Abstract: The subject matter of this specification can be embodied in, among other things, a linear-to-rotary apparatus that includes a linear actuator having an actuator housing including a piston chamber, a piston shaft disposed in the piston chamber, and a rotor apparatus. The rotor apparatus includes a rotary joint defining a rotational axis, a rotor arm extending radially from the rotary joint and configured to at least partially pivot about the rotary joint, and a torque linkage pivotably connected to the rotor arm. The torque linkage is also attached to an end of the piston shaft of the piston at a pivot connection joint, where the pivot connection joint defines a pivot axis that is substantially perpendicular to the translation axis of the piston shaft.

Abstract: A thermal management system includes a first hydraulic system for circulating a first hydraulic fluid at a first temperature and a second hydraulic system for circulating a second hydraulic fluid at a second temperature that is higher than the first temperature. The thermal management system also includes a sealed heat transfer device coupled between the first hydraulic system and the second hydraulic system. The sealed heat transfer device is not in flow communication with either of the first hydraulic system and the second hydraulic system. The sealed heat transfer device is configured to transfer heat from the second hydraulic fluid to the first hydraulic fluid.

Abstract: A rotary actuator includes a housing defining an arcuate chamber including a cavity, a fluid port in fluid communication with the cavity, and an open end. A rotor assembly includes an output shaft and a rotor arm extending outward. An arcuate-shaped piston is disposed in said housing for reciprocal movement in the arcuate chamber through the open end, wherein a seal, the cavity, and the piston define a pressure chamber, and a portion of the piston contacts the first rotor arm. A central actuation assembly includes a central mounting point formed in an external surface of the output shaft, said central mounting point proximal to the longitudinal midpoint of the shaft, and an actuation arm removably attached at a proximal end to the central mounting point, said actuation arm adapted at a distal end for attachment to an external mounting feature of a member to be actuated.

Abstract: A method of operating actuators simultaneously for moving at least two aircraft movable aerodynamic surfaces, the method comprising the steps of: controlling the actuators to move the aerodynamic surfaces towards a predetermined position; during the movement, detecting a slowest actuator; and adapting the control of the actuators to match the actions of the slowest actuator. A drive device for aerodynamic surfaces and an aircraft including such a device.

Abstract: The invention relates to an aircraft having at least one high lift system which is arranged at the wing of the aircraft and which comprises a drive for converting electrical or hydraulic energy into a speed-controlled rotational movement, wherein the aircraft furthermore has at least one control unit which controls the high lift system, wherein the drive comprises in accordance with the invention a main drive and an alternative drive, wherein the alternative drive is fed by a decentralized energy source.

Abstract: According to one embodiment, a radial fluid device includes a cylinder block having a plurality of radially extending cylinders, a plurality of pistons, and a cam disposed about the plurality of radially extending cylinders. A first linear control is coupled to the cam and operable to reposition the cam along a first axis. A second linear control is coupled to the cam and operable to reposition the cam along a second axis. A third linear control is coupled to the cam and operable to resist movement of the cam along a third axis.

Abstract: The present invention discloses a method for controlling a high-lift device or a flight control surface of an aircraft or spacecraft, especially with a system according to the present invention, comprising the steps of receiving, at least one first control unit, a command signal from a commander unit via a data network, providing a primary control signal to at least one secondary control unit via the data network, wherein the primary control signal depends on the received command signal, receiving, at the at least one second control unit, a sensor signal of one or more sensors of the high-lift device or flight control surface, and providing a secondary control signal to one or more actuators of the high-lift device or flight control surface, wherein the secondary control signal depends on the received sensor signal. Furthermore, the present invention discloses a system and an aircraft or spacecraft.

Abstract: A system for pneumatically actuating a split flap hingedly mounted near or at a trailing edge of an airfoil. The system includes a bladder system disposed between the split flap and the upper surface of the airfoil. The split flap is a small-chord (usually 1-3% of total wing chord) long-span lower panel which separates from the airfoil trailing edge by means of a hinge or a flexible lower skin. Deployment of the split flap is actuated pneumatically by the inflatable bladder system. The split flap may exist at a fixed wing trailing edge, a moving flap trailing edge, or an empennage trailing edge. The pneumatic bladder provides distributed force to extend and retract the split flap. This pneumatic approach eliminates extra drag, reduces cost and weight, and lessens flutter concerns.

