229results about "Without power ampliication" patented technology

The invention discloses a dual-redundancy motor dual-sensor control surface steering engine transmission system and a heavy-load unmanned aerial vehicle. The transmission system comprises a rotor wing structure; the planetary gear structure comprises an input rod, an output rod, a first planetary gear set and a second planetary gear set, the input rod is connected with the first planetary gear set, the first planetary gear set is in transmission connection with the second planetary gear set, the output rod is connected with the second planetary gear set, and the rotor wing structure is installed on the output rod; the input gear is mounted on the input rod and is in transmission connection with the input rod; the driving structure comprises two sets of stator windings and an output shaft, each stator winding comprises a rotor, and the rotors of the two sets of stator windings are both installed on the output shaft; and the transmission gear is connected with the output shaft and meshed with the input gear. According to the dual-redundancy motor dual-sensor control surface steering engine transmission system, high reliability and redundancy can be guaranteed, and meanwhile miniaturization, light weight, high power density and excellent dynamic characteristics are considered.

The invention discloses a high-thrust high-reliability linear steering engine transmission system and an eVTOL aircraft, and the transmission system comprises an output shaft which is provided with a center hole, and the inner side wall of the center hole is provided with a threaded section; the transmission assembly is mounted in the center hole and is in threaded connection with the threaded section; the input shaft penetrates through the transmission assembly and is in transmission connection with the transmission assembly; the input shaft rotates to drive the transmission assembly and the output shaft to axially move; the driving structure drives the input shaft to rotate; the gear assembly is in transmission connection with the driving structure and the input shaft; the position detection structure comprises a sliding block and a sliding rail parallel to the output shaft, and the sliding block is installed on the sliding rail in a sliding mode and connected with the output shaft; the output shaft axially moves to drive the sliding block to move along the sliding rail; the driving structure comprises a first driving piece and a second driving piece, and one of the first driving piece and the second driving piece is selected to be connected with the gear assembly. The transmission system is large in output force and torque and high in reliability, and the displacement recognition precision and reliability of the output shaft are high.

An aircraft wing is disclosed having a fixed wing with a tip, and a wing tip device rotatably mounted on a hinge at the tip of the fixed wing, such that the wing tip device is rotatable about the hinge, and an actuation system for rotating the wing tip device about the hinge. The actuation system includes a motor, at least one geared rotary actuator, a reduction gearbox, a clutch for selectively decoupling rotation of the motor from rotation of the geared rotary actuator, the geared rotary actuator is driveable by the motor and arranged to convert rotary motion into a different rotary motion and is arranged to rotate the wing tip device relative to the tip of the fixed wing. The clutch is coupled to the geared rotary actuator by a shaft and the shaft is fixed in rotation to a mass disposed radially outwardly from the shaft.

The utility model belongs to the technical field of fixed wings, and provides a manned fixed wing multi-mode control device which comprises an operation table. The mechanical control assembly is installed in the operation table, the mechanical control assembly is used for manually driving the fixed wings, the mechanical control assembly comprises a support installed in the operation table, a first linkage rod is rotationally installed at the upper end of the support, a first transmission connecting rod is rotationally installed at the front end of the first linkage rod, and a first rocker is installed at the upper end of the first linkage rod; and the intelligent control assembly is installed in the operation table, the intelligent control assembly is used for intelligently controlling the fixed wings, and the intelligent control assembly comprises a supporting frame installed in the operation table. According to the utility model, the mechanical control assembly, the intelligent control assembly, the separation assembly and the limiting assembly are integrated, and through cooperative operation of the components, diversified control modes of the fixed-wing aircraft are realized, including mechanical control, semi-automatic control and full-automatic control.

The invention discloses a built-in main control actuator connecting structure of an aircraft wing box, and relates to the technical field of aircraft structure design, an actuator is located in the aircraft wing box structure, an actuating rod of the actuator penetrates through an actuator rod avoiding hole of the aircraft wing box structure, and the actuating rod stretches out and draws back to cooperate with the actuator to rotate up and down in the wing box to achieve deflection of a control surface. The invention is suitable for small airfoils and small and medium-sized airfoils with low airfoil height, solves the problems that the small airfoils of the small airfoils are small, the trailing edge space is limited, and a control surface actuating system cannot be mounted, and avoids the problem that the aerodynamic efficiency is influenced as an actuating structure is arranged outside the airfoils; the structure of the actuator connector is designed and arranged on the front main bearing beam of the wing box, the load transfer continuity of the wall plate structure is not broken, meanwhile, the actuating load can be directly diffused to the skin through the support connector, and the load transfer efficiency is higher. By separating a daily inspection and maintenance passage from a disassembly and assembly passage, the opening size of the airfoil wall plate is reduced, the load transfer capacity of the lower airfoil skin is guaranteed, and structural weight reduction is facilitated.

