157results about "Conjoint controls" patented technology

An unmanned air vehicle is provided, which includes an airframe and a parallel hybrid-electric propulsion system mounted on the airframe. The parallel hybrid-electric propulsion system includes an internal combustion engine and an electric motor. A hybrid controller is configured to control both the internal combustion engine and the electric motor. A propeller is connected to a mechanical link. The mechanical link couples the internal combustion engine and the electric motor to the propeller to drive the propeller. An alternate unmanned air vehicle includes a second propeller driven by the electric motor. In this alternate unmanned air vehicle, the internal combustion engine is decoupled from the electric motor.

The invention belongs to the field of unmanned aerial vehicle technology, and more specifically relates to a multi-unmanned aerial vehicle route planning method which can be used for multi-unmanned aerial vehicle target tracking, route optimization, cooperative management, cooperative flight, and task distribution. The multi-unmanned aerial vehicle route planning method comprises following steps: detailed information of route planning task is initialized, and coordinate discrete transformation is performed; parameters of the bee colony collaborative foraging algorithm is initialized; route cost of each unmanned aerial vehicle is calculated; current route cost is calculated according to a current position by each unmanned aerial vehicle; guide aerial vehicles are selected by followed aerial vehicles, wherein one followed aerial vehicle is recruited by each guide aerial vehicle; the current route is abandoned and a new route is searched; parameters of optimized route are saved, and an optimal value is calculated; inspection is performed so as to determine whether an upper limit of iteration number is reached. The multi-unmanned aerial vehicle route planning method is capable of resolving the initialization sensitive problem of a traditional intelligent optimization route planning method, and can be used for improvement of accuracy of the algorithm so as to obtain an optimal solution or a second best solution sufficiently close to the optimal solution, and improve stability of route planning and search efficiency.

The invention discloses a flight control method for a multiple-rotor-wing aircraft. The multiple-rotor-wing aircraft comprises multiple rotor wing mechanisms; the multiple rotor wing mechanisms provide power for the multiple-rotor-wing aircraft. The flight control method comprises the following steps: detecting whether the output power of the multiple rotor wing mechanisms is abnormal or not; if detecting that the output power of partial rotor wing mechanisms is abnormal, adjusting the output power of other rotor wing mechanisms, so as to enable the multiple-rotor-wing aircraft to keep balanced. The invention further discloses a flight control system for the multiple-rotor-wing aircraft. Through the above mode, additional power can be saved to stabilize the attitude, therefore, that the multiple-rotor-wing aircraft lands in a loss minimization manner is ensured, and the damage of the parts of the aircraft and damage to vehicles and people on the ground, which are caused by direct crash of the aircraft are avoided.

A system is provided for controlling the speed of an aircraft during a deceleration segment between a first state associated with a first speed and a first time and a second state associated with a second speed and a second time. The system includes a navigation system configured to determine the first state; a guidance system configured to determine the second state; and an active deceleration system coupled to the navigation system and the guidance system and configured to construct the deceleration segment between the first state and the second state with at least one intermediate speed between the first speed and the second speed.

A cockpit rudder control mechanism is described for accepting aircraft control input utilizing rudder pedals depending from a movable carriage to rotate a shaft. The mechanism may be removed from the aircraft for maintenance and the position of the mechanism within the aircraft may be adjusted for a particular user. Movement of the rudder pedals is conveyed by linkages to a collar which rotates a shaft. The rotation of the shaft is measured for control input to the aircraft or is transferred to the mechanical control systems for the aircraft.

The invention relates to a side rod control device with three degrees of freedom. The device is provided with a pitch control mechanism, a roll control mechanism and a front wheel turning control mechanism which are arranged along three mutually vertical straight lines, wherein the front wheel turning control mechanism is rotatably arranged on a first fixing part through one of a pitch control shaft of the pitch control mechanism and a roll control shaft of the roll control mechanism; and the first fixing part is rotatably arranged on a second fixing part through the other of the pitch control shaft of the pitch control mechanism and the roll control shaft of the roll control mechanism.

