579results about "Without power ampliication" patented technology

A flow control device having: a composition gradient defining a first coefficient of thermal expansion and a second coefficient of thermal expansion that differs from the first coefficient of thermal expansion; and a thermoelectric junction operationally coupled to the composition gradient, wherein the composition gradient is formed of either of a plurality of dissimilar metals or of plastic with fillings or fibers.

A strain wave-type gearbox comprises: a flex spline component; an outer circular spline component; and a wave generator component, wherein the wave generator component is configured for adjustment of at least one dimension of the wave generator component, to cause engagement and disengagement of the flex spline component and the outer circular spline component. A drive arrangement for manipulating a vehicle control surface comprises: a first drive; a second drive; an output member for connecting to the vehicle control surface; first engaging means for selectively engaging and disengaging the first drive and the output member; and second engaging means for selectively engaging and disengaging the second drive and the output member, wherein the first engaging means comprises a strain wave-type gearbox.

A drive arrangement for manipulating a vehicle control surface, comprises: a first drive; a second drive; an output member for connecting to the vehicle control surface; first engaging means for selectively engaging and disengaging the first drive and the output member; and second engaging means for selectively engaging and disengaging the second drive and the output member, wherein the first engaging means comprises a first epicyclic gearbox.

A drive arrangement for manipulating a vehicle control surface, comprises: a first drive; a second drive; an output member for connecting to the vehicle control surface; first engaging means for selectively engaging and disengaging the first drive and the output member, and second engaging means for selectively engaging and disengaging the second drive and the output member.

An assembly for maintaining the pitch angle of a flight control surface for an aircraft. The assembly includes a shaft comprising a screw thread and defining a shaft axis, and a nut. The nut includes a housing, a barrel located in the housing and comprising a screw thread for engaging with the screw thread of the shaft, and at least one locking element located in the housing. The housing comprises a radially inner cam surface adjacent to the locking element. The barrel is configured to rotate relative to the housing when the screw thread of the barrel is engaged and rotating with the screw thread of the shaft so as to move the locking element along the cam surface from a disengaged position in which the locking element is spaced from the screw thread of the shaft, to an engaged position.

The invention belongs to an electric pump work control technology in the aviation field, and discloses an anti-error work control method and system for a multi-redundancy automatic electric pump, and the method comprises the steps: collecting an onboard state signal, a flight control system state signal and two paths of hydraulic pressure of each sub-hydraulic system in a hydraulic system in real time; according to the onboard state signal, a hydraulic pressure logic judgment instruction is output through the first control logic or an electric pump mistake proofing mechanism is started, and an electric pump connected to the sub hydraulic system is controlled to be closed through the electric pump mistake proofing mechanism; a flight control system state signal and two-way hydraulic pressure are obtained according to the pressure logic judgment instruction, and an electric pump connected to each hydraulic sub-system is automatically controlled through second control logic and an electric pump mistake proofing mechanism. Through the design of the first control logic, the second control logic and the electric pump mistake proofing mechanism, the working accuracy of the multi-redundancy automatic electric pump can be ensured, and the system safety is improved.

A wing for an aircraft, comprising a main wing and a high lift assembly which includes a body, and an assembly connecting the body to the main wing, so the body is movable between a retracted position and an extended position. The assembly comprises a track extending along a longitudinal axis between first and second ends, with an intermediate portion therebetween. A roller bearing ensures that the track moves along the longitudinal axis. The roller bearing comprises bearing units, each mounted to the main wing and having a bearing surface for engaging the elongate track. The units include a first and second end units and an intermediate unit wherein, relative to the longitudinal axis, the intermediate unit is positioned between the first end unit and the second end unit. The elongate track comprises a recessed portion which, in the retracted position, aligns the intermediate unit.

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.