A harmonic gear system, having: a housing; an input shaft having an aft end within the housing that defines a wave generator; a wave generator bearing; first and second nested cup-shaped flex splines that are grounded to the housing, the first and second flex splines are biased by the wave generator via the wave generator bearing; an output shaft having a forward end that is within the housing, and a radial flange at the forward end; first and second nested circular splined shafts within the housing, the first and second circular splined shafts are connected to the radial flange of the output shaft and are driven by ones of the first and second flex splines, to thereby drive the output shaft; and a bearing between the first and second circular splined shafts.

The invention relates to a trailing edge flap pneumatic structure for mechanical control and jet flow resistance reduction of an unmanned aerial vehicle. The trailing edge flap pneumatic structure comprises a deflection angle detection device, a jet flow spraying rod, a jet flow air supply pipe, a controller and a flight control plate. The two deflection angle detection devices are arranged at the two trailing edge flaps and used for detecting the deflection angles of the two trailing edge flaps; the two jet flow spray rods are transversely connected to the upper portions of the two wings and located in front of the two trailing edge flaps, and a plurality of backward jet flow holes are formed in the jet flow spray rods. An inlet of the jet flow air supply pipe is sequentially provided with a jet flow control electromagnetic valve and a jet flow meter and connected with a high-pressure air source, and the jet flow meter is used for detecting the air flow in the jet flow air supply pipe; an outlet of the jet flow air supply pipe is divided into two branches which are connected with the two jet flow spray rods; the controller is connected with the jet flow control electromagnetic valve; the flight control board is connected with the two deflection angle detection devices, the controller and the jet flow meter, and negative feedback control over the air flow of the jet flow air supply pipe based on the trailing edge flap deflection angle is formed.

This actuation system (22) comprises: an actuating device (26) comprising a first end (30) configured to be connected to a main section (18) and a second end (32), which is designed to be connected to a control surface (20), and at least one hinge device (28) comprising at least one hinge element (34) and designed to guide a movement of the control surface (20) relative to the main section (18), wherein the actuating device (26) is configured to move its first (30) and second end (32) translationally along an actuator axis (A-A'), wherein the hinge device (28) defines a plurality of hinge axes (H-H') which intersect to define a pivot axis (P-P') wherein the pivot axis (P-P') is offset from the hinge device (28) wherein the hinge device (28) is configured to guide the movement of the control surface (20) in a rotation about the pivot axis (P-P').

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 invention belongs to the technical field of aerospace, and particularly relates to a deflected and shrunk flexible dorsal fin suitable for heading control of a tailless aircraft and the aircraft, and the deflected and shrunk flexible dorsal fin comprises a middle separation beam, a flexible skin, an air bag and an air pressure control device; wherein the middle partition beam is of a rigid supporting structure and extends to the front end of the dorsal fin to form a partition, and the dorsal fin is divided into a left independent structure and a right independent structure; the flexible skin covers the two sides of the middle separation beam to form a flexible deformable surface; the air bags are arranged on the inner side of the skin and connected with the air pressure control device, the air pressure control device fills the air bag on one side with air, the air bag on the other side sucks air, and the asymmetric state of the wing face of the dorsal fin is formed. The heading control requirement of the tailless aircraft is met, and meanwhile the body heading stability of the tailless aircraft is improved.

Disclosed are a wing for an aircraft, a leading edge high lift assembly and an aircraft, the wing comprising a main wing and a leading edge high lift assembly, the leading edge high lift assembly comprising a slat and a connecting assembly movably connecting the slat to the main wing, and a connecting assembly for connecting the slat to the main wing such that the slat is movable between a retracted position, a first extended position, and a second extended position, where the connecting assembly comprises an elongate slat track, a first end and / or a middle portion of which is mounted to the slat, the slat track being mounted to the main wing by means of a roller support. The aim of providing a wing with a beneficial ratio of aerodynamic performance to noise resulting from the extension of the slat is achieved, in which the slat track and the roller support are configured such that, in a retracted position and a first extended position, the slat rests against the main wing and, in a second extended position, the slat rests against the main wing. A gap is formed between the trailing edge of the slat and the leading edge of the main wing.

