Abstract: Systems and methods for use in navigating aircraft are provided. The systems can use Geometry Redundant Almost Fixed Solutions (GRAFS) or Geometry Extra Redundant Almost Fixed Solutions (GERAFS) to compute high confidence error bounds for a heading angle estimate or pitch angle derived using signals received on at least two antennas.

Abstract: A Stability and Control Augmentation System (“SCAS”) module comprising one or more SCAS actuators, the or each SCAS actuator comprising a mechanical component that translates rotational motion to linear motion along a first axis of said SCAS; one or more electric motors for driving linear movement of the mechanical component in response to a command signal; and one or more angular transducers to detect the position of the SCAS actuator along the first axis.

Abstract: An aircraft comprises an aircraft structure, a flight control surface attached to the aircraft structure, and a local hydraulic power pack disposed proximal to the flight control surface. The local hydraulic power pack is configured to provide pressurized hydraulic fluid for actuating the flight control surface. The local hydraulic power pack comprises a reservoir for the hydraulic fluid and two hydraulic pumps for pressurizing the hydraulic fluid.

Abstract: A high-lift actuation system for differentially actuating a plurality of high-lift surfaces of an aircraft is disclosed. An exemplary high-lift actuation system includes a centralized drive device for centralized actuation control of an inboard high-lift surface of a first wing and a second wing, respectively, and at least two independent drive devices for individual actuation control of an outboard high-lift surface of the first wing and the second wing, respectively. The centralized drive device may include a central power drive unit (PDU) operably coupled to a common central driveline for driving the inboard high-lift surfaces, and the common central driveline may be separate and spaced apart from a respective driveline of the independent drive devices. The common central driveline may mechanically synchronize movement of the inboard high-lift surfaces, and a controller may electronically coordinate synchronized movement and controlled differential movement of the plurality of high-lift surfaces.

Abstract: A method for evaluating load performance of a rotor/stator test coupon tested within a sealable test chamber containing test fluid. The test coupon preferably includes at least a partial length of a PDM stage. Alternatively, the test coupon may include a splined rotor/stator. The method includes rotating either the rotor section or the stator section in order to actuate corresponding rotation of the other of the rotor section and the stator section. Non-linear torque in the form of an acceleration torque and/or a braking torque may be applied to either the rotor section or the stator section. Downhole stall conditions may be simulated by selectively engaging and disengaging a second motor and flywheel to vary rotational torque applied to the test coupon. Load performance of the test coupon may be evaluated over time in such simulated stall conditions.

Abstract: A check device is provided for a flight actuator primary load path failure detection device of the type that disconnects a position sensor from the primary load path in the event of a primary load path failure. The check device comprises: a mechanical linkage for simulating disconnection of the position sensor by permitting relative movement of at least first and second mechanical parts of the actuator that are unable to move relative to one another in normal use without failure of the primary load path, wherein these first and second mechanical parts, include a first mechanical part with movement detected by the position sensor of the primary load path failure detection device.

Abstract: An improved actuator arrangement for operating a driven element with different actuators. The actuator arrangement comprises a first actuator that drives a driven element in a first direction via a first output member and a second actuator that drives the driven element in a second direction opposite to the first direction via a second output member. The first and second actuators are configured to work in a counter-acting manner to move the driven element in the first and second directions. An additional connecting and supporting device is connected between the first and second output members to support the output members to bear transversal forces transversal to the moving direction and/or to bear pre-load force for pre-loading the actuators.

Abstract: A flexible response secured mechanical balancing device (D) for multiple hydraulic control actuators (1-2) with a common output member (5a-b). The balancing device (D) provides flexible/continuous fluid-pressure force fight compensation. Balance valves (17-20) are commanded to balance fluid-pressure from a first fluid-pressure network (52a) of a second actuator (2). Each fluid-pressure commanded balance valve (17-20) is arranged to allow relief of differential pressure, directly into a dedicated first (52a) or second fluid-pressure network (52b). A fluid-pressure controlled locking device (21-22) is allowing inhibiting the corresponding pairs of balance valves (17-20) e.g. in case of loss of fluid-pressure. The invention typically applies to vehicles (A) including: aircrafts, rotorcrafts, drones.