The invention discloses a self-resetting type speed brake handle structure. The middle of the connecting rod (8) penetrates into the rod sleeve (7), the adjusting anti-rotation block (6) and the spring A (4) to be inserted into the positioning pin A (5), the upper end and the lower end of the spring A (4) abut against the adjusting anti-rotation block (6) and the lower end of the rod sleeve (7) respectively, the rolling mechanism (3) is installed at the lower end of the connecting rod (8), the gear plate (1) is installed on the lower end face of the rod sleeve (7), and the rolling mechanism (3) is clamped into a groove formed in the lower end of the gear plate (1). A positioning pin B (10) is installed at the lower end of the gear plate (1) and makes contact with the limiting mechanism (2). The spoiler has a retraction position, the connecting rod can be reset to the retraction position within the range of-5 degrees to 0 degree, the preset function is achieved, the spoiler is rapidly opened, and potential safety risks are avoided under the emergency situation.

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.

The invention relates to a dual-motor input horizontal stabilizer balancing system which comprises a channel 1 motor assembly, a channel 2 motor assembly, a differential gear train, a final-stage gear, a reduction gear, an output gear, a ball screw and a lower connector. The dual-motor input horizontal stabilizer balancing system further comprises a main control electronic component PCE which is configured to be used for generating a corresponding balancing rate instruction according to a flight control law based on external data; and the motor control unit MCU comprises an MCU1 used for a channel 1 and an MCU2 used for a channel 2, and is configured to be used for generating a corresponding control instruction according to the balancing rate instruction and providing the control instruction to the channel 1 motor assembly (1) and the channel 2 motor assembly (2).

A drive arrangement for manipulating a vehicle control surface, comprises: a first drive; a second drive; an output member for connecting to the vehicle control surface; first engaging means for selectively engaging and disengaging the first drive and the output member; and second engaging means for selectively engaging and disengaging the second drive and the output member, wherein the first drive and the second drive are mounted on bearings.

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.

High-lift system for an aircraft with at least one mechanically adjustable flap which can be actuated by means of at least one actuator (1), wherein the actuator (1) is driven by the torque of a central power control unit (PCU) of the high-lift system, and wherein the at least one actuator (1) has an integrated release clutch (5) which interrupts the load path when a permissible response torque is exceeded, characterized in that a reset means is provided to manually close the release clutch (5) after an overload situation without removing the actuator (1).

An actuator (20) arranged to be connected to a mechanical chain (13) of a mechanical system (10) for controlling the flight of a vehicle, comprising a first motor (22a), a first output arrangement (23a) driven in rotation about an axis of rotation (X1-X1) by the first motor (22a) and forming a first mechanical pathway, a second motor (22b), a second output arrangement (23b) driven in rotation about the axis of rotation (X1-X1) by the second motor (22b) and forming a second mechanical pathway, an output member (24) configured to be connected to the mechanical chain (13), a first and a second coupling / decoupling device (25a, 25b) connecting, in terms of rotation about the axis of rotation (X1-X1), respectively the first output arrangement (23a) and the second output arrangement (23b) with the output member (24) in a coupled position. The actuator (20) comprising an electronic control unit (40) configured to detect a blockage of one of the mechanical pathways of the actuator (20) and to control the torques of the motors (22a, 22b) separately.

An assembly for maintaining a pitch angle of a flight control surface of an aircraft is disclosed. The assembly includes a shaft including a thread and defining an axis of the shaft; and a nut. The nut includes: a housing; a barrel located in the housing and including threads for engaging with the threads of the shaft; and at least one locking element, the at least one locking element being located in the housing. The housing includes a radially inner cam surface adjacent the locking element. The barrel is configured to rotate relative to the housing when the threads of the barrel engage with and rotate with the threads of the shaft, thereby moving the locking element along the cam surface from a disengaged position, in which the locking element is spaced from the threads of the shaft, to an engaged position, in which the locking element is wedged between the threads of the shaft and the housing, as a result, relative rotation between the shaft and the housing is prevented.

An actuator assembly for moving a movable aerodynamic surface of an aircraft is disclosed including an input section for introducing an externally provided rotary motion into the actuator assembly, and a geared rotary actuator having at least one actuator slice having a planetary gear with a sun gear couplable to the input section, a fixed gear attachable to a fixed component of the aircraft and an output gear couplable with the movable aerodynamic surface. At least one clutch is connected to the geared rotary actuator and is configured to couple the movable surface with the input section and to selectively decouple the movable surface from the input section, such that it is freely movable.

