319results about "With power amplification" patented technology

An aircraft wing includes a fixed structure and a linear hydraulic slat actuator attached to the fixed structure. The linear hydraulic slat actuator is oriented spanwise along the aircraft wing. The aircraft wing also includes an actuator output link and a slat attached to the actuator output link. The aircraft wing also includes a bell-crank coupled to the linear hydraulic slat actuator and to the actuator output link. In response to spanwise movement of the linear hydraulic slat actuator, the bell-crank is configured to move the actuator output link in a different direction that enables the slat to move between a retracted position and an extended position.

The embodiment of the invention provides a method and device for stably controlling an airplane in the auxiliary landing process, computer equipment and a medium, and the method comprises the following steps: judging whether the airplane is in a gliding stage, if so, activating a gliding stage control law, and if not, controlling the airplane to stably control the airplane through the gliding stage control law. The aircraft is controlled through the elevator centering instruction and the horizontal stabilizer automatic trimming instruction until the aircraft quits the gliding stage, and the gliding stage control law is quitted; if the aircraft enters the pull-up stage from the glide stage, activating the leveling stage control law, generating an elevator control instruction based on the steering column displacement and the steering column displacement nonlinear gain through the leveling stage control law, generating a rudder control instruction based on the pedal displacement signal and the pedal displacement nonlinear gain, and controlling the elevator control instruction and the rudder control instruction according to the pedal displacement signal and the pedal displacement nonlinear gain. And controlling the aircraft through the elevator control instruction and the rudder control instruction until the aircraft is grounded. According to the scheme, the landing quality and safety are improved through the staged aircraft auxiliary landing control law.

The invention belongs to the technical field of amphibious aircraft flight control, and discloses an amphibious aircraft fly-by-wire flight control system and a working method, a cockpit control device is connected with and controls a normal instruction sensor and a backup instruction sensor, the normal instruction sensor transmits a signal to an actuator control electron, and the backup instruction sensor transmits a signal to the actuator control electron; the actuator control electronics and the flight control computer transmit signals mutually, the actuator control electronics and the actuator control module transmit signals mutually, and the actuator control electronics further receive signals of the external signal sensor, the direct angular rate sensor, the flap position sensor and the flight control mode switch; the backup command sensor transmits the signal to the backup flight control computer, and the backup flight control computer transmits the signal to the actuator control module; the actuator control module is connected with and controls the actuator, and the angular displacement sensor is arranged on the actuator and transmits collected signals to the actuator control module. The problem that a current fly-by-wire flight control system is not matched with amphibious aircraft cross-medium application is solved.

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.

An inceptor apparatus for an aircraft comprises a primary inceptor member provided in the form of a stick member having a grip portion, at which the stick member can be gripped by a pilot's hand, and a secondary inceptor member provided at an upper portion of the primary inceptor member and having an actuating portion, at which the secondary inceptor member can be manually actuated by a pilot's thumb. Both inceptor members have associated a respective sensor assembly which is provided to generate electronic flight control signals or commands in response to at least one of i) pivoting movements of the respective inceptor member around each of two independent maneuvering axes associated to the inceptor member, ii) forces acting on or via the respective inceptor member in pivoting directions with respect to each of the maneuvering axes, and iii) lateral flexing or bending of the respective inceptor member.

The utility model discloses an integrated aileron driving system, which belongs to the field of aircraft control systems and comprises a fixed wing, an aileron is arranged on one side of the fixed wing, the aileron is rotatably connected with the fixed wing through a first linkage shaft, the upper surface of the aileron is rigidly connected with a first speed reducing plate, and a second speed reducing plate is arranged on the other side of the fixed wing. A power rotating assembly is arranged on the upper surface of the fixed wing, the power rotating assembly is connected with the upper surface of the first air brake and used for driving the first air brake to move, fixed limiting pieces are symmetrically arranged at the two ends of the surface of the side, close to the aileron, of the fixed wing, and a second air brake is arranged between the fixed limiting pieces. The second air brake is located on the lower surface of the aileron, and the first air brake and the second air brake are connected through a transmission assembly. According to the integrated aileron driving system, through the linkage design of the aileron and the air brake, the structure is simplified, pneumatic coordination control is achieved, and the control performance at different flight speeds is improved.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft driven by an electric propulsion system and in other types of vehicles. In some embodiments, an aircraft is disclosed, the aircraft comprising: at least one electric propulsion unit; at least one sensor configured to measure at least one aircraft condition; and at least one flight control computer configured to dynamically vary at least one torque command for the at least one electric propulsion unit based at least on the at least one aircraft condition; wherein the at least one electric propulsion unit is configured to generate thrust based on at least one dynamically changed torque command.

The present invention relates to an electric flight control system (5) for controlling an aircraft (1).The flight control system (5) comprises a plurality of functional subsets (95), each having at least one dissimilar primary computer (35) and at least one dissimilar secondary computer (36), at least one functional subset (95) being a simplex / duplex subset (96) comprising at least three identical simplex primary computers (41, 42, 43) and at least one duplex secondary computer (44, 45), each duplex secondary computer (44, 45) of a simplex / duplex subset (96) comprising an independent and synchronized control computing channel (46) and a monitoring computing channel (47), each simplex primary computer (41, 42, 43) of a simplex / duplex subset (96) having a single computing channel, and in that, barring failure, the three primary computers (41, 42, 43) of a subset simplex / duplex (96) form two virtual pseudo-duplex computers. Abbreviation figure: figure 1.