Abstract: The invention provides a motor-driven movement system for moving a movable element, the system comprising at least two actuators, each provided with means connecting it to the movable element and each dimensioned to be capable, on its own, of driving the movable element, a central control unit being connected to the two actuators in order to be capable of sending a position setpoint (Pos1, Pos2) to one or other of the actuators. According to the invention, the system further comprises control means for simultaneously controlling both actuators in terms of force in response to the position setpoint sent to one of the actuators. The invention also provides a method of driving such a system and a method of testing such a system.

Abstract: A control force transmission arrangement for an aircraft is provided. A transmission device has a first and a second force transmission point, and a coupling unit disposed between them. The coupling unit has a first and a second side element and a connecting element. The first and the second side elements are each connected to the first force transmission point and to the connecting element. At least the first side element has an element for altering the length of the side element. The second force transmission point is provided on the connecting element and is adjustable at least between a first and a second position.

Abstract: A system for computing flight commands, which comprises simplex computers coupled to form virtual command/monitor pairs. These virtual pairs send the commands computed to actuators whose controller implements an election of the pair whose command will be taken into account by the actuator.

Abstract: A control force transmission arrangement for an aircraft is provided. A transmission device has a first and a second force transmission point, and a coupling unit disposed between them. The coupling unit has a first and a second side element and a connecting element. The first and the second side elements are each connected to the first force transmission point and to the connecting element. At least the first side element has an element for altering the length of the side element. The second force transmission point is provided on the connecting element and is adjustable at least between a first and a second position.

Abstract: A device (13) for detecting the breakage of a primary path in a flight control actuator, said actuator having a primary path (1) comprising a rotary hollow screw (2), a secondary path (10) comprising a safety rod (3) that reacts the load passing through the screw (2), said device (13) being characterized in that it comprises a position sensor (15), connected to the screw (2), to measure information representative of the angular position thereof, and a disconnection system (17) able to disconnect the screw (2) position sensor (15) in the event of relative movement of the rod (3) with respect to the screw (2) if there is a break in the primary path (1).

Abstract: Systems and methods are provided for arbitrating sensor and actuator signals in various devices. One system includes input/output (I/O) circuitry, redundant computation circuits coupled to the I/O circuitry, and an arbitration circuit coupled between the I/O circuitry and the redundant computation circuits. The I/O circuitry is configured to be coupled to multiple non-redundant systems, and the redundant computation circuits are configured to be coupled to one of multiple system buses. One such device is an aircraft including multiple non-redundant systems and a plurality of system buses that are configured to transmit redundant messages to the non-redundant systems.

Abstract: According to one embodiment, a rotorcraft pitch link includes a housing and a piston assembly. The housing comprises a first chamber and a second chamber. A first housing opening allows a first fluid to flow into the first chamber, and a second housing allows a second fluid to flow into the second chamber. The piston assembly is at least partially disposed within the housing and comprises a piston head and a piston rod coupled to the piston head. The piston head separates the first chamber from the second chamber.

Abstract: A fault-tolerant actuating system with at least one flap, adjustable on a respective wing of an aircraft, and with a control and monitoring device, includes: drive devices that are functionally connected to the control and monitoring device, with one of these drive devices in each case being associated with a flap, where each flap in each case is associated with a drive device, in each case including: two drive motors, two brake mechanisms, where each drive motor is associated with a brake mechanism for stopping rotation of the output of the respective drive motor, a differential that couples the outputs of the aforesaid to the aforesaid in a summing manner, an output shaft for coupling the output of the differential to drive connections, and a differential lock that is functionally connected to the control and monitoring device, which differential lock is coupled to the control and monitoring device in this manner, where each of the brake mechanisms as well as the differential lock can be operated by way of

