739results about "With power amplification" patented technology

The invention discloses an aircraft control surface oscillation fault monitoring method based on multi-scale wavelet transformation, and the method comprises the steps: carrying out the calculation through a PID (Proportion Integration Differentiation) control law through an autopilot according to an input flight path command, and obtaining a control surface control instruction deltadepth; a control surface deflection prediction value delta est is generated based on the nonlinear model and the control surface control instruction, the control surface deflection prediction value and an actual deflection value delta mea collected by the sensor module are subtracted to generate a residual signal Rn, and the residual signal is decomposed into a plurality of sub-band signals through wavelet decomposition; wavelet energy calculation is carried out on the sub-bands possibly containing the control surface oscillation fault signals, whether a fault occurs or not is judged according to the magnitude of an energy comparison threshold value, and the judgment threshold value is determined through Monte Carlo simulation. According to the method, instruction signals are separated by generating residual errors, wavelet transform processing is carried out on residual error signals, interference of noise and irrelevant frequency signals on monitoring performance is reduced, meanwhile, small-amplitude oscillation signals can be effectively monitored according to sub-band wavelet energy serving as a fault criterion, and the monitoring coverage rate of oscillation faults is remarkably increased.

Heat exchangers for cryogenic fuel distribution systems and related methods are disclosed herein. An example aircraft disclosed herein includes a cryogenic fuel system including a cryogenic fuel, a hydraulic system including a hydraulic fluid, a gas turbine engine fluidly coupled to the cryogenic fuel system, and a heat exchanger fluidly coupled to the cryogenic fuel system, the heat exchanger is used to exchange heat between the cryogenic fuel and at least one of (1) hydraulic fluid or (2) working fluid thermally coupled to the hydraulic system.

The invention belongs to the field of airborne electromechanics, and discloses a hydraulic mechanism heterogeneous control method and system.The method comprises the steps that hydraulic systems are divided into two groups, and on-board state signals, flight control system state signals and first hydraulic pressure and second hydraulic pressure of sub hydraulic systems are collected; giving an on-board state threshold value, a low-pressure threshold value, a hydraulic safety threshold value and a sub-hydraulic system relation table; starting and stopping automatic control is conducted on the electric pumps in the first group through the established first control logic architecture according to the data; and the electric pumps in the second group are automatically controlled to start and stop by adopting the constructed second control logic architecture through the data. The system comprises a grouping module, a control logic construction module and two control units, wherein each control unit comprises an acquisition module, a data scheduling module, a staggered starting module, an anti-overvoltage protection module and an anti-error working module. All the electric pumps are controlled through two heterogeneous logic architectures, and the control precision and safety are improved.

A system configuration significantly reduces the number of interconnections needed between an electromechanical actuator (EMA) and an EMA controller (EMAC) while maintaining redundancy and other safety aspects needed by the target application without adding second or third independent redundant position sensors to the EMA for use with backup channel of the controller. The system configuration extends the sensor architecture used by COM-COM controllers with an additional sensor channel used by a backup controller without the need for adding a separate physical position sensor for the backup controller, which saves implementation effort, space, and weight that would otherwise be needed for integration of additional redundant sensors into the EMA.

Systems and methods for controlling an electric vertical take-off and landing (eVTOL) aircraft are disclosed. In one embodiment, a system includes a processor, a first manipulator communicatively coupled with the processor, the first manipulator configured to accept longitudinal and lateral linear movements as manual inputs and provide corresponding signals to the processor, and a second manipulator communicatively coupled with the processor, the second manipulator is configured to accept longitudinal and lateral linear movements as manual inputs and provide corresponding signals to the processor, wherein the processor is configured to control the heading of the aircraft using the signals received from the second manipulator corresponding to the lateral linear movement of the second manipulator. Some embodiments may additionally include at least one sensor and a thumb rocker for each manipulator.

The invention discloses an integrated linear steering engine, which belongs to the technical field of linear steering engines, and comprises a shell, and a motor assembly, a transmission assembly and a sensor which are arranged in the shell, the transmission assembly comprises a multi-stage gear pair, a lead screw pair and an output shaft assembly. The lead screw pair output shaft assembly and the motor assembly are arranged in parallel. The multi-stage gear pair comprises an input gear, a transfer gear and an output gear which are sequentially arranged in a meshed mode, and the input gear is connected with the output end of the motor assembly. The lead screw pair output shaft assembly comprises a lead screw and a nut output shaft. One end of the lead screw is connected with the output gear, the other end of the lead screw is connected with the nut output shaft, and the nut output shaft extends out of the shell to achieve linear output. The sensor comprises a rotary transformer sensor, and the rotary transformer sensor is arranged at the end, away from the output end, of the motor assembly. The size and weight of the steering engine can be greatly reduced, and the steering engine has the advantages of being large in output torque and high in transmission precision.

