46results about "Without power ampliication" patented technology

An electric flight control system comprising an actuation subassembly, having three external duplex computers for controlling an electric actuator. The electric actuator has three stators electrically connected to three respective switching devices, the switching devices being electrically connected to a same internal simplex computer of the electric actuator, each switching device being connected to a respective external duplex computers, each switching device being configured to connect the internal simplex computer or respective external duplex computer to the associated stator, the internal simplex computer being self-monitored and communicating with the three external duplex computers.

The present disclosure relates to a mechanical actuator having a modified locking mechanism and fewer components. The actuator has a cylinder, a locking recess formed on an interior wall of the cylinder, and a piston assembly that moves between an extended position and a retracted position responsive to fluid pressure within the cylinder. A lock is connected to the piston assembly and moves radially between a locked position and an unlocked position responsive to the fluid pressure within the cylinder.

This invention belongs to the field of propulsion and control technology for vertical takeoff and landing (VTOL) aircraft, and discloses a cycloidal propeller control mechanism and method for driving trailing edge flap deflection using a dual four-bar linkage mechanism. Based on the traditional single four-bar linkage mechanism used to control the periodic pitch motion of the cycloidal blades, this mechanism adds a four-bar linkage mechanism with a similar structure and principle, specifically designed to independently drive the trailing edge flaps for periodic deflection. By separately setting the pitch eccentricity and pitch eccentricity angle of the pitch control four-bar linkage, and the deflection eccentricity and deflection eccentricity angle of the flap deflection four-bar linkage, the pitch motion of the blades and the deflection motion of the flaps can be independently controlled. Through numerical optimization design, a preset phase difference can be established and maintained between the deflection motion of the flaps and the pitch motion of the blades, and the maximum deflection angle of the flaps is constrained within the target amplitude range, thereby achieving effective improvement and active control of the cycloidal propeller's aerodynamic performance.

A fixed-wing aircraft (10) includes: a fuselage (14), a closed wing (16), a first actuator device (22), and a first connector device (24). The closed wing (16) is coupled to the fuselage (14) and has a wingtip (18) and a first deformable wing portion (20). The first actuator device (22) is arranged to supply a first input torque to the first connector device (24) when the first actuator device (22) is actuated, and the first connector device (24) is arranged to apply a first output torque to the first deformable wing portion (20) when the first input torque is supplied to the connector device. The first deformable wing portion (20) is arranged to elastically deform when the first output torque is applied to the first deformable wing portion (20). The closed wing (16) forms an outer support (26) at the wingtip (18), which is arranged to counteract the first output torque.

The present invention relates to a body (G), at least one rotary part (2) arranged on the body (G) and rotatably connected relative to the body (G), at least one bracket (S) located on the body (G), at least one support rod (3) pivotably connected to the bracket (S), at least one actuator (4) connected to the rotary part (2), which is a trigger exerting a force along the direction along which it extends, and which causes the rotary part (2) to rotate, at least one reaction rod (5) which is a structural part connected at one end to the rotary part (2) and at the other end to the support rod (3), at least one shaft (6) rotating with the rotary part (2) about the axis along which it extends, in order to rotate the rotary part (2), a first lever (7), which is a structural part extending outwardly from the shaft (6) and connected to the actuator (4), allowing the force of the actuator (4) to be transmitted to the shaft (6).

The application discloses a folding rudder wing operating mechanism, which comprises a mounting seat body, a cavity is formed in the middle position of the mounting seat body, a rudder handle is mounted in the upper end of the cavity, a folding rudder wing is connected to the end of the rudder handle away from the cavity, an adapter plate is mounted outside the lower end of the cavity, a push rod is connected to the adapter plate, a linkage mechanism is arranged in the cavity, and the linkage mechanism is used for synchronously rotating the adapter plate and the folding rudder wing; the overall structure has a short thrust transmission path, and the response speed of folding rudder wing operation can be effectively improved.

The present invention relates to an electric flight control system (10) comprising an actuation subset (C), having three external duplex computers (21, 24, 27) for controlling an electric actuator (40). The electric actuator (40) has three stators (42, 43, 44) electrically connected to three respective toggle devices (50), the toggle devices (50) being electrically connected to the same internal simplex computer (80) of the electric actuator (40), each toggle device (50) being connected to one of the respective external duplex computers (21, 24, 27), each toggle device (50) being configured to connect to the associated stator (42, 43, 44) the respective internal simplex computer (80) or external duplex computer (21, 24, 27), the internal simplex computer (80) being self-monitored and communicating with the three external duplex computers (21, 24, 27).

A system configuration significantly reduces the number of interconnections needed between an electromechanical actuator (EMA) and an EMA controller (EMAC) while maintaining redundancy and other safety aspects needed by the target application. This allows simplification of communication interfaces between multiple sensors in the EMA and EMAC and unification of the interface to multiple sensors inside the EMAC, leading to cost and weight reduction that can be significant for many aerospace products and applications, including unmanned vehicle (e.g., UAM / UAV) applications.