Abstract: The present invention provides an actuator device which can prevent a rapid change in deflection angle when an active actuator and a standby actuator are switched. A first actuator 21A drives an aileron upon supply of hydraulic oil through a hydraulic system 31L from a hydraulic power source 35L. A second actuator 21B drives the aileron upon supply of hydraulic oil through a hydraulic system 31C from a hydraulic power source 35C. In normal cruising, the actuator 21A functions in an active state, and the actuator 21B functions in a standby state. When a hydraulic pressure PA of the hydraulic oil in the system 31L is equal to or less than a predetermined first threshold value PL in the above states, the actuator 21A is switched from the active state to a standby state, and the actuator 21B is switched from the standby state to an active state.

Abstract: A bearing assembly for mounting a pair of spaced parallel actuators between a wing and a control surface of an aircraft so that the actuators control deployment of the control surface from the wing in tandem. The bearing assembly includes a fixed member for attachment to the aircraft and a movable member attachable to the actuators. The fixed and movable members are coupled via a part-spherical bearing and are configured such that the part-spherical bearing is located in the space between the actuators.

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 aircraft reaction link includes a pair of linear portions, a coupling portion that couples ends of the pair of linear portions on the same side to each other, a fulcrum shaft attachment portion, and a cylinder attachment portion. A metal frame that is made of a metallic material and that is provided so as to extend across the coupling portion and the pair of linear portions and a composite frame that is made of fiber-reinforced plastic and that is provided so as to extend at least in the pair of linear portions are bonded together, and thereby, the pair of linear portions and the coupling portion are formed. The metal frame is provided so as to extend from the fulcrum shaft attachment portion to the cylinder attachment portion.

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 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: A pump unit installed inside a wing 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 the pump. A wing structure portion forming the surface structure of the wing is provided with an inlet port and an exhaust port that are formed therethrough. The inlet port is provided so as to be opened and closed by an inlet port opening/closing portion, and the exhaust port is provided so as to be opened and closed by an exhaust port opening/closing portion.

Abstract: The invention relates to a system for regulating a temperature of hydraulic fluid in at least one hydraulic circuit of an aircraft, wherein, in a periphery of the aircraft, additional hydraulic consumption units are each provided with at least one controllable valve for heating the hydraulic fluid above an adjustable minimum temperature limit value (Tmin), wherein the controllable valve is closed during a maneuver of the aircraft, wherein the controllable valves of the additional hydraulic consumption units are closed when the temperature of the hydraulic fluid exceeds an adjustable theoretical temperature value (Tsoll), and wherein a temperature control system is connected to a temperature sensor for adjusting the controllable valve on the basis of a local temperature of the hydraulic fluid, the temperature sensor being near the hydraulic consumption units which are arranged in wings and tail of the aircraft.

Abstract: An actuator system (14) comprising a valve assembly (40) having an inner spool (50), an outer spool (60), and a sleeve (70). An assembly (80) directly drives the inner spool (50) to move it relative to the outer spool (60), and thereby hydromechanically causes the outer spool (60) to move relative to the sleeve (70). A control assembly (90) provides current input to the drive assembly (80), which converts current input into mechanical motion. The control assembly (90) senses the position of the inner spool (50) and regulates current in accordance with the sensed position.

Abstract: When a loss or degradation in the function of one of a first aircraft central hydraulic power source and a second aircraft central hydraulic power source occurs, a controller performs a control so as to operate the other backup hydraulic pump, out of a first backup hydraulic pump and a second backup hydraulic pump, which is connected downstream of the other hydraulic power source, which is the other of the first and second aircraft central hydraulic power sources. Oil flowing through the other backup hydraulic pump is cooled by an oil cooler of the other hydraulic power source by operation of the other backup hydraulic pump in a state where the other hydraulic power source is being operated.

Abstract: An actuation system includes: two hydraulic actuators, each being coupled to a load-bearing structural component and to an elevator, and displaceable relative to the load-bearing structural component for the operation of the elevator; a first control valve device to control the first actuator; a second control valve device to control the second actuator; a control and monitoring unit that determines and transmits appropriate command signals to the control valve devices to execute appropriate control movements for the operation of the actuators; a locking device coupled to the first actuator to secure the first actuator a locked state, which is hydraulically connected with the first hydraulic system and the second hydraulic system via an hydraulic OR-circuit, where in normal operation the locking mechanism is in its unlocked state, and in the event of a failure of both hydraulic systems the locking mechanism secures the first actuator.