Techniques to control flight of an aircraft are disclosed. In various embodiments, a set of inputs associated with a requested set of forces and moments to be applied to the aircraft is received. An optimal mix of actuators and associated actuator parameters to achieve to an extent practical the requested forces and moments is determined, including by minimizing a weighted set of costs that includes costs associated with one or more errors each corresponding to a difference between a requested force or moment and a corresponding force or moment achieved by the computed solution.

A pedal system particularly tailored to a Fly-By-Wire (FBW) flight control system includes a double gradient linkage assembly. The double gradient linkage assembly includes a damper system and a spring system that improves yaw axis (azimuth) control of the aircraft. Control is only required when a change in the yaw axis state is demanded. Since only minimal displacement inputs to such a FBW flight control system is required, the travel of the pedals are exceeding compact, such that pilot workload is significantly reduced through the reduction in the frequency and magnitude of aircraft control inputs.

The invention discloses an aircraft stall protection system and a method thereof. The aircraft stall protection system include calculating a first angle of attack and a second angle of attack based on aircraft configuration and environmental conditions, the first angle of attack being greater than the second angle of attack. The system and method limit the actual aircraft angle of attack to the first angle of attack for a predetermined period of time and thereafter the system and method limit the actual aircraft angle of attack to the second angle of attack. The system and method allow the aircraft operator or pilot to extract maximum performance from the aircraft for any given set of flight conditions, without the risk of stalling or remaining in a high drag state for a prolonged period of time. This system and method are suitable for use in conjunction with a stall warning system.

A rudder pedal assembly exhibits sufficient ergonomic feel for a pilot, and is relatively small in overall size. The rudder pedal assembly includes a plurality of rails disposed non-parallel to each other within a common plane, and a rudder pedal. The rudder pedal is movably coupled to the plurality of rails and is configured to receive an input force in at least a first direction and a second direction. The rudder pedal is further configured, in response to the input force, to move along a path constrained by the plurality of rails in at least the first direction and the second direction, respectively, and within the common plane.

An aircraft encompasses a main-body, a frame-structure to support the main-body, main and auxiliary rotors provided to the frame-structure and a controller for controlling rotations of the main and auxiliary rotors. In a first mode, the controller delivers a same control signal for rotating the set of the main and auxiliary rotors, and when one of the main and auxiliary rotors becomes abnormal, the controller delivers the same control signal for compensating a decrease of the lift. In a second mode, the controller delivers a control signal only to the normal rotor for increasing the rotation of the normal rotor. Sets of the main and auxiliary rotors in divided regions adjacent to a subject divided region are rotated in a direction counter to the subject divided region. By the first and second modes, the lifts are equalized for balancing the aircraft.

A method is provided for improving aircraft engine operating efficiency during flight in aircraft driven on the ground by electric taxi systems that extends engine warm up and cool down time during ground operations without increasing the time an aircraft spends on the ground. Aircraft are driven by electric taxi systems between landing and take off with the main engines simultaneously maintained at throttle setting for time periods that ensure even and symmetrical heating or cooling for all engine components. When the aircraft reaches a location where idle or take off thrust is required before take off, engine thrust may be increased without the adverse effects of differential thermal expansion of engine components during flight. Aircraft engines may be designed to rely on electric taxi operation and extended warm up and cool down times without delaying push back or taxi-in or otherwise negatively impacting airport operations.

The invention discloses a parachute system and safe protecting method for an unmanned aerial vehicle and relates to the technical field of intelligent storage. The parachute system of the unmanned aerial vehicle comprises a sensor, a controller and a parachute; and the controller is connected with the sensor and the parachute in an electric mode and is used for detecting the flying state of the unmanned aerial vehicle, the controller is used for obtaining the flying state of the unmanned aerial vehicle form the sensor, and the parachute is controlled to be opened, when the unmanned aerial vehicle is in an unstable state. Thus, safety of the unmanned aerial vehicle is improved.

