508 results about "Fly control" patented technology

The present disclosure relates generally to flight control of electric aircraft and other powered aerial vehicles. In one embodiment, a method is disclosed, comprising: receiving a descent rate command from a pilot input device, determining a proximity of each propeller of at least two propellers to a vortex ring state; and controlling the aircraft's descent rate to be less than the commanded descent rate when at least one of the at least two propellers is within a first threshold proximity to the vortex ring state.

The invention discloses a flight control method and system for an electric power inspection unmanned aerial vehicle, and relates to the technical field of unmanned aerial vehicle inspection flight control. According to the flight control method of the electric power inspection unmanned aerial vehicle, the current flight position of the unmanned aerial vehicle is obtained in real time, deviation analysis is performed on the current flight position and a target inspection position, and when the deviation exceeds a preset threshold value, disturbance sensing data in a set area is further obtained and a disturbance sensing intensity index is calculated; according to a comparison result of the index and a disturbance attribution judgment threshold value, an offset cause is judged, and a corresponding disturbance response control strategy or a flight error correction strategy is executed respectively, so that the flight stability and the control precision of the electric power inspection unmanned aerial vehicle are dynamically guaranteed; according to the invention, through dual discrimination of the disturbance perception intensity index and the flight error adjustment index, external interference and flight control error causes are accurately distinguished, path offset cause identification and strategy matching are realized, and the flight stability and control precision of the electric power inspection unmanned aerial vehicle in a complex environment are improved.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a computer-implemented method for controlling an aircraft is disclosed, comprising: receiving, from a source processor, a first copy of a signal corresponding to an input device; sending a second copy of the signal to all other processors; receiving a number of second copies of the signal from all other processors, the number of second copies being equal to the number of all other processors excluding the source processor; determining a consensus signal based on the first copy and the second copies of the signal; and determining a command signal for an effector of the aircraft based on the consensus signal, and wherein no two processors are configured to receive signals from a same input device.

A fixed-wing unmanned aerial vehicle large-attack-angle longitudinal aerodynamic characteristic test device and a test method relate to the technical field of aircraft aerodynamic characteristic test, and mainly comprise a wind tunnel simulation device, an unmanned aerial vehicle bearing and motion simulation device, a sensor arranged on an unmanned aerial vehicle or a model, a data acquisition system and the like. The simulation module is used for simulating longitudinal motion states and attitudes of the unmanned aerial vehicle in various pneumatic environments; the device is used for experiments, nonlinear longitudinal aerodynamic characteristics of the unmanned aerial vehicle under various attack angles are obtained, a phase diagram reflecting the attack angle-pitch angle speed change relation of the system state is obtained, a balance point under the normal working condition and the limit working condition is determined, and a basis is provided for flight control of the unmanned aerial vehicle under the large attack angle limit working condition.

The invention discloses an aircraft voice instruction security analysis method and system based on natural language processing, and aims to realize intelligent identification and execution control of aircraft operation instructions. The method comprises the steps of receiving a voice instruction audio input by an aircraft operator, and performing preprocessing to generate a standardized voice segment; converting the standardized voice segment into text content, and extracting structured semantic information related to the instruction; calculating a semantic intention confidence score of the instruction in combination with the structured semantic information, the aircraft operation state, the historical instruction and the current task mode; according to the confidence score and the security policy rule, jointly evaluating the security of the instruction and outputting an executable state identifier; when the state identification is executable, calculating a flight control command and submitting the flight control command to a flight control system for execution; and when the state identifier is non-executable, the interception operation is automatically executed, and abnormal prompt information is fed back to an operator through the man-machine interaction module.

The invention provides an air-ground cooperative wetland air patrol method and device and electronic equipment, and the method comprises the steps: making a patrol flight plan based on an operation region and a flight limiting region defined by an operation control management center; daily first flight confirmation is carried out, the unmanned aerial vehicle is controlled to take off under the condition that the flight control right is obtained, and automatic patrol is carried out according to the patrol flight plan. In the automatic patrol process of the unmanned aerial vehicle, image video data are collected in real time, the operation condition is monitored, and the real-time equipment state and alarm information of the unmanned aerial vehicle are obtained. Analyzing according to the image video data to obtain an inspection result, and generating an alarm signal when an abnormal condition exists in the inspection result. And judging whether the unmanned aerial vehicle has a fault based on the real-time state of the equipment and the alarm information, and controlling the unmanned aerial vehicle according to a preset fault handling strategy when the unmanned aerial vehicle has the fault. According to the method, an air-ground cooperation technology and an intelligent technology are combined, the wetland inspection efficiency is effectively improved, and intelligent management of the wetland protection area is realized.

