449results about "Without power ampliication" patented technology

A system for identifying a failure of the primary load path in an actuator. The actuator includes a screw shaft and a nut assembly moveable along the screw shaft. The nut assembly comprises a primary nut for transmitting load through the actuator along a primary load path, and a secondary nut for transmitting load through the actuator along a secondary load path. The nut assembly is configured such that, in normal operation, the secondary nut is not engaged with the screw shaft and load is transmitted through the actuator along the primary load path, and such that, upon failure of the primary load path, the secondary nut is configured to engage the screw shaft and transmit load through the actuator along the secondary load path. The system includes a strain gauge configured to be mounted to the actuator and to measure vibrations of the actuator.

The present disclosure relates to a system for detecting the position and / or the operating state of a movable component of an aircraft, wherein the system has one movable component, two actuators, two motion sensors, one asymmetry sensor, one drive train, and one monitoring device; wherein the component can be actuated by the at least two mutually offset actuators; wherein at least one motion sensor is provided for one respective actuator; wherein the motion sensor can monitor a mechanical movement of the drive train; wherein the monitoring device is connected to the motion sensors and to the asymmetry sensor, and wherein the system is configured such that an initialization of the monitoring device and / or of the motion sensors and / or of the asymmetry sensor can take place after a booting of the system and / or when the component has adopted a fully retracted and / or a fully extended position.

A lead screw control actuation system. A system according to an embodiment includes a motor configured to rotate a lead screw that is configured to increase torque provided by the motor. The lead screw is coupled to a nut through a threaded opening such that rotation of the lead screw causes the nut to move along a longitudinal axis of the lead screw. The nut includes a slot oriented in a direction perpendicular to the direction of motion of the nut. The system further includes a lever arm including a pin at a first end of the lever arm, the pin configured to slide in the slot of the nut to cause the lever arm to rotate in response to the motion of the nut. A second end of the lever arm is coupled to a canard such that rotation of the lever arm causes the canard to rotate.

An unmanned aerial vehicle (UAV) having wings stowed against a fuselage of the UAV in a first arrangement is disclosed. Methods and systems for deploying the wings into a second arrangement are disclosed. For example, after a launch of the UAV, the UAV monitors for at least one precondition. The at least one pre-condition being a pre-condition associated with deploying wings of the UAV into the second arrangement. Upon detecting the at least one precondition, the wings of the UAV are deployed into a second arrangement. Deploying the wings comprises activating, in response to detecting the at least one precondition associated with the UAV, a gearbox configured to transition the wings from the first arrangement to the second arrangement. Roll control may be maintained throughout launch and deployment.

An aircraft includes a wing and a leading-edge aerodynamic surface coupled to the wing. An actuator is disposed within the wing and coupled to the leading-edge aerodynamic surface via a linkage system. Two or more hinge systems are coupled to the leading-edge aerodynamic surface and configured to rotate about a fixed axis internal to the wing to move the leading-edge aerodynamic surface between a retracted position and a drooped position.

A mechanism for moving an aircraft wing slat between a retracted position and a deployed position includes a rib, an arm, a first bar and a second bar. The arm is disposed generally parallel with the rib and is configured for attachment with the wing slat. The rib, the arm and the first and second bars cooperate to form a four-bar linkage which is rotatable about connection points in a negative pitch rotational direction to dispose the arm in the deployed position, and in a positive pitch rotational direction to disposing the arm in the retracted position. The mechanism includes additional structures for maintaining the arm in the deployed position, including a magnetic extend stop and a standoff, an arm lobe and a cross-bolt, and a bar lobe and a linear member.

A device may include a locking assembly configured for coupling with an aileron of an aircraft, the locking assembly includes a lock configured for coupling with the aileron. The locking assembly may include a lock mechanism operatively coupled with the lock and in communication with an accelerometer, where the lock mechanism is configured to transition the lock from a locked configuration to a released configuration. The lock mechanism opens the lock based on a specified acceleration measured with the accelerometer, and the lock frees the aileron to a dynamic configuration. A device may include a tailored aileron deployment system including a biasing member coupled with the aileron, in the released configuration the biasing member controls movement of the aileron in the dynamic configuration.