The present invention provides a protection control method of an air vehicle, a device and an air vehicle, wherein the method comprises the steps of detecting the flight state of the air vehicle (S101); when the flight state of the air vehicle is detected to be a recycle state, generating a recycle trigger signal, and sending the signal to an external protection device for the air vehicle, thereby triggering the external protection device to start the safety protection to the air vehicle (S102), wherein the flight state detection of the air vehicle comprises analyzing the flight data of the air vehicle, detecting a ground control signal received by the air vehicle, if the analysis result to the flight data is abnormal, or the detected ground control signal is a recycle trigger signal, determining the flight state of the air vehicle as the recycle state. The method can avoid an error starting of a safety device in the air vehicle more effectively, better guarantees the normal flight of the air vehicle, and reduces the damage to the air vehicle and the outside world when the air vehicle falls down rapidly due to the system failure or other reasons.

A pedal system particularly tailored to a Fly-By-Wire (FBW) flight control system includes a double gradient linkage assembly. The double gradient linkage assembly includes a damper system and a spring system that improves yaw axis (azimuth) control of the aircraft. Control is only required when a change in the yaw axis state is demanded. Since only minimal displacement inputs to such a FBW flight control system is required, the travel of the pedals are exceeding compact, such that pilot workload is significantly reduced through the reduction in the frequency and magnitude of aircraft control inputs.

The invention belongs to the technical field of an unmanned aerial vehicle, and discloses a novel composite rolling attitude control system and method. The control method integrates a principle for adjusting a flight attitude by utilizing a rudder surface and a conventional helicopter tipping tray periodically-pitch-variable principle; by utilizing a redundant control concept, an aerocraft rolling passage is controlled by a rudder system, the auxiliary compensation for the rolling passage is carried out by virtue of a rotary wing system, an airflow direction is limited by virtue of a duct, and the rudder system is controlled; and the control system comprises a controller, a rudder system, a rotary wing system and a duct. The novel composite rolling attitude control system is high in air passing capacity, compact in structure, reasonable in design and capable of realizing automatic hovering and precise attitude control and has advantages of small space size, large effective load, high control response sensitivity, high system stability, high robustness and the like.

Systems, methods, and devices are provided herein for controlling one or more movable objects (302) via a graphical user interface. A method for controlling a movable object (302) may be provided. Themethod may comprise obtaining one or more parameters of a target object (308), and generating a motion path (310) for the movable object (302) based on the one or more parameters of the target object(308). The motion path (302) may comprise a plurality of spatial points that are defined relative to the one or more parameters of the target object (308). The plurality of spatial points may be configured to be on one or more planes.

An assisted control system for controlling the attitude of a rotorcraft comprises a flight control device; at least one control member configured to act on an aerodynamic element of the rotorcraft; a mechanical connection connecting the flight control device to the at least one control member, the mechanical connection including at least one connecting rod moveable in translation, a crank device pivotably disposed about a stationary support shaft passing through the crank device, and a motor including a drive rotor and a stator configured to facilitate a pivoting movement of the crank device about the stationary support; a motor control device configured to control a rotation of the drive rotor relative to the stator; and at least one sensor configured to measure information representative of an operation performed by the flight control device under pilot control and to deliver a signal to the motor control device.

The present invention relates to a drone capable of wirelessly communicating with a mobile terminal, the mobile terminal, and a control method for the drone and the mobile terminal. The drone according to one embodiment of the present invention comprises: a communication unit for wirelessly communicating with a first mobile terminal; and a control unit for performing a function corresponding to a control signal received from the first mobile terminal, on the basis of the control signal, wherein the control unit transmits a control request signal to a second mobile terminal other than the first mobile terminal on the basis of the strength of a connected signal and / or the reception of a preset control signal in a state of being connected to the first mobile terminal so as to wirelessly communicate therewith.

A rudder pedal mechanism is configured with foreign object strike tolerance and / or with an articulating brake. The foreign object strike tolerance is implemented via a force transfer mechanism that is responsive to an input force supplied to the rudder pedal. The force transfer mechanism is configured to supply a transfer force to a rudder position command unit when the input force is supplied in a first direction, and to not supply the transfer force to the rudder position command unit when the input force is supplied in the second direction and exceeds at least a predetermined force magnitude. The articulating brake is implemented via a brake rod assembly that is rotationally coupled to the rudder pedal, and includes a first rod that is rotationally coupled to the rudder pedal, and a second rod that surrounds at least a portion of the first rod.