Abstract: An actuator includes a plurality of cylinder units each including a cylinder and a rod, and is pivotably attached, at a first end thereof, to a control surface. A reaction link is pivotably attached, at a first end thereof, to a fulcrum shaft rotatably supporting the control surface. The reaction link is pivotably attached, at a second end thereof, to a second end of the actuator between the plurality of cylinder units.

Abstract: A projectile is guided by employing heat, generated from the aerodynamic heating of a surface of the projectile while the projectile is in flight, to vaporize a material stored within the projectile. The resulting gas is used to guide the projectile.

Abstract: A hydraulic actuator system 1 is provided for driving a movable portion attached to a wing 5 of an airplane. A plurality of first system hydraulic actuators 2 and a plurality of second system hydraulic actuators 3 are included and are connected to the movable portion. The first system hydraulic actuators 2 are attached to the wing 5 and connected to a first hydraulic pressure source, whereas the second system hydraulic actuators 3 are attached to the wing to be in parallel to the first system hydraulic actuators 2 and are connected to a second hydraulic pressure source. The first system hydraulic actuators 2 and the second system hydraulic actuators 3 are alternately provided at a predetermined part of the wing 5.

Abstract: The present invention relates generally to flow regulation in aircraft systems. More particularly, the present invention relates to a flow regulator 100 and a high-lift system 200 for an aircraft incorporating such a flow regulator 100. The flow regulator 100 comprises a fluid input port 102 for receiving a pressurised fluid, a fluid output port 104 for providing fluid having a regulated flow rate, and a regulator valve 106 connected in fluid communication between the fluid input port 102 and the fluid output port 104. The regulator valve 106 is operable to provide regulated fluid at a substantially constant output flow rate to the fluid output port 104. The flow regulator 100 also includes a flow switching mechanism 108 for switching the substantially constant output flow of the flow regulated fluid provided by the regulator valve 106 between a first flow rate and a second flow rate, the first flow rate being less than the second flow rate.

Abstract: A bearing assembly for mounting a pair of spaced parallel actuators (3) between a wing and a control surface of an aircraft so that the actuators control deployment of said control surface from the wing in tandem is disclosed. The bearing assembly comprises a fixed member (18) for attachment to the aircraft and a movable member (8) attachable to the actuators. The fixed and movable members are coupled via a part-spherical bearing (14) and are configured such that the part-spherical bearing is located in the space between the actuators (3).

Abstract: A structural assembly for an aircraft comprising a cover with outer and inner faces, an inner support element and urging means on the inner support element, wherein opposing edges of the cover are fixedly mounted relative to the inner support element, and the urging means acts on the inner face of the cover to urge the cover to distend outwardly and locate in a predetermined position. Another aspect of the invention relates to a method of forming a structural assembly for an aircraft comprising an cover with outer and inner faces, an inner support element and urging means on the inner support element, the method comprising the steps of fixedly mounting opposing edges of the cover relative to the inner support element and operating the urging means to act on the inner face of the cover so that the cover is urged to distend outwardly into a predetermined position.

Abstract: An actuation system for trailing-edge flap control suitable for use in reducing vibration in rotorcraft blades as well as primary flight control and noise mitigation employing an antagonistic pair of fluidic artificial muscles (FAMs) located and operated inside the rotor blade. The FAMs are connected to a force transfer mechanism such as an inboard bellcrank and engaged to an outboard bellcrank by one or more linkages running spanwise out through the spar. The outboard mechanism translates the spanwise linkage motion into chordwise motion of a flap control rod which is connected to the trailing-edge flap. A torsion rod flexure (TRF) device is included connecting the trailing-edge flap to the blade. The actuation system can produce large flap deflections at relatively high operating frequencies for vibration reduction and noise cancellation and is capable of larger flap deflections at lower operating frequencies for embedded primary control of the rotorcraft.