An aerial mobility vehicle includes a wing extending from a fuselage of the aerial mobility vehicle; a folding portion disposed at an edge portion of the wing, configured to extend from the wing to form a portion of the wing during unfolding of the folding portion, and configured to move to coincide with the wing to reduce an air resistance area of the wing in a vertical direction during folding of the folding portion; an actuator connected to the folding portion and configured to power the folding portion to unfold or fold to the wing; and a controller connected to the actuator and configured to control the actuator to fold the folding portion during vertical takeoff and landing of the fuselage, and configured to control the actuator to unfold the folding portion during cruising of the fuselage.

The subject matter of this specification can be embodied in, among other things, a rotary actuator having a housing with a first mounting assembly, a first mounting bracket adapted for attachment to the first mounting assembly and attachment to a first external mounting connector of a first mounting surface, a second mounting bracket adapted for attachment to the first mounting assembly and attachment to the first external mounting connector, a rotor assembly rotatably journaled in said housing and having a rotary output shaft having a second mounting assembly, a third mounting bracket adapted for attachment to the rotary output shaft and adapted for attachment to a second external mounting connector of a second mounting surface, and a fourth mounting bracket adapted for attachment to the rotary output shaft and adapted for attachment to the second external mounting connector.

A mechanical flight control system of an aircraft provided with an actuator and a force-measuring device, the actuator comprising at least one electric motor and one duplex calculator per motor, each duplex calculator comprising a control channel, a monitoring channel and two input interfaces, the force-measuring device comprising a plurality of measurement channels each comprising a sensor, an electronic module and an output interface connected to one of the input interfaces. the electronic module is connected to the sensor in order to generate an analog measurement signal and is configured to digitize the analog measurement signal, at least one of the measurement channels applying a delta-sigma modulation, and then to convert it into a digital transfer signal carrying a transfer value representative of a force in the mechanical flight control system.

The present invention relates to a release unit (1) for the controlled release of a mechanical lock, comprising: an actuator housing (10); at least one release element (2) extending through the actuator housing (10) and arranged to be axially displaceable within the actuator housing (10); at least one release unit (7) arranged in the actuator housing (10); at least one tolerance compensation spring (4) supported in the actuator housing (10) on one side and on the release element (2) on the other; at least one actuator spring (5) supported in the actuator housing (10) on one side and in the release unit (7) on the other; and a contact surface (9) provided in the actuator housing (10), wherein the at least one release unit (7) comprises a release lever (30); each release lever (30) is operatively connected to a first and a second transmission lever (34, 35);at least one structural component (23) fixed in the actuator housing (10), providing an element holder (21) having a release element (22) and a guide track for the release lever (30), the first translation lever (35) and the second translation lever (34); and comprising at least one stop element (8) connected to the at least one locking element (2), as well as a method for controlled release of a mechanical locking element by means of a release unit (1) and a missile positioning system comprising a release mechanism with at least one such release unit (1).

The invention relates to an aircraft wing comprising a wing structure and a spoiler. The spoiler is movable between a stowed configuration and a deployed configuration in a "pop-up" manner. The spoiler comprises a hinged top flap movable between a first position, in which the hinged top flap is constrained by an actuator, and a second position, in which the hinged top flap is unconstrained by the actuator. When the hinged top flap is in the second position, airflow across the top surface of the aircraft wing acts to pull the spoiler to the deployed position. The invention also relates to a method of deploying a stowed spoiler in an aircraft wing and a method of retracting a deployed spoiler in an aircraft wing.

The present invention relates to a mechanical flight control system (2) for an aircraft equipped with an actuator (30) and a force measurement system (40), said actuator (30) comprising at least one electric motor and one duplex computer per motor, each duplex computer comprising one control channel, one monitoring channel, and two input interfaces, said force measurement system (40) comprising several measurement channels, each comprising a sensor, an electronic module, and an output interface connected to one of said input interfaces. Said electronic module is connected to said sensor to generate an analog measurement signal and is configured to digitize said analog measurement signal, at least one of said measurement channels applying delta-sigma modulation, and then to transform it into a digital transfer signal carrying a transfer value representative of a force in said mechanical flight control system (2).

The invention relates to a mechanical flight control system of an aircraft for measuring forces in the mechanical flight control system. The mechanical flight control system (2) is provided with actuators (30) comprising at least one electric motor and one duplex calculator per electric motor, each duplex calculator comprising a control channel, a monitoring channel and two input interfaces, and with a force measuring device (40) comprising a plurality of measuring channels, each measuring channel comprising a sensor, an electronic module and an output interface connected to one of the input interfaces. The electronic module is connected to the sensor so as to generate an analog measurement signal and is configured to digitize the analog measurement signal and then convert it into a digital transmission signal carrying a transmission value representative of a force in the mechanical flight control system (2), at least one of the measuring channels applying a delta-sigma modulation.

