with the pressure of the hydraulic circuit; and a programmable central control unit, which controls the pump to deliver the second flow when the quantity detected by the sensor corresponds to a pressure of the hydraulic circuit below the threshold value.

Abstract: An actuator control system has a dual concentric servo valve having a spool and at least one motor adapted to selectively displace the spool.

Abstract: The invention relates to a more electric flight control system onboard an aircraft provided with flight control surfaces and means of controlling these surfaces, that comprises at least one local electro-hydraulic generator (HPP1, HPP2) to supply hydraulic servocontrols (34) connected to flight control surfaces (31, 32, 33).

Abstract: A control surface drive system having a plurality of actuator assemblies are coupled to first and second supply lines and to a return line. The first and second supply lines are connected to a source of hydraulic fluid. At least one of the actuator assemblies has a hydraulic actuator movably connectable to an aircraft control surface. A flow control assembly is connected to the return line and to at least one of the first and second supply lines. A bypass line is in fluid communication with the first and second supply lines and positioned to recycle the hydraulic fluid from one of the first and second supply lines back into the other one of the first and second supply lines when the hydraulic actuator moves toward the first position. A computer controller operatively interconnects the plurality of actuator assemblies and the flow control assembly. It is emphasized that this abstract is provided to comply with the rules requiring an abstract.

Abstract: A backup system is provided that has a local electric motor and pump for some or all of the hydraulic actuators on an aircraft. A local backup hydraulic actuator has two power sources, hydraulic as primary and electrical as backup. During normal operation, the hydraulic actuator receives pressurized fluid from a hydraulic system and the fluid flow to the chambers is controlled by a servo valve. If the hydraulic system fails, the electronic controller detects the failure by observing the signal indicative of the pressure from the pressure sensor, and the controller powers the local hydraulic pump to provide high pressure hydraulic fluid to the hydraulic actuator via the servo valve.

Abstract: A flight control actuation system comprises a controller, electromechanical actuator and a pneumatic actuator. During normal operation, only the electromechanical actuator is needed to operate a flight control surface. When the electromechanical actuator load level exceeds 40 amps positive, the controller activates the pneumatic actuator to offset electromechanical actuator loads to assist the manipulation of flight control surfaces. The assistance from the pneumatic load assist actuator enables the use of an electromechanical actuator that is smaller in size and mass, requires less power, needs less cooling processes, achieves high output forces and adapts to electrical current variations. The flight control actuation system is adapted for aircraft, spacecraft, missiles, and other flight vehicles, especially flight vehicles that are large in size and travel at high velocities.

Abstract: A flight control actuation system comprises a controller, electromechanical actuator and a pneumatic actuator. During normal operation, only the electromechanical actuator is needed to operate a flight control surface. When the electromechanical actuator load level exceeds 40 amps positive, the controller activates the pneumatic actuator to offset electromechanical actuator loads to assist the manipulation of flight control surfaces. The assistance from the pneumatic load assist actuator enables the use of an electromechanical actuator that is smaller in size and mass, requires less power, needs less cooling processes, achieves high output forces and adapts to electrical current variations. The flight control actuation system is adapted for aircraft, spacecraft, missiles, and other flight vehicles, especially flight vehicles that are large in size and travel at high velocities.

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: A system is provided for increasing autopilot control authority in a vehicle, such as aircraft including mechanical control systems. The system includes an autopilot system, a control element and an autopilot supplement assembly. The autopilot system can automatically control the vehicle by applying a variable autopilot force to control surfaces. The control element can control the vehicle by applying a variable control force to control surfaces, where the control force acts counter to the autopilot force. The autopilot supplement assembly can measure the control force and, in turn, determine a variable supplemental force. Thereafter, the autopilot supplement assembly can apply the supplemental force to the control surfaces such that the sum of the supplemental force and the autopilot force is greater than the control force. Thus, the system may provide the benefits of systems that require full control authority to vehicles such as aircraft that include mechanical control systems.

Abstract: The control device (1) comprises at least one first and one second control system (4, 5) which comprises control elements (A1 to A8, B1 to B6) actuable by an operator, so as to control the monitoring system (2). These first and second control systems (4, 5) are in interaction so that an actuation of a control element of one of the control systems (4, 5) automatically brings about an action on a control element of the other control system (5, 4) which is intended to control the same function (F1 to F6) in such a way as to bring it into a control state representative of the controlled state of this function (F1 to F6).