Abstract: A hydraulic system for aircraft wherein at least one circuit includes two hydraulic pumps, at least one of which is driven by an engine that may suffer uncontained failure that can damage hydraulic lines close to the pump, is equipped with pressure sensors for the hydraulic fluid in the lines close to the pump and a sensor for the hydraulic fluid level in a hydraulic tank of the circuit supplied by the pump. A control system for a cut-out valve installed on a suction line wherein the fluid arrives at the pump includes logic that determines the occurrence of an uncontained engine failure requiring the isolation of the pumps from line elements that may have been damaged from measurements of the fluid pressures in the lines, from measurement of the level of fluid in the tank and from information supplied late by a uncontained engine failure detection system to command the closure of the cut-out valve.

Abstract: An electrical system for driving flight control actuators, electrical accumulator, and methods for managing electrical power for an aircraft electrical system and methods of retrofitting an aircraft electrical system with an electrical accumulator are provided. The electrical system can include a bidirectional accumulator positioned in electrical communication with the plurality of flight control actuators to absorb excess electrical power regenerated by the plurality of flight control actuators and to provide supplemental power to the plurality of flight control actuators to thereby satisfy transient power requirements.

Abstract: The disclosure relates to a device for testing means of detecting the loading of the secondary path of a flight control actuator of the type including a primary path and a secondary path able to take up the effort of the primary path in the event of a breakdown of the latter, wherein a tool is able to press against at least one zone of at least one element of the secondary path, as well as a lever arm which by tilting allows the tool to press against the zone so that it exerts a traction and/or compression force on the secondary path, the lever arm being mounted pivotingly on a support able to be fastened on a structure that is mechanically independent of the secondary path.

Abstract: A method of operating a control system for a machine is disclosed. The method may include receiving machine control settings for a first operator station from an operator and receiving an input reflecting a desire to switch control to a second operator station. The method may also include automatically adjusting control settings of the second operator station to match the control settings of the first station. The method may further include deactivating the first operator station once the control settings of the second operator station substantially match those of the first operator station.

Abstract: A system and method are provided for controlling aircraft flight control surfaces. The system may include at least three pilot sensor channels, each pilot sensor channel including a set of pilot sensor data. The system may also include at least three aircraft sensor channels, each aircraft sensor channel including a set of aircraft sensor data. The system may further include an actuator control component configured to synchronously receive and vote on the pilot sensor data and the aircraft sensor data, such that a voted output of the at least three pilot sensor channels is transmitted to a flight control computer and augmented before being transmitted to remote electronics units. The voted output of the at least three pilot sensor channels providing for the control of the aircraft surfaces coupled to the remote electronics units.

Abstract: For emergency piloting of a manual flight control system (7) of an aircraft (1), in the event of flexible connection means (26) breaking, an upstream control device (8) takes authority over the electronic control unit (29) of series actuator (21). An output pivoting quadrant (25) is subjected by a centering rod (22) to a blocking and centering action. In the event of the system breaking, the output pivoting quadrant (25) forms a bearing point for the series actuator (21), so that an emergency control actuate a airfoil surface (11).

Abstract: A floating piston actuator for use with multiple hydraulic systems includes a housing and a piston assembly. The piston assembly includes an elongate shaft or ram, at least one secured piston coupled to the ram, and a set of moveable pistons moveably disposed within the housing chamber on the ram such that the set of moveable pistons are translatable relative to a longitudinal axis of the ram. The set of moveable pistons form at least three hydraulically separate actuator cylinders where each of the three hydraulically separate actuator cylinders are disposed in fluid communication with a corresponding pressurized fluid source. The floating piston actuator allows hydraulic redundancy while maintaining a similar length compared to existing multi-system actuators. The floating pistons can be arranged in multiple configurations to allow operation with a hydraulic system failure without force fight or binding and with a higher output force than conventional multi-system actuators.