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: 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: 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: An actuation system for a hydraulically operable aircraft elevator, includes: two actuators, supplied from at least one hydraulic supply system, of which each is coupled to a load-bearing structural component, and to an elevator, displaceable relative to the load-bearing structural component, for the operation of the elevator, a first control valve device, hydraulically connected to the first actuator to control the latter, and a second control valve device, hydraulically connected to the second actuator to control the latter, which devices are respectively supplied from an hydraulic system, a control and monitoring unit functionally connected with the control valve devices, which determines appropriate command signals for the control valve devices and transmits the signals to the latter to execute appropriate control movements for the operation of the actuators, a locking device, coupled to the first actuator, with a locking mechanism to secure the extended position of the pushrod of the first actuator in it

Abstract: A hydraulic actuator includes, but is not limited to a first outflow and a second inflow and outflow as well as a closure device for closing the second inflow and outflow in the direction of outflow. The first outflow an outflowing hydraulic fluid is subjected to flow resistance. When a switching position is reached, the second inflow and outflow is closed in the direction of outflow so that the hydraulic fluid has to be discharged through the first outflow while overcoming the flow resistance. This damps the movement of the hydraulic actuator in the region of an end position.

Abstract: A fluid contact surface actuation system for a vehicle, including a first fluid contact surface constructed and arranged to act against a first fluid passing over the first fluid contact surface; and a first fluid actuator coupled to the first fluid contact surface to move the first fluid contact surface between a first position and a second position to enable control of the vehicle in a predetermined manner, the first fluid actuator having a first resilient bladder that receives a second fluid such that pressure of the second fluid moves the first bladder between a contracted configuration and an expanded configuration.

Abstract: A method for actuating at least one positioning flap on each wing of an aircraft, which positioning flap is actuated by at least two flap coupling devices each including a flap actuator device, where at least on one flap actuator device for each positioning flap a brake mechanism is arranged by means of whose actuation an adjustment state of the respective flap actuator device can be locked, including the steps of: actuating each brake mechanism of a positioning flap individually, subsequently actuating the flap actuator device by means of the drive motor, in the case of a change in the adjustment state of the positioning flap by a predetermined extent, terminating actuation of the positioning flap concerned, and a system for implementing the method.

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: An electric actuator of the present invention includes a hydraulic fluid supply device and an actuator actuated in response to an input of hydraulic fluid from a variable-volume pump. The hydraulic fluid supply device includes an adjustable-speed motor, the variable-volume pump which is driven by the adjustable-speed motor and ejects hydraulic fluid, an electric motor control unit which controls the adjustable-speed motor so as to achieve an intended rotation speed, and a pump control unit which controls the variable-volume pump so that the ejection volume of the variable-volume pump decreases with an increase in the ejection pressure of the variable-volume pump.

Abstract: A flow control actuator 10 for a body exposed to fluid flow comprises first and second flow surfaces 14, 16 spaced to define an elongate gap 20 therebetween. An elongate control element is disposed in or adjacent the gap and has an externally facing arcuate surface and defines a first slot 24 between the control element and said first flow surface and a second slot between said control element and said second flow surface 26. The control element 22 is moveable to allow the width of said slots to be adjusted generally proportionally. The flow of pressurised flow control fluid passes from a plenum chamber 30 out through one or both slots 24, 26 to be deflected around the arcuate surface under the influence of the Coanda effect. The flow control actuator 10 may be provided in the trailing or leading edge of an aerofoil; alternatively it may be provided between two surfaces spaced to provide said gap but otherwise providing a generally continuous surface.

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: 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: Adjustment mechanism for adjustment area of variable-shape flow surface with two opposite skin surfaces includes plurality of whirl chambers swivelably arranged next to one another so that the whirl chambers are swivelable relative to one another, and the whirl chambers include lateral stiffening elements and longitudinal stiffening elements. Joints are structured and arranged to jointedly couple lateral stiffening elements and longitudinal stiffening elements of the whirl chambers, and a first drive tube section and a second drive tube section are arranged between adjacent longitudinal stiffening elements. A pump is coupled to first and second drive tube sections and control device is functionally connected to pump to swivel adjacent whirl chambers around joint axes via complementary volume changes in first and second drive tube sections. The instant abstract is neither intended to define the invention disclosed in this specification nor intended to limit the scope of the invention in any way.

Abstract: A cellular actuator device suitable for a control surface of an aircraft or space craft has a number of elementary cells which are combined to a common arrangement and are formed by pressure-tight chambers. The elementary cells can be acted upon by a pressure medium and, when acted upon by the pressure medium, can be deformed in at least one work direction while changing their length. The elementary cells are coupled for combining their length changes in the at least one working direction to an overall movement of the elementary cell arrangement. The actuator device is particularly suitable for actuating a control surface of an aircraft or spacecraft.