190results about "Actuated personally" patented technology

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.

A craft is provided with a lever that can control the power / speed of the craft and semi-automatically transition the craft in various modes of operation. This makes control of the craft somewhat similar to control of a boat, thereby bringing a maritime flight control experience to a wing-in-ground effect vehicles, hydrofoiling vessels, seagliders, and seaplanes. Further, various audio, visual, and / or haptic devices in the cockpit can provide alerts / feedback to the operator on the use of the lever and mode transitions.

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 invention discloses a flight control device, an aircraft control method and a vertical take-off and landing aircraft, and relates to the technical field of aircraft control. A main manipulator configured to receive manipulation information input by a pilot; the backup manipulator and the main manipulator are independently arranged relatively, and the backup manipulator is configured to receive manipulation information input by a pilot; the control system is configured to identify a target manipulator from the main manipulator and the backup manipulator, generate a vector control instruction mapped with the manipulation information by using a signal corresponding to the manipulation information received from the target manipulator and combining a preset manipulation mapping relation, and control the target manipulator according to the vector control instruction. And controlling the aircraft to perform vector motion according to the vector control instruction. According to the scheme, the complexity of aircraft control can be effectively reduced, and the safety of aircraft control is improved.

The utility model belongs to the technical field of fixed wings, and provides a manned fixed wing multi-mode control device which comprises an operation table. The mechanical control assembly is installed in the operation table, the mechanical control assembly is used for manually driving the fixed wings, the mechanical control assembly comprises a support installed in the operation table, a first linkage rod is rotationally installed at the upper end of the support, a first transmission connecting rod is rotationally installed at the front end of the first linkage rod, and a first rocker is installed at the upper end of the first linkage rod; and the intelligent control assembly is installed in the operation table, the intelligent control assembly is used for intelligently controlling the fixed wings, and the intelligent control assembly comprises a supporting frame installed in the operation table. According to the utility model, the mechanical control assembly, the intelligent control assembly, the separation assembly and the limiting assembly are integrated, and through cooperative operation of the components, diversified control modes of the fixed-wing aircraft are realized, including mechanical control, semi-automatic control and full-automatic control.

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.

The invention relates to a pitch angle adjusting mechanism, a control system and a hovercar, the pitch angle adjusting mechanism is used for being installed between a steering wheel and a seat body part, and the pitch angle adjusting mechanism is used for adjusting the pitch angle of the steering wheel so that the steering wheel can be switched between a land running state and a flying state. Only one control system needs to be configured to meet the control requirements of the hovercar during land running and flying at the same time, the universality of the control system is improved, the complex structure of the whole vehicle and the reduction of the use number of the control system are avoided, the installation and arrangement space and weight can be saved, the lightweight and miniaturized design of the hovercar is facilitated, and the cost is reduced. The use performance of the hovercar is ensured.

An electric flight control system comprising an actuation subassembly, having three external duplex computers for controlling an electric actuator. The electric actuator has three stators electrically connected to three respective switching devices, the switching devices being electrically connected to a same internal simplex computer of the electric actuator, each switching device being connected to a respective external duplex computers, each switching device being configured to connect the internal simplex computer or respective external duplex computer to the associated stator, the internal simplex computer being self-monitored and communicating with the three external duplex computers.

An aircraft rudder control system having one or more rudder pedal adjustment systems, each including a rudder pedal actuatable to control movement of a rudder, and further including a support frame component, a rudder pedal arm, a pin, a pin receiving component, a first lever, a second lever, and / or a solenoid. The rudder pedal arm is rotatably attached to the support frame component and the rudder pedal is attached to the rudder arm. While the pin is received in the pin receiving component, the rudder pedal arm is prevented from rotating. When the pin is withdrawn from the pin receiving component, the rudder pedal arm is freed to be rotatably repositioned by user foot actuation. The first lever can manually actuate the pin into and out of the pin receiving component, and the second lever is operated by the solenoid, actuating the pin when the solenoid is activated and deactivated.

An eVtol aircraft manages different command inputs from various sources such as multiple inceptor types from the pilot (physical or digital), automatic command from autopilot or autonomous system, voice command, and / or remote piloting from an external source linked to the system. A Flight Control Command Source Management (FCCSM) can manage multiple dissimilar sources of commands for the Flight Control System automatically and based on pilot specified selections and / or a priority list.

The invention relates to a joystick having first and second redundant return mechanisms (110, 112). The first return mechanism comprises a first elastically deformable assembly (120) inserted between a first base (126) rigidly connected to a stationary frame (6) and a first rim (40) rigidly connected to the stick. The second return mechanism (112) comprises: - a contact face rigidly connected to the stationary frame; - a shuttle (132) free to slide along the stick, this shuttle comprising, on one side, a pad (142) bearing on the contact face and, on the opposite side, a second base; and - a second elastically deformable assembly (130) inserted between the second base and a second rim (70) rigidly connected to the stick such that the second elastically deformable assembly continuously urges the pad against the contact face.

The invention discloses a wind pendulum and propeller device of a helicopter, belongs to the technical field of helicopter instrument production, and aims to solve the technical problems of torsion generated by rotation of a propeller, turning and flight power of the helicopter, flight control and the like by utilizing the wind pendulum and propeller device. According to the technical scheme, a propeller rotates anticlockwise to generate lift force, and the attack angle of blades is fixed; two groups of wind pendulums are mounted on a wind pendulum mounting frame below the propeller, are bilaterally symmetrical, are perpendicular to the longitudinal axis of the fuselage and droop naturally; one end of the long connecting rod is connected with the piston rod; the hydraulic system is composed of two oil cylinders and an oil pipe, and a tail end piston rod is connected with a control handle rod. One handle rod controls the deflection direction and angle of one wind swing set through the hydraulic system and the long connecting rod. The device is suitable for flight power and flight control devices of helicopters and unmanned aerial vehicles.