Abstract: A system and method of mitigating a force fight between hydraulically-operated actuators that are coupled to a single flight control surface is provided. The differential fluid pressure across each hydraulically-operated actuator is sensed. The position of a user interface is sensed using a plurality of user interface position sensors. Flight control surface position is sensed using one or more position sensors. The sensed differential pressures, the sensed user interface positions, and the sensed flight control surface position are used to generate a plurality of substantially equal actuator commands.

Abstract: The present invention relates to an on-board aeronautic system intended to be dynamically reconfigured, especially an on-board information system, and to an associated method as well as to an aircraft comprising such a system. In particular, the system comprises a plurality of heterogeneous equipment items, at least part of the said equipment items being reconfigurable, and comprises a reconfiguration management center set up to receive state messages from the said plurality of equipment items and to emit reconfiguration messages destined for the said reconfigurable equipment items as a function of at least the said received messages, the said state messages being emitted according to the same format by monitoring means encompassed in each of the said equipment items, the reconfigurable equipment items encompassing reconfiguration means capable of processing at least one of the said reconfiguration messages in order to reconfigure the said associated equipment item.

Abstract: The invention relates to a flight control system for an aircraft comprising control surfaces and actuators associated with the control surfaces to control flight functions, including roll, yaw, pitching, and aerodynamic braking of the aircraft. The actuators associated with the control surfaces that control at least one of the flight control functions are electromechanical actuators. A part of the control surfaces associated with the electromechanical actuators is divided, where each of the divided control surfaces is comprised of at least two independent surfaces.

Abstract: A hydraulic pump is installed inside a wing at which a control surface is provided, and supplies pressure oil to an actuator when a loss or degradation occurs in the function of aircraft central hydraulic power sources. An electric motor is installed inside the wing and drives the backup hydraulic pump. A driver is installed inside the wing and drives the electric motor. A cooling device is installed inside the wing and simultaneously cools the backup hydraulic pump, the electric motor, and the driver.

Abstract: A flight control system includes a dual stage actuator for moving a control surface. Each stage includes several control valves that are controlled independently to provide a desired redundancy. A flight controller generates a position command that is indicative of the position desired for the control surface. A first communication link is provided between several flight controllers to share information. Each of the flight controllers forwards the position command to actuator remote processing unit. The actuator remote processing unit receives position commands and generates a command signal that controls movement of the actuator using the control valves. Each of the actuator remote processing units is linked through a second communication link. Feedback and balancing of the different control valves is provided by the visibility accorded each actuator remote processing unit by the second communication link.

Abstract: A first actuator drives a control surface by being operated by supply of pressure oil from a first aircraft central hydraulic power source including a first aircraft central hydraulic pump. A second actuator drives the control surface by being operated by supply of pressure oil from a second aircraft central hydraulic power source including a second aircraft central hydraulic pump. A backup hydraulic pump is installed inside a wing of the aircraft and is provided so as to be able to supply pressure oil to the first actuator when a loss or degradation in a function of at least one of the first aircraft central hydraulic power source and the second aircraft central hydraulic power source occurs. A maximum discharge pressure of the backup hydraulic pump is set to be greater than maximum discharge pressures of the first aircraft central hydraulic pump and the second aircraft central hydraulic pump.

Abstract: A pump unit installed inside a wing, and a panel body constituting part of a wing structure portion forming a surface structure of the wing are provided. The pump unit includes a backup hydraulic pump that can supply pressure oil to an actuator when a loss or reduction occurs in the function of an aircraft central hydraulic power source and an electric motor that drives this pump. Except for at least the panel body, the wing structure portion is formed from fiber reinforced plastics. The panel body is formed of a metallic material.