Abstract: The invention relates to a servo-control device comprising at least one servo-control loop which receives a servo-control signal as input and whose output acts on an element to be servo-controlled, said device comprising means for determining a parameter characterizing the real force generated by the servo-control, the device also including a loop having means for determining an error characteristic of the difference between said parameter characterizing the real force and a parameter characterizing an acceptable theoretical force which is a function of the servo-control signal, and also means for correcting the processing of the servo-control loop as a function of said error. The means for determining the error make use of a reference model which is advantageously a transfer function relating force to a flow rate control signal.

Abstract: A bias correction system for use in a neutrally-stable aircraft having a control column and a pitch control system is provided. The position of the control column is generally represented by a control column position signal. The bias correction system is for removing control column bias when the control column is in a neutral position. The bias correction system includes a first combining unit for combining the control column position signal and a correction signal, and a switch. The switch includes activated and deactivated states. The switch is set to the deactivated state when the control column is physically displaced from its neutral position. The deactivated state allows the correction signal to remain at its last value, the activated state allows the correction signal to equal approximately the control column position signal.

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: 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: Herein disclosed is a flight control system comprising a control member, a control surface controllable to assume different angle positions within a control angle range, a transmission linkage provided between the control member and the control surface to transmit the control force from the control member to the control surface, and a power assist actuator associated with the control member and the transmission linkage to control the control surface.

Abstract: A fault tolerant actuation system (12) for flight control systems is provided. The fault tolerant actuation system (12) includes a plurality of primary flight computers (14a, 14b, and 14c) with corresponding power control units (24a, 24b, and 24c). Each power control unit includes a remote electronic unit (18a, 18b, and 18c), an electro-hydraulic servo valve (26a, 26b, and 26c), and an actuator (28a, 28b, and 28c). The actuators are linked to a flight control surface (30) to control its position. The electro-hydraulic servo valves (26a, 26b, and 26c) and the actuators (28a, 28b, and 28c) include sensors that monitor their operation. Each RE (18a, 18b, and 18c) generates a control current (i.sub.1, i.sub.2, and i.sub.3) based upon commands of the corresponding primary flight computer as well as feedback data transmitted from the sensors of the corresponding electro-hydraulic servo valve and actuator only. The feedback data is transmitted along separate servo loops having separate compensations (66a, 70a).

Abstract: A device for controlling a rudder of an aircraft has at least two servo-systems each including at least one electric control input. At least one of the servo-systems is a mixed servo-system that also has a mechanical control input. When an engine fault occurs, an electric control system causes at least two of the servo-systems to simultaneously operate the rudder. When the electric control system is not operational due to an electric fault, the mixed servo-system operates the rudder based on control signals from the mechanical control input.

Abstract: System for control of an aerodynamic surface (1) of an aircraft, comprising an intentional actuation member (5) and a hydraulic jack (2) intended to deflect said aerodynamic surface (1).According to the invention, the system comprises means (14) for detecting the difference in pressures prevailing in said jack (2), on either side of the piston (2C) of the latter, and calculation means (11) comprising:a computer delivering values of a coefficient which is a decreasing function of said difference in pressures; anda multiplier receiving an electrical signal (dp) representing a demanded deflection and said coefficient and addressing a first deflection command (C2) formed by the product of said signal and said coefficient to said servovalve.

Abstract: An oscillatory servo-valve fault monitor is provided for identifying faults attributable to servo-control system components. In particular, both low frequency and high frequency (oscillatory) fault responses are identified. Such oscillatory fault responses are distinguished from normal position control adjustments and transient responses. As the reaction to a fault condition is to shut down the faulty actuator and corresponding servo-control, such capability avoids inadvertent shut-downs. Independent detection paths are provided for sensing low frequency responses and high frequency responses. Each path includes a respective window comparator and up-down counter. For a given path, the window comparator defines two thresholds--in effect, a window. If the monitored signal amplitude is outside the window, then a counter or count signal is incremented. If the signal amplitude is within the window, then the counter or count signal is decremented.

Abstract: A system for cooling individual integrated actuator packages (IAPs) positioned throughout the aircraft in a power-by-wire actuation system is disclosed. A low-pressure hydraulic fluid circulates throughout the aircraft to control the temperature of the individual IAPs and to maintain and replenish the hydraulic fluid within the individual IAPs. A thermo-control loop containing low-pressure hydraulic fluid includes a pump, low-pressure hydraulic lines, a heat exchanger, a reservoir and a filter. The entire system provides heat transfer for a high- or low-temperature environment, as may be present on an aircraft. The individual IAPs may be coupled together in series or, alternatively, in parallel, depending upon the design constraints of the system. Compared with distributed hydraulic systems, energy consumption, weight, and cost are significantly reduced by using a low hydraulic pressure for the circulating fluid. Survivability is significantly increased over that provided by a distributed hydraulic system.