A system and a method for controlling a rotorcraft (1) in yaw (Y). The control system comprises a plurality of drive channels that operate an anti-torque rotor (4) of the rotorcraft (1). A first drive channel makes use of a rudder bar (13) including a set of pedals operated by the human pilot of the rotorcraft in order to control the anti-torque rotor (4). A second drive channel includes an autopilot. A third drive channel includes objective-type flight control means comprising a movable control member (12) that is operated by the human pilot and that continuously issues control signals (11) relating to a state of progression that the rotorcraft (1) is to achieve. Use of the second drive channel and of the third drive channel depends on inhibit means (19) for inhibiting operation thereof, which means are activatable by detector means (37) for detecting operation of the rudder bar (13).

The invention provides a tilting type vertical take-off and landing fixed-wing unmanned aerial vehicle and a flight control system. The tilting type vertical take-off and landing fixed-wing unmanned aerial vehicle comprises an unmanned aerial vehicle body, wings and a tail fin; the unmanned aerial vehicle body and the wings are in the form of a fixed-wing aircraft layout; the tilting type verticaltake-off and landing fixed-wing unmanned aerial vehicle comprises two supporting arms and four rotor structures; the two supporting arms are mounted on the unmanned aerial vehicle body or the wings in a direction parallel to the unmanned aerial vehicle extending direction so that a front end and a rear end distributed on the front and rear sides of the wing are formed on each supporting arm; therotor structures are correspondingly mounted at the front end and the rear end of each supporting arm; the geometric centers of the four rotor structures are coincided with the gravity center of the whole unmanned aerial vehicle; and the four rotor structures can all adjust their tilting angles so as to form a tilting power device of the unmanned aerial vehicle. The invention further provides theflight control system based on the tilting type vertical take-off and landing fixed-wing unmanned aerial vehicle.

A flapping-wing aircraft includes a support frame, a motor coupled to the support frame, a pair of wings coupled to the support frame, and a linkage assembly coupled to the support frame and configured to translate an output torque of the motor into flapping motion of the wings, wherein the linkage assembly includes a first link coupled to a rotational output of the motor, a second link pivotably coupled to the first link at a first pivot joint, a third link pivotably coupled to the second link at a second pivot joint, and a fourth link pivotably coupled to the support frame and slidably coupled to the third link, and wherein the fourth link is coupled to a first wing of the pair of wings.

The invention relates to unmanned helicopter bi-motor controlling method. It utilizes each motor torque detector to measure output torque to form loop locked inner looping of which given torsion moments are connected in parallel. The output torque is used to drive rotary wing of which rotating speed can be formed the outer looping. The outer looping controller output can be used as parallel inner looping given torsion moment to realize speed control for unmanned helicopter. Thus the invention can effectively balance the bi-motor output torque and increase rotor speed smoothness at the same time.

The invention relates to the field of control of controllers and provides a method for controlling pitch of an aircraft. The method comprises the steps that S1, a culvert fan facing a fuselage is arranged in the belly of the fuselage of the aircraft; and S2, mixed control is added to the rotating speed of the culvert fan to be used for controlling the pitch of the aircraft. The invention further provides a device for controlling the pitch of the aircraft correspondingly. The stability of the aircraft can be improved.

The invention discloses a dynamic inverse-based large-angle-of-attack control method for a fighter plane. For a large-angle-of-attack flying state, a time-scale separation method is adopted, the aircraft state variables are decomposed into two groups of subsystems based on different time scales, and a control law is solved by utilizing a dynamic inverse method respectively. According to the method, a control rudder surface deflection angle is deduced by aiming at the attitude angle speed loop, and the attitude angle speed is deduced by aiming at an airflow angle loop. According to the attitudeangle rate loop bandwidth selection, the influence of coordination control on factors such as side sliding is considered, and a frequency band bandwidth phase and the like are selected. Aiming at anairflow angle loop command model, a winding speed shaft roll angle instruction model is selected as a first-order model for meeting the flight quality requirement of the fighter plane, and an angle ofattack and a side slip angle instruction model are a second-order model. According to the method, the good tracking performance and stability of a flight control system under the large attack angle of the fighter plane are guaranteed, the dangerous states such as deep stall and tail rotation can be timely changed, and the method has good reference significance on practical application of engineering.