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 pulley frame assembly includes a pulley frame and rollers adjustablely mounted on the pulley frame. The pulley frame assembly further includes a roller adjustment device comprising: a transition frame rotatably supporting the rollers and movably connected to the pulley frame via a rotating connector; an adjustment member including an eccentric shaft passing through the transition frame and the pulley frame, the eccentric shaft including a first shaft section and a second shaft section eccentrically arranged, wherein the first shaft section is connected to the transition frame, and the second shaft section is connected to the pulley frame; and a rotating connector passing through the transition frame and the pulley frame at a distance spaced relative to the eccentric shaft, allowing the transition frame to rotate about the rotating connector. This device reduces the adverse effects of roller stress on the overall lifespan of the device, allows for quick confirmation of whether the rollers are in the correct position, and enables adjustment of the roller position over a wide range.

The invention discloses a wind pendulum and propeller device of a helicopter, belongs to the technical field of helicopter instrument production, and aims to solve the technical problems of torsion generated by rotation of a propeller, turning and flight power of the helicopter, flight control and the like by utilizing the wind pendulum and propeller device. According to the technical scheme, a propeller rotates anticlockwise to generate lift force, and the attack angle of blades is fixed; two groups of wind pendulums are mounted on a wind pendulum mounting frame below the propeller, are bilaterally symmetrical, are perpendicular to the longitudinal axis of the fuselage and droop naturally; one end of the long connecting rod is connected with the piston rod; the hydraulic system is composed of two oil cylinders and an oil pipe, and a tail end piston rod is connected with a control handle rod. One handle rod controls the deflection direction and angle of one wind swing set through the hydraulic system and the long connecting rod. The device is suitable for flight power and flight control devices of helicopters and unmanned aerial vehicles.

The invention provides a microminiature folding wing aircraft steering engine structure which comprises a body, at least four mounting positions are arranged on the body, an airfoil unfolding mechanism and an airfoil deflection mechanism are arranged on the mounting positions, the airfoil unfolding mechanism comprises a mounting seat, a pressure spring, a push block and a wing panel, the mounting seat is fixedly connected to the middle of the body, and the pressure spring is fixedly connected to the middle of the body. A guide hole horizontally and transversely formed along the body is formed in the mounting seat, the compression spring penetrates through the guide hole, the push block is slidably connected into the guide hole, the two ends of the compression spring are connected with the bottom of the guide hole and the push block respectively, the wing piece comprises a wing body and a first connecting part, one end of the first connecting part is fixedly connected to the side face of the wing body, and the other end of the first connecting part is rotatably connected to an orifice of the guide hole. The rotating axis of the first connecting part is horizontally arranged along the body, and the first connecting part is provided with a second connecting part connected with the push block. The technical problems that an existing micro-miniature folding wing steering engine is insufficient in space utilization rate, structural rigidity and system integration degree are solved.

Return brake for at least one control surface of an aircraft, comprising a spring arrangement (2), a ball ramp mechanism (5) and at least one coupling mechanism (4) for coupling a drive side with an output side of the return brake, wherein exactly one constant brake (6a, 6b) activated by means of the spring arrangement is provided, characterized in that the spring arrangement (2) is arranged between two parts (5a, 6b) of the ball ramp mechanism (5), between which a plurality of balls of the ball ramp mechanism (5) are also arranged.

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.

Described is a mechanical transmission (T) comprising a first rotary element (200) rotatable about a main axis (X) and a second rotary element (300) rotatable about the main axis (X). The mechanical transmission (T) also comprises a connecting element (400) coupled to the first rotary element (200) by a short-pitch threaded connection (50) and to a containment frame (100) by means of a long- pitch threaded connection (60). The long -pitch threaded connection (60) is located on two angular zones angularly opposite each other relative to the main axis (X) and aligned along a first direction (A) perpendicular to the main axis (X). The two angular zones are separated from each other by zones without the long-pitch threaded connection (60) in such a way that, in a plane (T) perpendicular to the main axis (X), the mechanical transmission (T) has a maximum dimension, along a second direction (B) perpendicular to the first direction (A), which is less than the maximum dimension along the first direction (A).