An aircraft hydraulic system 100 comprises common hydraulic pressure sources 11 and 12 and pressure reduction valves 21 and 22. The common hydraulic pressure sources 11 and 12 supply a shared hydraulic pressure to a low pressure operation unit 103 that is disposed in an aircraft 101 and that operates by a relatively low operating hydraulic pressure and to a high pressure operation unit 102 that is disposed separately from the low pressure operation unit 103 in the aircraft 101 and that operates by a relatively high operating hydraulic pressure. The pressure reduction valves 21 and 22 reduce the hydraulic pressure that is supplied to the low pressure operation unit 103.

The invention relates to an aircraft system for suppressing movement of a structural flight element. The aircraft system includes a movable aircraft structure, a power source, a controller, a resistor, and a first transistor connected in series to the resistor. When the power source of the aircraft provides power, the controller controls the first transistor to prevent current from flowing through a terminal of the first transistor and through the resistor. Further, when the power supply of the aircraft does not provide power, the first transistor allows a current generated by a back electromotive force (EMF) voltage generated by movement of the movable aircraft structure to flow through a terminal of the first transistor and through the resistor.

The invention relates to the technical field of aircraft control, and discloses a control method of a rotary actuator integrated with a gear vernier encoder and related equipment, and the method comprises the steps: firstly obtaining a control instruction of a flight control system to a rudder arm, and synchronously collecting the single-circle angle values of a driving gear and a driven gear fed back by the gear vernier encoder and the real-time current of a motor; then, the actual absolute position of the rudder arm, the number of rotation turns of the motor and the actual rotation speed are calculated based on the angle value; then, a position tracking error is generated by comparing the control instruction with the actual position of the rudder arm, and the expected current is calculated according to the position tracking error and the actual rotating speed of the motor; and finally, combining the position of the motor rotor, the expected current and the real-time current to generate a motor driving signal. According to the invention, absolute position memory and full-closed-loop high-precision servo control of the aircraft control surface actuator under the condition of no standby battery are realized, and the system has the advantages of compact structure, high reliability, good dynamic response and low cost.

The invention relates to an eVTOL nacelle tilting steering engine based on high reduction ratio transmission and a working method, and relates to the field of electric vertical take-off and landing aircrafts. According to the technical scheme, the design that the double-cycloid speed reducer is combined with the parallel shaft speed reduction gear set is adopted, the wiring harness passing ability problem from the aircraft end to the nacelle is solved, the single-point failure risk that a lead screw of a traditional linear steering engine is stuck is avoided, and the safety and reliability of the steering engine are improved; moreover, depending on the dual-redundancy winding, the dual-redundancy electromagnetic brake and the dual-redundancy rotating speed feedback of the dual-redundancy servo motor, the dual-redundancy design is realized from the electrical level, and the safety performance is further enhanced. In addition, external disturbance can be reduced through high-reduction-ratio transmission, the influence of aerodynamic load disturbance on nacelle actuation is reduced, and the tilting stability is greatly improved.

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.

The application discloses a dual-redundancy steering engine control system for a UAV and a configuration method, comprising M steering engine controllers and N steering engines; the M steering engine controllers are connected with a flight pipe computer through a bus; each of the steering engine controllers has n channels and is connected with n steering engines; N>n, N>=M; the steering engines comprise dual-redundancy linear steering engines, dual-redundancy rotary steering engines, single-redundancy linear steering engines and single-redundancy rotary steering engines; wherein the dual-redundancy linear steering engines are connected with one channel of two different steering engine controllers respectively, the dual-redundancy rotary steering engines are connected with one channel of two different steering engine controllers respectively, and the single-redundancy linear steering engines and the single-redundancy rotary steering engines are connected with one channel of corresponding steering engine controllers. The dual-redundancy isosmorphism design is adopted, the two redundancies do not interact, work independently, the structure is simple, and the reliability is high.

An electrical flight control system for controlling an aircraft, the flight control system comprising at least one subassembly for acquiring a piloting command in order to acquire the piloting command from an operation of an associated piloting member that can be operated by a pilot of the aircraft along at least one piloting axis, the flight control system comprising at least three position sensors per piloting axis generating position information of the piloting member and being in communication with the at least one acquisition subassembly. Each acquisition subassembly comprises at least three simplex computers each having a single calculation channel connected to one of the at least three position sensors.