An apparatus for adjusting aircraft inceptor trim includes an inceptor centring device mechanically coupled to a control shaft that rotates in response to inceptor inputs. The inceptor centring device includes an inceptor centring spring preloaded with a preload force to provide a first biasing force to bias the control shaft to an inceptor neutral position. The apparatus has a trim mechanism, mechanically coupled to the inceptor centring device, and includes a trim locking device having a locked state and an unlocked state. A resistance force of the trim mechanism with the trim locking device in the unlocked state is less than the preload force, such that when the trim mechanism is in the unlocked state, actuation of the inceptor adjusts the position of the inceptor centring device to alter the degree of trim.

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.

A differential actuator system for controlling a flight control surface is provided, which includes a first control lane subsystem, a second control lane subsystem, a differential subsystem and a differential lock subsystem. The differential subsystem includes a first gear set forming a first load path from the first control lane to the flight control surface, and a second gear set forming a second load path from the second control lane to the flight control surface. The differential lock subsystem operably couples to the first gear set and the second gear set and is switchable between a locked state to couple the first and second gear sets in a torque-summing mode for moving the flight control surface and an unlocked state to couple the first and second gear sets in a velocity-summing mode for moving the flight control surface. Control signals issued from a controller operate the differential actuator system.

An emergency autoland braking system for aircraft includes an electromechanical actuator configured to provide a braking force to rudder pedals for actuating a braking system without altering the existing braking system of the aircraft. An equalized braking force may be applied to both rudder pedals via a force balancing mechanism and a single linear actuator, or differential braking force may be applied via independent linear actuators. The system is compatible with both open and closed loop control for commanding a braking force.

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.

The invention discloses an aircraft flap and an aircraft flap control method. The aircraft flap comprises a flap body, a translation mechanism, two rotating mechanisms and a controller. The translation mechanism comprises a first connecting assembly and a first driving assembly, one end of the first connecting assembly is fixedly connected to the fixed rear edge of the wing, the other end of the first connecting assembly is rotationally connected with a connecting joint of the flap, and the first driving assembly can drive the flap to unfold or fold; the two rotating mechanisms correspond to two end ribs of the flap in a one-to-one mode, each rotating mechanism comprises a second connecting assembly and a second driving assembly, one end of each second connecting assembly is movably installed on the fixed rear edge, the end ribs are movably connected with the second connecting assemblies, and the second connecting assemblies are further in transmission connection with the second driving assemblies; the second driving assembly can drive one end, connected with the fixed rear edge, of the second connecting assembly to rotate; the first driving assembly and the second driving assembly are electrically connected with the controller. According to the airplane flap, translation and rotation of the flap can be independently controlled, and the working performance of the flap is effectively improved.

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.

A flap drive mechanism for moving an actuating flap of an aircraft is provided. In one aspect, a flap drive mechanism includes a base mount, links, an actuator, a first track having a first rail, a second track pivotably coupled with the first track and having a second rail, a third track pivotably coupled with the second track and having a third rail. The second and third tracks and a mounting link each provide a flap segment mount. The actuator is coupled with the base mount and arranged to drive the links so that the first rail translates relative to the base mount, the second track pivots relative to the first track while a bearing coupled with the third track slides relative to the second rail, and the third track pivots relative to the second track while a bearing coupled with the mounting link slides relative to the third track.

A system for identifying a failure of the primary load path in an actuator. The system includes a first strain gauge configured to be mounted to a static portion and to measure vibrations of the static portion and a second strain gauge configured to be mounted to a nut assembly and to measure vibrations of the nut assembly. The system comprises a processor configured to identify a failure of the primary load path by comparing the measured vibrations of the static portion and the measured vibrations of the nut assembly.

In the event of a failed engine, an automatic takeoff thrust asymmetry compensation system (“ATACS”) for an aircraft improves capabilities to reduce VMCG and deal with the potential side-effects simultaneously. The system commands selected control surfaces (which can be e.g., rudder and / or ailerons and / or spoilers or any combinations thereof) for a short period of time, improving the capability to reduce the VMCG without increasing the penalty on system failures or poor handling qualities.