Abstract: A control system and method for controlling a flight control member employs a hydraulic control valve which is responsive to input from the pilot in a manner to start and stop a hydraulic motor which in turn starts and stops movement of the control member so that the control member assumes the desired flight control position.

Abstract: An attitude control system for controlling a plurality of attitude control devices, which cooperate to establish the attitude of a vehicle includes redundant control units for outputting attitude control signals to each of the attitude control devices. Additionally, a plurality of discreet power servo units are provided which are located in proximity to and operatively connected to the plurality of attitude control devices. The redundant control units and the plurality of discreet power servo units are connected electrically and/or optically. Each of the discreet power servo units includes an adder for adding an attitude control signal and a positional feed back signal, a rotational drive source for providing a rotational motion in response to the output from the adder, a fluid pressure circuit for hydraulically driving an attitude control device coupled thereto, and a sensor for sensing the position of the attitude control device coupled thereto and for providing to the adder the positive feed back signal.

Abstract: A control system for a working machine which has a boom equipped with a working implement at its forward end and pivoted to a vehicle body movably about a vertical axis, and control valve for controlling the movement of the boom. The system comprises a variable resistor for setting a target position where the boom is to be stopped to produce a setting signal, a sensor for detecting the moved position of the boom to produce a detection signal, differential device for determining the difference between the setting signal and the detection signal to produce a difference signal, a detector for detecting the direction of movement of the boom from the magnitude of the difference signal, a pulse-width modulator for subjecting the difference signal to pulse-width modulation to produce a pulse signal, and a driver for driving the control valve by the pulse signal in the detected direction in proportion to the difference signal.

Abstract: A dual redundant servomechanism for moving aircraft control surfaces is disclosed. The servomechanism is of the type whose input commands, from the pilot of the aircraft, are transmitted electrically. Force fight, which is associated with such dual servomechanisms when they are connected to a common aircraft control surface, is minimized. This is accomplished by providing the control system for each servomechanism with input signals which are electrically summed. Each control system includes electrical transducers which provide a signal indicative of actuator position and the pressure associated with the hydraulic motor used in each servomechanism.

Abstract: An elevator control system especially for an aircraft such as an airplane, s equipped with an electrically controllable elevator flap drive system for the elevator flap on each side of the aircraft. Additionally, the system is equipped with a mechanically controlled auxiliary drive system which is responsive to a hand wheel operated by the pilot. A single hand wheel has mechanical connecting links to the elevator flaps on both sides of the aircraft. Monitoring features enable the pilot to test the auxiliary mechanical drive system without actually switching on that system. The auxiliary mechanically operated system is functional even if all other power systems failed, since a ram air turbine is provided for generating the hydraulic pressure for driving the auxiliary system.

Abstract: A vertical rudder control and drive system for an aircraft, such as an aiane, is equipped with electrically controllable vertical rudder drive systems on each side of the aircraft. Additionally, the control system is equipped with a mechanically controllable auxiliary drive system for the operation of the vertical rudder in response to foot pedals operated by the pilot when there should be a failure in the electrically controlled drive systems. A mechanical control signal transmitting link is provided between the foot pedals and the hydro-mechanical drive for the vertical rudder. Monitoring features enable the pilot to test the mechanical drive system without actually using that system in flight. Preflight tests may be performed.

Abstract: In a control system having a manipulatable unit, such as an aircraft stick, which is biased toward a neutral position to provide control "feel", an input which at each instant signifies the position of displacement of the stick from its neutral position is processed to provide an output signifying the magnitude of the biasing force to be applied to the stick. In a range of positions spaced to one side of its neutral position (or in such a range at each side of neutral) the output provides for a steep biasing force gradient as the stick is moved through a predetermined distance in either direction from a turning point to which the stick had been brought by movement in the opposite direction. For all other movements of the stick within that range the biasing force has a low gradient. The higher gradient of bias for small trimming movements provides better feel for trimming without requiring stick bias to be unduly stiff during coarser stick displacements.

Abstract: A system which provides feel-force to the control stick of an aircraft by means of hydraulic pressure is provided with non-nulling, proportional, direct feedback loop, and a limited integral feedback loop in the drive of the pressure control servo valve that commands the pressure-generating hydraulic force actuator. Static nulls are compensated in the integral path; integration is corrected when the static null and integral output exceed a limiting value. Embodiments include software and/or hardware portions.