The invention relates to the technical field of flight control, in particular to a hybrid control FPV unmanned aerial vehicle flight control system and method. The method comprises the following steps: acquiring triaxial angular velocity data of the unmanned aerial vehicle in real time by using an inertial navigation system, eliminating high-frequency noise and performing zero offset calibration; constructing a corresponding linear extended state observer to adjust a gain parameter in real time based on an absolute value of an observation error, and filtering tiny noise disturbance through a dead zone function to realize high-precision estimation of the system state and total disturbance of the unmanned aerial vehicle so as to obtain an error and an error change rate; and designing a self-adaptive PID control law to dynamically adjust PID control quantities corresponding to proportion, integral and differential gains, carrying out dynamic weighted stacking according to a flight mode, generating a three-axis control quantity, carrying out flight control on the FPV unmanned aerial vehicle, and meanwhile, realizing real-time visual verification of a control effect corresponding to the flight control. According to the invention, high-precision and high-robustness attitude control can be realized.

The invention discloses an unmanned autonomous cluster flight control method based on a bionic alert mechanism, which is applied to the field of unmanned autonomous cluster control, and is characterized in that a W-MSR algorithm and a dynamic weighted bionic alert mechanism are fused, the W-MSR filters and eliminates extreme values of neighbor individuals, the dynamic weighted bionic alert mechanism strengthens input from informed individuals, and the dynamic weighted bionic alert mechanism is used for improving the robustness of the unmanned autonomous cluster flight control. And meanwhile, the influence of suspicious individuals is inhibited, so that the local information aggregation degree is effectively improved, and group splitting can be prevented. Different from a scheme depending on explicit attacker recognition, the method can improve the motion accuracy and connectivity of the unmanned autonomous cluster in a confrontation environment by probabilistically suppressing the influence of suspicious individuals and amplifying a consistent signal, provides a new idea for improving the security of a swarm intelligence system, and has a wide application prospect. Damage of malicious individuals to the cluster is effectively defended, and robustness and recovery capability of the cluster in a complex environment are improved.

An actuator system for a flight control surface of an aerial vehicle includes a motor, and an inner shaft. The inner shaft includes a first portion at a first end and a second portion that extends from the first portion to a second end of the inner shaft. The first portion of the inner shaft is coupled to the motor. The second portion of the inner shaft is coupled to at least one roller that includes roller threads. The actuator system includes an inverted nut defining an inner bore and having an outer diameter that defines an output rod. The inner bore includes nut threads. The inverted nut is coupled about the inner shaft such that the nut threads are to engage with the plurality of roller threads to move the output rod, and the output rod is to be coupled to the flight control surface.

The invention relates to the technical field of flight control, in particular to an unmanned aerial vehicle single-shot failure control method, system and device and a medium, and the method comprises the steps: single-shot failure control in the take-off stage of an unmanned aerial vehicle, namely, taking-off is interrupted, an engine on the other side is closed, a brake is used, and a rudder and a front wheel are used for carrying out deviation correction control on the unmanned aerial vehicle; single-engine failure control in the air cruise stage of the unmanned aerial vehicle: compensating additional yaw moment brought by a single-side engine through a transverse heading automatic compensation controller and a flight control system together to enable the aircraft to reach a balanced state, and adjusting the heading of the unmanned aerial vehicle to a designed emergency course to complete recovery; the single-shot failure control in the landing stage of the unmanned aerial vehicle comprises the following steps of: keeping the balance of the unmanned aerial vehicle by utilizing a transverse heading automatic compensation controller before landing, immediately closing an engine on the other side after touching the ground, and adopting a rudder and a front wheel to rectify deviation to control and track a ground route. The flight safety of the unmanned aerial vehicle is guaranteed, and the single-shot failure emergency processing capacity of the unmanned aerial vehicle is improved.

The invention relates to the field of unmanned aerial vehicle inspection, in particular to a multifunctional mechanical arm system for distribution network unmanned aerial vehicle inspection. The system comprises an unmanned aerial vehicle platform, a mechanical arm module, a bionic control system and a replaceable operation assembly. The unmanned aerial vehicle platform provides flight control and load bearing functions; the mechanical arm module executes grabbing, spraying and laser operation through a multi-degree-of-freedom joint structure and a driving unit; the bionic control system collects arm motion trail data of an operator in real time and converts the arm motion trail data into a mechanical arm motion instruction, and synchronous motion control is achieved. The replaceable operation assembly is connected through a standardized interface, and different operation requirements are met. Through the bionic control system, the problems of complex operation and insufficient precision in the prior art are solved, and the precision and response speed of the unmanned aerial vehicle in tasks such as laser foreign matter removal, insulating paint spraying and foreign matter grabbing are remarkably improved.

The folding type rotor unmanned aerial vehicle comprises a cylindrical shell, a flight control cabin, a folding mechanism cabin, a communication and load cabin and an image cabin are arranged in the shell from top to bottom in a longitudinal row mode, a multi-rotor unmanned aerial vehicle body with longitudinally-folded machine arms is arranged in the folding mechanism cabin, and containing openings are formed in the positions, corresponding to the machine arms, of the side wall of the shell. The outer ends of the vehicle arms are connected with propellers through motors, and the propellers can be clamped into the corresponding containing openings after being folded. According to the folding type rotor wing unmanned aerial vehicle, the blades of the propellers can be automatically clamped in the containing openings of the shell after being folded, meanwhile, the folded vehicle arms are restrained in the shell together, other special restraining structures are not needed, and therefore extra space occupation and restraining operation are reduced, and the carrying performance is greatly improved; and the blades of the propeller can be automatically unfolded after being separated from the containing opening, the vibration problem caused by the fact that the propeller is not unfolded in place when the motor rotates is avoided, and the take-off stability of the unmanned aerial vehicle is improved.