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 an actuator jamming monitoring scheme. The actuator jamming monitoring scheme comprises the steps that a remote electronic unit REU transmits a current signal to an electro-hydraulic servo valve EHSV; the REU receives a displacement instruction signal from the computer; an actuator displacement signal is collected from an LVDT sensor; acquiring an actuator extension cavity pressure signal from a first pressure sensor; acquiring an actuator retraction cavity pressure signal from a second pressure sensor; and the data of the REU, the LVDT sensor, the first pressure sensor and the second pressure sensor are summarized and analyzed, and the analyzed result is compared with a corresponding threshold to judge whether to trigger jamming monitoring alarm.

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 invention discloses a novel bird-imitating empennage multi-degree-of-freedom platform, and relates to the technical field of flying robots, the novel bird-imitating empennage multi-degree-of-freedom platform comprises a first driving structure, a support, a second driving structure, an empennage driving structure and an empennage assembly, the second driving structure and the empennage driving structure are both arranged on the support, the first driving structure can drive the empennage assembly to rotate in the heading direction, and the second driving structure can drive the empennage assembly to rotate in the pitching direction and the rolling direction. And the empennage driving structure can control the area change of the empennage assembly. According to the novel bird-imitating empennage multi-degree-of-freedom platform, control over the heading direction, the rolling direction and the pitching direction can be achieved, and the area of the empennage assembly can be changed.

A flight control system, including a flight control computer (FCC) configured for providing a first and second flight function, and sending, to effectors of an aircraft, first command signals based on the first and second flight functions, and a second computer configured for providing one or more autonomy functions associated with operation of the aircraft, and that are different from the first and second flight functions, monitoring operation of the FCC, determining whether the FCC has failed, and providing a fail-safe mode for at least the first flight function in response to determining that the FCC has failed. Providing the fail-safe mode includes providing one or more critical flight functions that includes the first flight function, and sending, to the effectors, second command signals that replace a first command signal associated with the first flight function and that are based on the critical flight functions.

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.

This invention discloses a servo-elastic stabilization method based on multi-sensor information fusion, belonging to the field of flight control. Specifically, it involves: extracting and modeling the main load-bearing structure of the aircraft; calculating the modal frequencies, mode shapes, and mode shape slopes of the aircraft under unconstrained conditions; then combining the aircraft, a first sensor, servo motors, and the control system to form a closed-loop servo-elastic system; calculating the open-loop system transfer function from the control input to the feedback input of the control system, drawing an open-loop Bode plot, and determining the dangerous frequencies and corresponding dangerous natural modes; based on this, placing a second angular velocity sensor on the aircraft; then calculating and correcting the weighting coefficients of the two angular velocity sensor signals, and introducing the weighted signal into the control system instead of the original signal; comparing the system characteristics to ensure the servo-elastic stability of the aircraft without affecting the rigid body phase and amplitude.

The invention provides a fly-by-wire aircraft transverse soft control system and method. The fly-by-wire aircraft transverse soft control system comprises a left steering wheel and a right steering wheel. The left steering wheel and the right steering wheel are respectively connected with a left steering wheel force sensor and a right steering wheel force sensor through a drum wheel, and the two drum wheels are respectively connected with one end of a left control steel cable and one end of a right control steel cable. The other end of the left control steel cable and the other end of the right control steel cable sequentially pass through corresponding decoupling pulleys and reversing pulleys and then are connected with fan-shaped wheels of the left load module and the right load module. The left load module and the right load module are respectively connected with a left damper, a right damper, a left command position sensor, a right command position sensor, a left balancing mechanism, a right balancing mechanism and output rocker arms, the fan-shaped wheels are connected with the corresponding output rocker arms, and the two output rocker arms are connected through a release device; and the left balancing mechanism and the right balancing mechanism are synchronously connected. Layout is flexible, the requirement for the aircraft structure is low, instruction consistency is good, and integration is high.

A method is provided for determining a gas charge pressure PcalcGC of a hydraulic accumulator within a hydraulic fluid system. The method includes receiving pressure decay data representing the pressure decay within a self-pressurizing hydraulic reservoir of the hydraulic fluid system as the self-pressurizing hydraulic reservoir depressurizes from an operating pressure of the self-pressurizing hydraulic reservoir towards atmospheric pressure. The method also includes calculating the gas charge pressure PcalcGC of the hydraulic accumulator using the received pressure decay data of the self-pressurizing hydraulic reservoir.

The invention relates to the technical field of aircraft attitude control, and discloses an aircraft attitude auxiliary braking control system based on a high-speed permanent magnet synchronous motor, which comprises the following steps of: defining an aircraft attitude braking instruction as a target energy dissipation power instruction, and acquiring a motor operation angular speed in real time; a target braking current instruction is dynamically calculated based on the energy dissipation power instruction and the angular velocity, and a motor is controlled to execute braking. Through energy management mode reconstruction, a braking system becomes a self-adaptive energy dissipater, the current is automatically adjusted under turbulence disturbance to maintain constant power, conversion from force tracking control to energy flow management is achieved, and the braking efficiency is improved. And meanwhile, by controlling an input linear mapping and global working condition switching mechanism, the control consistency at different speeds is ensured, and the attitude stability of the aircraft is improved.