A rotary electromechanical actuator (EMA) includes: a first EMA part comprising a first electric motor and a first rotary shaft (S1) extending along an axis and arranged to be rotated about the axis by the first electric motor, and a first 1st stage gearbox providing a geared transmission from the first electric motor to the first output shaft, and a first brake assembly for applying a braking force to the first output shaft; the EMA further comprising: a second EMA part comprising a second electric motor and a second rotary shaft (S2) extending along the axis and arranged to be rotated about the axis by the second electric motor, and a second 1st stage gearbox providing a geared transmission from the second electric motor to the second output shaft, and a second brake assembly for applying a braking force to the second output shaft.

The utility model belongs to the technical field of aircraft tests, and particularly relates to an aircraft course control stability augmentation system. The system comprises a composite rocker arm (4), wherein the composite rocker arm (4) comprises a first rocker arm (401), a second rocker arm (402) and a fan-shaped wheel (404); the fan-shaped wheel (404) is connected with a pedal through a steel cable (3), a second rocker arm (402) is fixedly connected to the disc face of the fan-shaped wheel (404), the tail end of the second rocker arm (402) is hinged to the middle of a first rocker arm (401) through a shaft B, the first end of the first rocker arm (401) is hinged to an output shaft of the stability augmentation steering engine (2) through a shaft C, the second end of the first rocker arm (401) is hinged to a pull rod (5) through a shaft A, and the tail end of the pull rod (5) is connected with a control valve rocker arm (6). The control valve (7) is used for driving the hydraulic actuator (8), and the control surface (9) is driven by the hydraulic actuator (8) to deflect. According to the invention, the course control stability is improved, and the onboard space and weight can be saved.

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.

An actuator unit includes first and second link mechanisms. The first link mechanism has first and second arms that extend in first and second directions when viewed in a direction parallel to an axis of rotation. The first arm is connected at a first connection point to a power source for powering the first link mechanism. The second arm is connected at a second connection point to the second link mechanism. The second link mechanism is connected at a third connection point to a locking collar. The locking collar is configured to reciprocate as the first link mechanism rotates in a reciprocating manner. The distance from the first connection point to the axis of rotation is greater than the distance from the second connection point to the axis of rotation.

A drive system of an aircraft, including a fuel cell, which can be supplied with hydrogen from a hydrogen tank and with air from a blower, the fuel cell being configured to provide drive power for operational flight after takeoff and before landing dependent on a hydrogen mass flow supplied by the hydrogen tank and dependent on an air mass flow supplied by the blower, and an electrical energy store, which is configured to provide additional drive power for takeoff and landing, wherein an additional hydrogen tank and an air or oxygen tank are configured to interact with the fuel cell such that the fuel cell can be supplied with an additional hydrogen mass flow and with an additional air or oxygen mass flow, thereby compensating at least partially for a loss of the additional drive power provided by the electrical energy store for landing.

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.

An aircraft wing including a wing structure supporting a wing cover having an outer aerodynamic surface, and a pop-up spoiler unit including an actuation mechanism and a spoiler coupled to the actuation mechanism. The spoiler is moveable by the actuation mechanism to deploy through an aperture in the wing cover. The pop-up spoiler unit is mounted to the wing structure to be substantially isolated from wing bending deflections. The wing structure has a neutral plane which is substantially neither stretched nor compressed during bending of the wing structure. The pop-up spoiler unit is mounted to the wing structure at only two pin joints each having a respective pin axis about which the pop-up spoiler unit is substantially free to rotate, and each pin axis substantially lies in the neutral plane.

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').

Describes a micro spoiler for enhancing the effectiveness of a lateral control surface of an aircraft wing. An example aircraft includes a wing, a lateral control surface, and a micro spoiler. The lateral control surface is movably coupled to the wing. The lateral control surface is movable between a neutral position, a first upward deflection position, and a second upward deflection position that extends beyond the first upward deflection position. The micro spoiler is located on or in front of the lateral control surface. The micro spoiler is movable between a retracted position and a deployed position. The micro spoiler is configured to move from the retracted position to the deployed position based on the lateral control surface moving from the neutral position to the first upward deflection position or moving toward the first upward deflection position.