An aircraft wing includes a main wing and a high lift assembly having a high lift body and a connection assembly movably connecting the high lift body to the main wing, such that the high lift body is movable between retracted and extended positions. The connection assembly includes an elongate track extending along a longitudinal axis between first and second ends and has an intermediate portion. The first end is mounted to the high lift body, and the second end is mounted to the main wing by a roller bearing such that the track is movable along the longitudinal axis. The roller bearing includes a roller unit mounted to the main wing and having a circumferential roller surface engaging an engagement surface of the track and an internal elastic bearing allowing the roller surface, in a cross section parallel to the roller rotation axis, to adapt its orientation or form.

A system for cross-channel communication for effectors in an electric aircraft is presented. The system may include a flight component of an electric aircraft. The system may include a plurality of effectors, wherein the plurality of effectors may be configured to control the flight component. The system may include a plurality of flight controllers communicatively connected to the plurality of effectors, wherein the plurality of flight controllers may be configured to receive an input, generate a command as a function of the input and transmit the command to the plurality of effectors. The system may include a plurality of networks communicatively connected to the plurality of effectors and the plurality of flight controllers, wherein the plurality of networks may be configured to receive the command from the plurality of flight controllers and transmit the command to the plurality of effectors.

The invention relates to the technical field of aircraft control, and discloses an aircraft and a double-shaft rotary type control surface thereof, and the double-shaft rotary type control surface comprises a circular rotating disc which is rotatably installed on the surface of the aircraft through a rotating disc rotating shaft; the control surface is rotatably mounted on the circular turntable through a control surface rotating shaft, the control surface rotating shaft is perpendicular to the turntable rotating shaft, and when the control surface is folded, the control surface is flush with the surface of the aircraft; the rotary table power mechanism is connected with the rotary table rotating shaft and used for driving the circular rotary table to rotate by 360 degrees in the surface of the aircraft around the rotary table rotating shaft; the control surface power mechanism is connected with the control surface rotating shaft and used for driving the control surface to rotate around the control surface rotating shaft; and the turntable power mechanism and the control surface power mechanism are both in signal connection with the control device. Therefore, when the rotating shaft of the control surface and the incoming flow direction form an included angle, the control surface has the control capability in three directions of pitching, yawing and rolling at the same time, the compound control capability is strong, and the stability and the control performance of the aircraft can be improved.

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.

The invention relates to an aircraft force comprehensive adjusting device, and belongs to the field of flight control. The device comprises an upper rocker arm, a left rocker arm, a right rocker arm, a middle rocker arm, a bolt and a ball bearing, the upper end A of the upper rocker arm is hinged with the control line system; the upper ends of the left rocker arm and the right rocker arm are fixedly connected with two ends of the upper rocker arm respectively; the bolt penetrates through the middle parts E of the left rocker arm, the middle rocker arm and the right rocker arm; the left rocker arm, the middle rocker arm and the right rocker arm rotate around the bolt through respective ball bearings; the upper end F of the middle rocker arm is hinged with the input end of the booster; the lower ends B of the left rocker arm, the right rocker arm and the middle rocker arm are coaxial and concentric, the lower ends B of the left rocker arm and the right rocker arm are hinged to the structure installation support, and the lower end B of the middle rocker arm is hinged to the output end of the stability augmentation steering engine.

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.

A leading edge assembly for an aircraft wing. The leading edge assembly includes a housing configured to be connectable to a fixed wing section of the wing, the housing formed with a first opening connecting an exterior of the housing with an interior of the housing, an actuating element movably connected to the housing, such that the actuating element is movable between a first position and at least one second position. The actuating element extends through the first opening and includes a first section arranged in the interior of the housing and a second section arranged at the exterior of the housing. The actuating element is configured to be connectable to a high lift device.

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.