Abstract: The actuator locking mechanism is a bi-directional secondary load path lock that operates under both compression and tension loads via a radial locking wedge and axial slide lock system. This locking mechanism is designed to effectively protect against structural disconnect after a primary load path failure occurs and to minimize the axial play in the engaged secondary load path. To avoid the dormancy of such a failure, the activated mechanism jams/stalls the actuator drive thereby alerting the aircrew and the maintenance crew of the fault condition and allowing safe continued flight and landing. This locking mechanism is impervious to relative deflection between the primary and secondary load path such that under a non-failed condition, the mechanism will not jam. A jam only occurs when axial displacement (tension or compression direction) is greater than a predetermined relative deflection. This axial displacement is eliminated by the slide locking system minimizing the axial play.

Abstract: A device (13) for detecting the breakage of a primary path in a flight control actuator, said actuator having a primary path (1) comprising a rotary hollow screw (2), a secondary path (10) comprising a safety rod (3) that reacts the load passing through the screw (2), said device (13) being characterized in that it comprises a position sensor (15), connected to the screw (2), to measure information representative of the angular position thereof, and a disconnection system (17) able to disconnect the screw (2) position sensor (15) in the event of relative movement of the rod (3) with respect to the screw (2) if there is a break in the primary path (1).

Abstract: An actuator assembly having a primary load path for tightly coupling an actuated surface to a reference structure and a secondary load path having a backlash portion for coupling the actuated surface to the reference structure with backlash, wherein the secondary load path is unloaded during an operative state of the primary load path and loaded during a failure state of the primary load path. The actuator assembly includes at least one sensor configured to sense the failure state of the primary load path when a relative displacement between a portion of the primary load path and a portion of the secondary load path exceeds a predetermined value or is within a predetermined range of values.

Abstract: A regenerative integrated actuation system and method are provided to efficiently supply power to a plurality of actuators and to make efficient use of power regenerated by the actuators. The system may include a plurality of actuators in communication with a power distribution system. Each actuator may include an energy storage element. The system may also include a controller for determining if a respective actuator is in a consumption mode or a regenerative mode. If in the consumption mode, the controller may determine whether the respective actuator should draw power from an energy storage element or from the power distribution system. Conversely, if in the regenerative mode, the controller may determine whether any excess power from the respective actuator should be provided to one of the actuators, such as to another actuator which is in the consumption mode or to the energy storage element associated with the respective actuator.

Abstract: A drive arrangement comprising first and second actuators (24) arranged to drive respective output members (22) for movement, first and second synchronization arrangements (34) associated with respective ones of the first and second actuators (24), and a synchronization shaft (36) interconnecting the synchronization arrangements (34).

Abstract: An aircraft control surface failsafe hinge configuration is disclosed. The arrangement comprises a drop link configuration which is adapted to attach a control surface to an aircraft structure where the drop link additionally incorporates a failsafe hinge means which provides a backup hinge connection between the control surface and the aircraft structure to which it is attached. The improved failsafe hinge construction may be applied to any aircraft control surface such as an aileron, elevator, vertical tail plane and the like.

Abstract: A hybrid actuator assembly includes an actuator housing, an electromechanical actuator, and a hydraulic actuator. The electromechanical actuator is disposed at least partially within the actuator housing and is adapted to be controllably energized. The electromechanical actuator is configured, upon being controllably energized, to supply a first drive force. The hydraulic actuator is disposed at least partially within the actuator housing and is adapted to receive pressurized hydraulic fluid. The hydraulic actuator is configured, upon receipt of pressurized hydraulic fluid, to supply a second drive force.

Abstract: A hydraulic pump is installed inside a wing at which a control surface is provided, and supplies pressure oil to an actuator when a loss or degradation occurs in the function of aircraft central hydraulic power sources. An electric motor is installed inside the wing and drives the backup hydraulic pump. A driver is installed inside the wing and drives the electric motor. A cooling device is installed inside the wing and simultaneously cools the backup hydraulic pump, the electric motor, and the driver.