A hydraulic energy retention system includes a body with a piston separating a hydraulic fluid from a pressurized gas, a discharge port in fluid communication with the hydraulic fluid, and a stop operable to an engaged position to block movement of the piston past the stop and toward the discharge port. The hydraulic energy retention system may be implemented with a hydraulic fluid reservoir, such as a bootstrap reservoir.

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.

A digital surface position sensor includes a position sensor, an adaptable hardware interface, and a processing circuit. The position sensor is adapted to be coupled to an aircraft flight control surface and is configured to sense a position of the aircraft flight control surface and supply a position signal representative thereof. The adaptable hardware input supplies an identification signal that identifies the aircraft flight control surface to which the position sensor is coupled. The processing circuit is coupled to receive the position signal and the identification signal. The processing circuit is configured, upon receipt of the signals, to process the position signal and the identification signal and generate (i) a first digital position signal representative of the position of, and the identification of, the aircraft flight control surface and (ii) and independent second digital signal representative of the position of, and the identification of, the aircraft flight control surface.

The invention relates to a flight control system which comprises three flight control computers, and each flight control computer comprises an I / O board used for receiving input signals of one path of a cockpit control device signal and a sensor signal of a cross-linked interface system and outputting the input signals of the other path of the cockpit control device signal and the sensor signal of the cross-linked interface system; after the input signals are simply processed, the processed input signals are forwarded to the computing board, the backup board and other flight control computers, meanwhile, remaining cockpit control device signals and interface signals sent by other computers are received, and control instructions are output to actuators of corresponding control surfaces; the three calculation boards are used for calculating a corresponding control instruction based on a control law according to the received input signal and forwarding the control instruction to the I / O board; the backup board is used for entering a pre-position state when the computing board of the flight control computer fails; wherein when the backup boards of all the flight control computers enter the pre-position state, the flight control system enters the backup mode, and the backup boards calculate backup control instructions and transmit the backup control instructions to the actuators of the corresponding control surfaces.

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.

An aircraft wing assembly includes a wing box including an upper surface, a lower surface, and a leading edge structure connecting the upper surface to the lower surface and secured to the upper surface at a connection point. The aircraft wing assembly includes a drive mechanism secured to the wing box and a high lift device mounted to the drive mechanism. The drive mechanism is configured to move the high lift device between a retracted position and an extended position in use, and wherein the high lift device is configured to overlap the connection point when in the retracted position.

A control surface system for a body located in an air and / or space vehicle. The control surface system having at least one control surface which is located on the body and which, by moving relatively to the body, enables air flow to be controlled and the air vehicle to maneuver accordingly. The control surface system having at least one actuator which is made of an electroactive polymer material which is located on the body and which changes its form depending on electrical energy and based thereon, triggers the control surface.

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 control system for rotating an airfoil along one axis that is offset from a perpendicular axis that varies the angle of attack (AOA), dihedral, and center of lift based on the position of the attachment to precisely maneuver an aircraft. This technique can eliminate the need for traditional control surfaces such as ailerons, elevators, and rudders. allowing for an aerodynamically advantaged shape.

A flight control system, including a flight control computer (FCC) configured for providing a first and second flight function, and sending, to effectors of an aircraft, first command signals based on the first and second flight functions, and a second computer configured for providing one or more autonomy functions associated with operation of the aircraft, and that are different from the first and second flight functions, monitoring operation of the FCC, determining whether the FCC has failed, and providing a fail-safe mode for at least the first flight function in response to determining that the FCC has failed. Providing the fail-safe mode includes providing one or more critical flight functions that includes the first flight function, and sending, to the effectors, second command signals that replace a first command signal associated with the first flight function and that are based on the critical flight functions.

The invention relates to the technical field of airborne rocker arm assembly control, in particular to an aircraft control surface deflection angle limiting mechanism which comprises a fixed part and a movable part, an overturning groove is formed in the fixed part, the movable part and the inner wall of the overturning groove are rotationally installed, a fixed plate is fixedly installed on the inner wall of the overturning groove, and the movable part is fixedly installed on the fixed plate. A rotating plate is fixedly installed on the movable part, a connecting shaft is rotatably arranged on the rotating plate in a penetrating mode, the connecting shaft rotatably penetrates through the fixed plate and is rotatably connected with the inner wall of the overturning groove, a fan-shaped limiting ring is fixedly installed outside the connecting shaft, and an adjusting mechanism used for limiting the maximum deflection angle of the fan-shaped limiting ring is arranged in the fixed plate. A two-way pipe is installed on the movable component, and a limiting and compensating mechanism used for limiting and compensating the deflection angle of the movable component is arranged in the two-way pipe. Compared with the prior art, the application range is wide, the lifting height and the steering angle of the aircraft can be compensated, and the influence of external wind power is reduced.