The invention belongs to the technical field of aircraft design, and discloses a high-safety double-anchoring cabin control system which comprises a left control system and a right control system, the left control system is provided with a left steering wheel, a left damper, a left longitudinal module mechanical assembly and a left longitudinal load mechanism, and the right control system is provided with a right steering wheel. A left steering wheel is connected with a left longitudinal module mechanical assembly and a left damper through a set of rocker arm structures, the left longitudinal module mechanical assembly is connected with a left longitudinal load mechanism and a right control system through one rocker arm assembly, and a right longitudinal module mechanical assembly is connected with a right longitudinal load mechanism through the other rocker arm assembly. A load pull rod is further connected between the left longitudinal load mechanism and the right longitudinal load mechanism through a rocker arm in a switching mode, the load pull rod is further connected with a balancing motor, and the load pull rod is axially connected with an anchoring pull rod. According to the invention, the fault-tolerant design of the cabin control system is innovatively realized through the designed double-anchor-point redundancy architecture.

The invention relates to a rotorcraft (100) having a main rotor (110) with at least one main rotor actuator (114), a series actuator arranged upstream of the main rotor actuator, an engine (180) for powering the main rotor, and a flight control system (160) for controlling flight operations in an AEO mode or in an AFR mode, the flight control system comprising a control logic (165) for pre-conditioning the series actuator upon entering the AFR mode to enable rotorcraft control in case of an IFSD, wherein the pre-conditioning of the series actuator comprises shifting a reduced nominal range of series actuator travel authorized for the series actuator in the AEO mode into a reduced shifted range of series actuator travel authorized for the series actuator in the AFR mode to enable an authorization of an increased range of travel in case of an IFSD to provide sufficient control authority for recovery maneuver.

The present invention relates to a method for measuring a damping function, preferably of a damper of an electromechanical actuator or drive system, and to an actuator or drive system with a control unit designed to carry out a method according to the invention.

The flight control system (10) includes at least one aircraft flight control computer (16) intended to control an aircraft elevator (5). The at least one flight control computer is configured to implement the following steps repeatedly during an aircraft acceleration phase while taxiing on the ground in preparation for takeoff: - determination (30) of an elevator deflection angle (5), corresponding to a load (Fz) to be applied to a nose landing gear of the aircraft such that said load is within a predetermined load range ([Fzmin; Fzmax]), - control (40) of an elevator actuator (18) so as to apply said elevator deflection angle to the elevator.

The present invention relates to a method for detecting play in an aerodynamic control system (12), which comprises a fixed part (18), a movable control surface (20), and two servo control blocks (A, B) for generating a displacement of the control surface. The method comprises a play detection sequence including: - a control phase comprising: * asymmetric control of the servo control blocks (A, B); * acquisition of an evolution, during said asymmetric control, of a force variable representative of an opposing force exerted in one of the servo control blocks (A, B) during the asymmetric control; - a play detection phase (106) comprising the verification (114) of a play detection condition as a function of said acquired evolution.

The invention provides a backup control apparatus and a control method. The backup control device comprises a first activation module, a second activation module and an instruction calculation module. The first activation module and the second activation module generate a first activation signal and a second activation signal indicating activation of the backup control device, respectively, in response to the aircraft speed being higher than a first threshold and the number of host computer failures being higher than a second threshold. In response to the first activation signal and the second activation signal, the instruction calculation module is configured to generate a backup control command to control a control surface actuator to move.

The application discloses a distributed large eVTOL electric governor control method, relates to the technical field of aircraft control systems, and discloses a distributed large eVTOL electric governor control method, which realizes rapid detection and isolation of faults, maintains system stability through accurate dynamic thrust compensation, can improve flight safety, reduce fault propagation risk, and ensure that the system can quickly recover to stability when hardware faults occur.

The present disclosure relates to spoiler actuation systems and methods for aircraft. An aircraft includes a wing, a spoiler, a spoiler movable between a cruise position and an up position and between the cruise position and a droop position, and a spoiler actuation system coupled to a hydraulic system of the aircraft, the spoiler actuation system including a first piston and a second piston, a rack connected between the first piston and the second piston (the rack movable between a first position and a second position), a pinion coupled to the rack (the pinion rotating between a third position and a fourth position as the rack moves between the first position and the second position), a first crank arm coupled to the pinion (the first crank arm rotating with the pinion between the third position and the fourth position), and a second crank arm coupled to the first crank arm and the spoiler (the second crank arm moving the spoiler between the cruise position and the up position as the first crank arm rotates between the third position and the fourth position).

An aircraft includes actuators coupled to a lift device. The aircraft also includes a lift device control interface disposed on a flight deck of the aircraft. The lift device control interface is configured to receive a pilot input. The aircraft further includes a flight control system configured to generate a command corresponding to a target lift device position. The flight control system is also configured to send the command to the lift device control interface to cause the lift device control interface to indicate the target lift device position. The flight control system is further configured to send a control signal to the actuators to set the lift device to the target lift device position.

The present invention relates to an electric flight control system (5, 15, 25, 35, 45) for controlling an aircraft (1), the flight control system (5, 15, 25, 35, 45) comprising at least one flight control input subset (2) for acquiring said flight control input from a maneuver of an associated control element (20) maneuverable by a pilot of said aircraft (1) along at least one control axis, said flight control system (5, 15, 25, 35, 45) comprising at least three position sensors per control axis generating position information for said control element (20) and communicating with said at least one input subset (2). Each input subset (2) comprises at least three simplex computers, each having a single processing channel connected to one of said at least three position sensors.