Abstract: An actuator arrangement comprises a first screw shaft, a first nut cooperating with the first screw shaft such that rotation of the first screw shaft causes the first nut to translate thereon, a second screw shaft, a second nut cooperating with the second screw shaft such that rotation of the second screw shaft causes the second nut to translate thereon, and first and second linkages connecting the first and second nuts, respectively, to an output member, rotation of the first and second screw shafts to cause translation of the first and second nuts towards one another causing extending movement of the output member and rotation of the first and second screw shafts to cause translation of the first and second nuts away from one another causing retracting movement of the output member.

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: 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: Herein disclosed is an actuator control apparatus, comprising primary control means for producing a control signal used to control an operation of an actuator in accordance with an instruction signal indicative of an operation of the actuator, secondary control means for producing a control signal in accordance with the instruction signal; and switching means for selectively switching one of the primary control means and the secondary control means to the other, and in which the secondary control means is less in the number of functions than the primary control means, and each of components constituting the secondary control means is easier in fault verification than a component constituting the primary control means the most difficult in fault verification among all of components constituting the primary control means.

Abstract: The assembly comprises a first and a second nut intended to cooperate with a screw in particular provided in an adjustable tailplane actuator for an aircraft, and a mechanical connection between the first and the second nut comprising means for fusible driving of the second nut by the first nut, in which the mechanical connection also comprises adjustable spacing means forming a wedge between the first and second nuts.

Abstract: A hydraulic system for aircraft operating devices provides a secondary pressure source for direction and deceleration control in case of loss of pressure from a primary pressure source. The system includes multiple pressure accumulators that include a primary accumulator, which supplies pressurized fluid to a first actuator, and first and second secondary accumulators, which supply pressurized fluid to second and third actuators, respectively. If the primary accumulator is disabled, the first secondary accumulator will supply pressurized fluid to the first actuator. If the first secondary accumulator is disabled, the second secondary accumulator will supply pressurized fluid to the second actuator. By providing multiple, substantially independent accumulators, direction and deceleration control is maintained even in the event of loss of the primary accumulator.

Abstract: The invention relates to a control surface (3) for an aerodynamic lifting surface (2) of aircraft comprising a main closing rib (9) located at one end of the control surface (3) to which a main torsion bar (8) is joined, the mentioned torsion bar (8) being joined at its other end to a lever system (14) on which at least one actuator (15) acts, such that it is possible to act on the rotation of the mentioned control surface (3) during the flight of the aircraft. The invention also relates to an actuation method for actuating such a control surface (3).

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: The use of at least one attachment screw (12) that includes means (14, 15, 17a, 17b, 20) for detecting the secondary path loading of a flight-control actuator, or associated with a bush that includes such means, for the attachment of the top coupling of the secondary path of a flight-control actuator.

Abstract: A fault-tolerant actuating system with at least one flap, adjustable on a respective wing of an aircraft, and with a control and monitoring device, includes: drive devices that are functionally connected to the control and monitoring device, with one of these drive devices in each case being associated with a flap, where each flap in each case is associated with a drive device, in each case including: two drive motors, two brake mechanisms, where each drive motor is associated with a brake mechanism for stopping rotation of the output of the respective drive motor, a differential that couples the outputs of the aforesaid to the aforesaid in a summing manner, an output shaft for coupling the output of the differential to drive connections, and a differential lock that is functionally connected to the control and monitoring device, which differential lock is coupled to the control and monitoring device in this manner, where each of the brake mechanisms as well as the differential lock can be operated by way of

Abstract: Actuator systems and associated methods are disclosed herein. One aspect of the disclosure is directed toward an actuator system that includes a first structure and a second structure movable relative to the first structure. The system further includes an actuator apparatus and an actuator device coupled in series between the first structure and the second structure. The system can still further include a controller operably coupled to the actuator apparatus and the actuator device. The controller can be programmed with instructions to automatically actuate the actuator apparatus and the actuator device so that a position of the first structure relative to the second structure after the actuator apparatus and the actuator device have been actuated is at least approximately the same as a position of the first structure relative to the second structure before the actuator apparatus and the actuator device have been actuated.

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.