515results about "With power amplification" patented technology

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.

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.

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 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.

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.

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.

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.

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.

The present disclosure is generally directed to a flight control system and method of flight control for an electric aircraft. The system includes a pilot input communicatively connected to an electric aircraft, wherein the pilot input is configured to receive an input datum, and plurality of flight components communicatively connected to the electric aircraft, wherein the plurality of flight components includes a plurality of control surfaces. The system also includes a flight controller, wherein the flight controller is configured to determine a phase of flight, determine a command datum to control a position of the plurality of control surfaces as a function of the input datum, and command, when the phase of flight is determined to be hover, the plurality of control surfaces using the command datum.

An inceptor apparatus for an aircraft having 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.

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 subject matter of this specification can be embodied in, among other things, a method that includes controlling, by a first fluid valve, a first fluid flow to a first fluid actuator, actuating, by the first fluid actuator, an output, controlling, by a second fluid valve, a second fluid flow to a second fluid actuator, and actuating, by the second fluid actuator, the output.

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.

A flight control system for an aircraft including a first set of primary flight control computers and a second set of secondary flight control computers, where each primary flight control computer implements a first autopilot functionality. The secondary flight control computers jointly take control of control surface actuators when the primary flight control computers fail. The flight control system further includes another computer, to which the secondary flight control computers are connected, which implements a simplified second autopilot functionality, the second autopilot functionality being able to be activated only when the second set of flight control computers has control of the control surface actuators. Thus, an autopilot functionality is widely available, while at the same time ensuring the robustness sought by the installation of the secondary flight control computers.

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

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 application belongs to the field of electromechanical control, and relates to electric pump working control technology, and discloses an electric pump working comprehensive control method and system.The working comprehensive control method comprises the following steps: collecting machine state signals, flight control system state signals and hydraulic pressure of each sub-hydraulic system in a hydraulic system in real time; controlling the opening and closing of the electric pump connected to the sub-hydraulic system according to the machine state signals through a first electric pump control logic and an error prevention working mechanism; and automatically controlling the electric pump connected to each sub-hydraulic system according to the flight control system state signals, the hydraulic pressure and the rest of the electric pump control instructions through a second electric pump control logic, the error prevention working mechanism and a staggered working mechanism.The application can ensure the accuracy of the electric pump working, avoid system damage caused by simultaneous starting of the electric pump, and improve system safety.

A flight control system may include active sidesticks with active sidestick computers which may execute flight control functions to determine control surface commands for remote electronic units. The active sidesticks may also include remote data concentrators which concentrate digital data for processing by the active sidestick computers. The flight control system may also include additional remote data concentrators and flight control computers which are located outside of the active sidesticks. The flight control computers may also generate the control surface commands for the remote electronic units.

An active side lever controller for controlling pitching and rolling motions of an aircraft, having a mechanical backup, comprising: a cross-shaped hinge assembly connected to an operating handle and implementing a two-degree-of-freedom rotation of a side lever in pitching and rolling directions through a pitching hinge shaft and a rolling hinge shaft arranged perpendicularly crossing each other; the multiple spring damping assemblies are coaxially arranged with the pitching hinge shaft and the rolling hinge shaft respectively, and can provide reverse damping force in the rotation process of the pitching hinge shaft and the rolling hinge shaft; and the multiple motor assemblies are coaxially arranged with the pitching hinge shaft and the rolling hinge shaft correspondingly, reverse driving force is provided during normal work, and rotors of the motor assemblies freely rotate in the mode of being disengaged from a transmission system after power failure. According to the active side lever controller, the problems that the fault risk of an electromechanical assembly is high in the past, and a traditional mechanical backup scheme is complex in structure and large in occupied space can be solved.

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.

Electromechanical drive system (1) to control a transmission shaft (2) connected to the mechanical operating systems of an aircraft, comprising an electric machine (3), a two-way transmission rotation (8) and a clutch assembly (9) wherein the two-way transmission rotation (8) is configured, in any direction of rotation of the transmission shaft (2), to selectively allow the passage of torque either only between the electric machine (3) and the transmission shaft (2) when the torque is supplied by the electric machine (3) to the transmission shaft (2) or between the transmission shaft (2), the clutch assembly (9) and the electric machine (3) when the torque is supplied by the transmission shaft (2) to the electric machine (3).