A system for controlling actuators in an aircraft. The system including at least a first actuator and a second actuator. Each of the actuators including a movement member, at least one measurement sensor, and a computing member internally connected to the movement member and configured to be connected to at least one electronic computer of the aircraft. The computing member having a function for controlling the movement member, the measurement sensor of the first actuator being externally connected to the computing member of the second actuator, the computing member of the second actuator also having a monitoring function for monitoring the movement member of the first actuator.

An electromechanical actuation arrangement for a non-mass-balanced primary flight control surface establishes and retains a torsional preload between distributed actuator outputs to eliminate surface freeplay and increase mechanical impedance. Preload is retained mechanically via an interconnect shaft and / or electronically by closed-loop synchronization that maintains a commanded differential torque. An independent flutter-suppression device provides damping and / or locking upon oscillation detection or disconnection. The architecture supports multiple embodiments, including fewer motors than transmissions and a transmission without its own motor, and enables reduction or elimination of hydraulic systems for primary flight controls while meeting aeroelastic requirements under normal and single-fault conditions.

A safety element for an aircraft actuation system comprising an actuator, an inceptor, and a shear pin connecting the motor to the inceptor is described. The safety element comprises an absorber zone which is configured to deform in response to an excess force event in the aircraft actuation system, and is locatable between the shear pin and the actuator. The safety element reduces a force applied to an element such as the shear pin during the excess force event. Especially, where the excess force event is caused by the actuator, the shear pin is preventable from breaking.

High-lift systems (300, 1300) for aircraft (100), and associated methods are provided. A high-lift system includes a movable high-lift device (20A), an inboard track (33) to guide an inboard portion (46) of the high-lift device, and an outboard track (34) to guide the outboard portion (48) of the high-lift device. An actuator (40) drives the high-lift device between the retracted position and the deployed position. The actuator is disposed inboard of the inboard track. An inboard bell crank (42) connects the actuator to the inboard portion of the high-lift device. An outboard bell crank (44) connects the actuator to the outboard portion of the high-lift device. A connecting rod (58, 62) connects the inboard bell crank to the outboard bell crank so that the outboard bell crank is drivingly connecting the actuator to the outboard portion of the high-lift device via the connecting rod and the inboard bell crank.

The present invention relates to an electric flight control system (10) comprising an actuation subset (C), having three external duplex computers (21, 24, 27) for controlling an electric actuator (40). The electric actuator (40) has three stators (42, 43, 44) electrically connected to three respective toggle devices (50). The toggle devices (50) are electrically connected to the same internal simplex computer (80) of the electric actuator (40). Each toggle device (50) is connected to one of the respective external duplex computers (21, 24, 27). Each toggle device (50) is configured to connect to the associated stator (42, 43, 44) either the internal simplex computer (80) or the respective external duplex computer (21, 24, 27). The internal simplex computer (80) is self-monitoring and communicates with the three external duplex computers (21, 24, 27). (Shorthand figure: Figure 1)

A no-back brake includes a torque tube and a brake. The brake is configured to be axially displaced in response to a force to prevent the torque tube from rotating. The brake includes a rotor that is concentric with the torque tube and that is fixed to the torque tube so that the rotor rotates with the torque tube. The brake includes a stator that is concentrically mounted on the torque tube so that the torque tube is rotatable relative to the stator. The brake includes a friction device located between the rotor and the stator. The friction device is configured to form a frictional coupling between the rotor and the stator that opposes relative rotational motion between the rotor and the stator when the brake is axially displaced. The frictional device includes a composite material.

The present invention relates to a mechanical flight control system (2) for an aircraft equipped with an actuator (30) and a force measurement system (40), said actuator (30) comprising at least one electric motor and one duplex computer per motor, each duplex computer comprising one control channel, one monitoring channel, and two input interfaces, said force measurement system (40) comprising several measurement channels, each comprising a sensor, an electronic module, and an output interface connected to one of said input interfaces. Said electronic module is connected to said sensor to generate an analog measurement signal and is configured to digitize said analog measurement signal, at least one of said measurement channels applying delta-sigma modulation, and then to transform it into a digital transfer signal carrying a transfer value representative of a force in said mechanical flight control system (2).Abbreviated figure: figure 1.

This invention relates to tools and components for adjusting and holding a chain. The chain (26) has links (26a) that engage with a sprocket (24). The tool (100) includes: a boot-shaped member (102) intended to be fixedly mounted relative to the sprocket (24); and a finger (104) fastened to the boot-shaped member (102) and intended to be inserted into the link (26a). With this arrangement, the chain is adjusted into place and held in position.

A screw actuator (16) is disclosed. The screw actuator (16) includes a primary ball nut (18), a secondary nut (20), and a ball screw (22). The primary ball nut (18) and secondary nut (20) are both in threaded engagement with the ball screw (22), in which the inner functional plays of the screw actuator (16) are configured such that in normal operation the threads of the secondary nut (20) are in contact with the ball screw threads.

Gearbox assemblies for actuators and aircraft actuator systems including the same. The gearbox assemblies include a housing, a gear assembly arranged within the housing, and a torque limiting sleeve arranged on an exterior of the housing, wherein the torque limiting sleeve is axially positioned radially outward from the gear assembly.