A pedal cartridge for an aircraft. The pedal cartridge is connectable to a pedal mechanism and includes: a pedal. The pedal includes a pedal axle receiving portion configured to receive a pedal axle of the pedal mechanism such that the pedal is rotatably connectable to the pedal axle. The mechanism also includes a first rotation sensor component, the first rotation sensor component configured to interact with a second rotation sensor component of the pedal mechanism to permit rotation of the pedal about the pedal axle to be detected.

The invention relates to a force application device for an aircraft control stick, in which the computer is configured: a) for an angle (ANG1) smaller than or equal to a prescribed virtual stop threshold (S1, S2), to control the motor so that it applies a resistive force (EFF1) to the lever, according to the angle (ANG1), and to deactivate the brake; b) for an angle (ANG1) which is increasing beyond the prescribed threshold (S1, S2), to control the motor and activate the brake so that they apply a resistive force (EFF2) to the lever that is higher than the value (V1); c for an angle (ANG1) which is decreasing while being higher than the prescribed virtual stop threshold (S1, S2), to control the motor so that it applies a return force (EFF3) to the lever according to the angle (ANG1) and to deactivate the brake.

An aircraft rudder pedal includes a frame, a pair of pedals, an output shaft and a mechanical kinematic chain for transmitting to the output shaft a displacement of at least one of the pedals relative to the frame. The mechanical kinematic chain comprises a central transmission part able to be rotated relative to the frame by a displacement of one of the pedals relative to the frame. The mechanical kinematic chain also includes a transmission mechanism joining the central transmission part to the output shaft. The transmission mechanism includes a homokinetic association of two universal joints between said central transmission part and the output shaft.

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.

An improved inceptor includes a grip configured to move a shaft and a mechanical linkage coupling the shaft to a sensor. The inceptor also includes a redundant mechanical system that includes a second shaft movable by the grip and a second mechanical linkage coupling the second shaft to the sensor. Accordingly, movement of the grip is relayed to the sensor by two discrete mechanical systems. In examples, redundant mechanical systems can be provided for any of three or more rotational directions, e.g., roll, pitch, yaw, and / or the like.

Pitch and roll input control devices for aircraft are described that utilize a multi-bar linkage mechanism to translate forward or aft movement of the steering wheel into flight controls in a compact volume. The devices replicate the motion of a large mechanism but use a more compact mechanism that does not require volume below the floor and eliminates the linkage that normally would reside between a pilot's legs.

The present invention relates to systems and methods for autonomous control of overrunning devices. In an example, a system is described that includes a control lever for controlling operation of a device, a first actuator coupled to the control lever via a rod, and a resettable frangible link coupling a second actuator to the control lever via the rod. Based on a force applied to the rod by the first actuator, the resettable frangible link enables decoupling of the second actuator from the control lever.

The invention belongs to the field of application and development of aircraft accelerator control systems, and relates to an interlocking device for forward and reverse thrust hubs of an aircraft accelerator platform. According to the interlocking device for the forward and reverse pushing hubs of the accelerator platform, the backpack type forward and reverse pushing hubs are installed on the accelerator platform shell, the interlocking assembly is installed on the forward pushing hubs, and a four-connecting-rod mechanism in the interlocking assembly moves at different positions to drive a clamping claw to achieve interlocking of the forward and reverse pushing hubs; a thrust rod interlocking function is provided for a pilot in a manual mode and an automatic mode, the pilot is prevented from touching a throttle lever by mistake, and safety is improved.

A control device comprising a lever provided with a magnetic dipole and able to move, at least in rotation, relative to a support about an axis. The control device includes a measuring device configured to measure a magnetic field generated by the magnetic dipole, and a monitoring system for monitoring the integrity of the control device. The monitoring system is configured to signal an anomaly using an alarm when a first norm of a first vector difference between the current magnetic field and a first reference magnetic field relating to an estimated orientation of the lever, is greater than a first magnetic field threshold for a duration greater than a duration threshold.

An aircraft includes a first inceptor assembly. The first inceptor assembly has a first base and a first control stick extending away from the first base, the first control stick defining a first longitudinal axis that extends through the first control stick, the first longitudinal axis forming a non-zero angle with a vertical direction. The aircraft includes a second inceptor assembly, the second inceptor assembly having a second base and a second control stick extending away from the second base, the second control stick defining a second longitudinal axis that extends through the second control stick and through the second base, the second longitudinal axis being generally aligned with a vertical direction.

The present invention relates to an electric flight control system (10) comprising an actuation subset (C), having three external duplex computers (21, 24, 27) for controlling an electric actuator (40). The electric actuator (40) has three stators (42, 43, 44) electrically connected to three respective toggle devices (50), the toggle devices (50) being electrically connected to the same internal simplex computer (80) of the electric actuator (40), each toggle device (50) being connected to one of the respective external duplex computers (21, 24, 27), each toggle device (50) being configured to connect to the associated stator (42, 43, 44) the respective internal simplex computer (80) or external duplex computer (21, 24, 27), the internal simplex computer (80) being self-monitored and communicating with the three external duplex computers (21, 24, 27).