Pitch and roll input control devices for aircraft are described that utilize a multi-bar linkage mechanism to translate forward or aft movement of the steering wheel into flight controls in a compact volume. The devices replicate the motion of a large mechanism but use a more compact mechanism that does not require volume below the floor and eliminates the linkage that normally would reside between a pilot's legs.

Drive Component A drive arrangement for operating a vehicle control surface includes a first drive device, a second drive device, an output member for connecting to the vehicle control surface, first engagement means for selectively engaging and disengaging the first drive device and the output member, and second engagement means for selectively engaging and disengaging the second drive device and the output member.

The utility model belongs to the technical field of aircraft control surface structures, and particularly relates to an aircraft control surface transverse control device which comprises an angle rocker arm, an actuator, a pull rod and a control surface. One end of the angle-shaped rocker arm is hinged to an output shaft of the actuator, the other end of the angle-shaped rocker arm is hinged to the pull rod, the pull rod is hinged to the upper end of the control surface, and the lower end of the control surface is hinged to the machine body; the actuator stretches out and draws back to drive the angle-shaped rocker arm to rotate, and the angle-shaped rocker arm pulls the control surface to swing through the pull rod. The aircraft control surface adopting the transverse control device can meet the control suspension arrangement requirements of the control surface under the conditions that the design space is narrow and the deflection angle is large.

The invention provides a planetary aileron actuator, and belongs to the technical field of servo actuators. The aileron actuator comprises a double-winding motor used for providing a power source, a transmission assembly used for transmission and supporting, an output assembly connected with the transmission assembly and used for output, a double-LVDT position feedback assembly used for position feedback and a shell used for containing the output assembly and the position feedback assembly. According to the actuator, the extreme compactness and miniaturization of the actuator are realized, and the actuator has electrical dual redundancy, extremely high transmission precision, extremely high power density and extremely high load capacity.

An actuator for a control surface is provided. The actuator includes a frame, a motor, an epicyclic gear, a power-off brake, a power-on brake, and a one-way bearing. The sun gear is operably coupled to an output of the motor. A planet carrier can be coupled to the control surface such that rotation of the planet carrier in a first direction transitions the control surface towards a deployed position and rotation of the planet carrier in a second direction transitions the control surface towards a stowed position. The power-off brake is coupled to the sun gear to fix the sun gear upon power off. The energized brake is coupled to the ring gear to secure the ring gear when energized. The one-way bearing allows the ring gear to rotate in the second direction and prevents the ring gear from rotating in the first direction when the power-on brake is powered off.

A strain wave drive includes: an elliptical wave generator shaft rotatable about an axis (X); a flexible flex spline mounted around the wave generator shaft; a ring gear assembly mounted around the flex spline, the ring gear assembly comprising an output ring gear having a circular inner periphery and being sandwiched between two earth ring gears each having a circular inner periphery. The flex spline has a first set of a first number of radially outwardly extending teeth around its outer periphery. Each earth ring gear has the first number of radially inwardly extending teeth to engage with the first set of teeth of the flex spline. The flex spline has a second set of the first number of radially outwardly extending teeth around its outer periphery wherein the output ring gear has a second number of radially inwardly extending teeth.

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).

This invention belongs to the field of aerospace equipment design and relates to a device for the retraction and extension of a single-sided flap system and its arbitrary angle limiting mechanism. It mainly includes a mounting and fixing mechanism, a speed reduction and torque amplification mechanism, and a limit control mechanism. The mounting and fixing mechanism connects the device to the mounting bracket of the single-sided flap control system and the universal joint of the flap transmission rod. Then, a crank or other power tool is used to input high-speed, low-torque onto the input shaft of the speed reduction and torque amplification mechanism, which is transmitted to the flap transmission rod through the output shaft of the speed reduction mechanism, thus achieving flexible retraction and extension of the flap system. Furthermore, a limit control mechanism is added to the speed reduction mechanism to control the rotation direction of the device and the flap transmission rod, achieving arbitrary angle limiting of the single-sided flap system. This invention solves the problems of difficult retraction and extension and unlimited arbitrary angle limiting in existing single-sided flap control systems without a control mechanism.