314results about "Actuated personally" patented technology

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 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.

An information prompting method and an information prompting device are used for prompting the state of a side stick during aircraft driving, and the method comprises the following steps: based on the operation of a plurality of drivers, judging whether one of the drivers obtains the priority control authority of the side stick or not; when it is judged that one of the drivers obtains the priority control authority of the side rod, the authority identifier corresponding to the driver obtaining the priority control authority is lightened; and based on the side rod state of the driver who does not acquire the priority control authority, judging whether to lighten the prohibition identifier corresponding to the driver who does not acquire the priority control authority. As a result, the side lever state can be presented intuitively and concisely.

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.

An apparatus for adjusting aircraft inceptor trim includes an inceptor centring device mechanically coupled to a control shaft that rotates in response to inceptor inputs. The inceptor centring device includes an inceptor centring spring preloaded with a preload force to provide a first biasing force to bias the control shaft to an inceptor neutral position. The apparatus has a trim mechanism, mechanically coupled to the inceptor centring device, and includes a trim locking device having a locked state and an unlocked state. A resistance force of the trim mechanism with the trim locking device in the unlocked state is less than the preload force, such that when the trim mechanism is in the unlocked state, actuation of the inceptor adjusts the position of the inceptor centring device to alter the degree of trim.

A 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.

A side lever manipulating device and a manipulating method thereof, the side lever manipulating device for controlling pitching motion and rolling motion of an aircraft, comprising: a manipulating handle which is a main body part of the side lever manipulating device and is held by a pilot; a pitch control unit which is provided to the control handle, and which controls the pitch movement of the aircraft by means of the pitch control unit; and a rolling control unit which is provided on the control handle and controls the rolling motion of the aircraft by means of the rolling control unit, the pitch control unit and the rolling control unit being independent from each other and not interfering with each other during operation. The side rod control method is used for controlling the side rod control device. By means of the side rod control device and the control method thereof, pitching and rolling control can be decoupled, independent pitching and rolling control is achieved, and therefore the control problem of cross coupling of the side rod is solved.

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 discloses a control device, a cockpit and a vertical take-off and landing aircraft, and relates to the technical field of vertical take-off and landing aircrafts, the control device comprises a main control device, a backup control device and a device switching module; the main control equipment and the backup control equipment are arranged corresponding to a driver's seat of the vertical take-off and landing aircraft, are relatively independently arranged, and can be independently used for controlling the flight of the vertical take-off and landing aircraft; the equipment switching module is in communication connection with the main control equipment and the backup control equipment, and is at least used for switching the control authority of the vertical take-off and landing aircraft from the main control equipment to the backup control equipment; according to the technical scheme provided by the utility model, the safety of the whole flight can be improved.

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.

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.

The utility model provides a general comprehensive adjustor for helicopter pull rods, which relates to the technical field of helicopters and comprises a guide block seat and a sliding guide block, a guide block groove is arranged on the guide block seat, the sliding guide block and the guide block groove are slidably mounted, a first screw is mounted at one end of the sliding guide block in a threaded manner, and a second screw is mounted at the other end of the sliding guide block. A second screw rod is mounted at one end of the guide block seat in a threaded manner; the first screw and the second screw are both provided with discs in a threaded mode. According to the angle of the connecting lugs at the upper ends of the pull rods, the discs can be rotated, the angle of the round rods can be changed, the positions of the sliding guide blocks in the guide block grooves can be movably adjusted according to the length needed by the pull rods, and the sliding guide blocks are fixed through the stop nails, so that the distance between the two discs is matched with the distance between the connecting lugs at the ends at the two ends of the pull rods; when a new pull rod is installed, the end connecting lugs at the two ends are respectively sleeved on the round rods on the discs at the two ends, so that whether the length of the pull rod meets the requirement or not can be conveniently and quickly measured, and the field emergency repair time is shortened.

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.

In an aspect a system for fly-by-wire flight control configured for use in electric aircraft includes at least a sensor, wherein the sensor is communicatively connected a pilot control and configured to detect a pilot input from the pilot control and generate, as a function of the pilot input, command datum. A system includes a flight controller, the flight controller including a computing device and configured to perform a voting algorithm, wherein performing the voting algorithm includes determining that the sensor is an allowed sensor, wherein determining that the sensor is an allowed sensor includes determining that the command datum is an active datum, determining the command datum is an admissible datum, generating, as a function of the command datum and the allowed sensor, a control surface datum wherein the control surface datum is correlated to the pilot input.

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.