Disclosed here are unmanned aerial vehicle embodiments including some embodiments having a fuselage, tail, and wings including example embodiments with an adaptable payload section, alternatively or additionally, modular flight surfaces including tail, wings and motor, alternatively or additionally the vehicle configured for short landings with reversible thrust, alternatively or additionally, the unmanned aerial vehicle configured with direct connection to moveable flight control surfaces.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a system of an aircraft is disclosed, configured to determine one or more desired commands for the aircraft, retrieve engine information for at least one electric propulsion unit (EPU) of a plurality of EPUs of the aircraft, wherein the engine information includes at least one time-based metric for temperature associated with the at least one EPU, generates control commands based on the received engine information, and controls effectors according to the generated control commands to meet the desired commands of the aircraft.

The invention provides a personal aircraft chassis and a personal aircraft. The chassis comprises a main framework and a chassis body, the personal aircraft comprises a main framework and a thrust array assembly, the thrust array assembly is fixed to the main framework, thrust of the thrust array assembly is achieved through a plurality of first turbojet engines, the geometric position relation between the thrust action point and the thrust direction of the thrust array assembly and the main framework is kept fixed, and when the personal aircraft stops on the ground, the thrust array assembly is fixed to the main framework. The thrust direction of the thrust array assembly is perpendicular to the ground. The vector nozzle assemblies are fixed to the main framework, thrust of the vector nozzle assemblies is achieved through the second turbojet engines, the central axes of the second turbojet engines are perpendicular to the central axis of the first turbojet engine, and vector nozzles of the vector nozzle assemblies at least have one degree of freedom relative to the main framework. And high-dynamic, high-fault-tolerance and reconfigurable flight control which cannot be realized by utilizing pure thrust difference control can be realized.

An assembly for maintaining a pitch angle of a flight control surface of an aircraft is disclosed. The assembly includes a shaft including a thread and defining an axis of the shaft; and a nut. The nut includes: a housing; a barrel located in the housing and including threads for engaging with the threads of the shaft; and at least one locking element, the at least one locking element being located in the housing. The housing includes a radially inner cam surface adjacent the locking element. The barrel is configured to rotate relative to the housing when the threads of the barrel engage with and rotate with the threads of the shaft, thereby moving the locking element along the cam surface from a disengaged position, in which the locking element is spaced from the threads of the shaft, to an engaged position, in which the locking element is wedged between the threads of the shaft and the housing, as a result, relative rotation between the shaft and the housing is prevented.

The invention discloses a flight control method for a tilt transition stage of a tilt-rotor aircraft. The flight control method comprises the following steps: constructing a tilt corridor of the tilt-rotor aircraft; acquiring the vertical control quantity of the aircraft in the tilting transition stage in real time; the tilt angle corresponding to the upper boundary of the tilt corridor serves as a tilt angle feed-forward instruction of the aircraft in the tilt transition stage; a tilt angle beta is obtained; taking the flight speed of the upper boundary of the tilting corridor corresponding to the tilting angle beta as a flight speed instruction of the aircraft in the tilting transition stage; correspondingly obtaining a final accelerator instruction of the tilting rotor; and performing flight control on the aircraft in the tilting transition stage. The tilt angle can be determined based on the tilt corridor so as to correspondingly obtain the final accelerator instruction of the tilt rotor, meanwhile, the flight height is kept unchanged at the moment of the tilt transition stage, and the flight speed is controlled within the preset range, so that the height and speed control precision of the aircraft is improved, and the control precision of the aircraft is improved. And therefore, the controllability and the safety in the tilting flight process are improved.

The invention relates to an air docking reliable control method based on additive state decomposition, and belongs to the field of flight control. Comprising the steps that a docking aircraft model, a docking device model on a docking aircraft and a docking device model on an aircraft to be docked are constructed, and an air docking comprehensive model is obtained based on the three sorts; carrying out equivalent decomposition on the air docking comprehensive model based on additive state decomposition to obtain an accurate nonlinear subsystem, an accurate linear subsystem and a linear subsystem with interference; designing a first controller based on the precise nonlinear subsystem, designing a second controller based on the precise linear subsystem, and designing a third controller based on the linear subsystem with interference; and adding the three controllers to obtain an air docking reliable controller. According to the method, complex interference and nonlinear characteristics in an air docking system can be effectively processed, high-precision tracking docking control is realized, and the safety, reliability and robustness of air docking are improved.

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 discloses a four-rotor aircraft trajectory tracking control method based on an improved event triggering mechanism. In a four-rotor aircraft trajectory tracking task, an anti-saturation auxiliary system is firstly designed to solve the problem of actuator input saturation of a four-rotor aircraft. The saturation phenomenon is modeled and compensated in advance, so that the influence of the problem on the control performance of the system is reduced. And secondly, through combination of a self-adaptive gain item and a fuzzy controller, a sliding mode reaching law hyperbolic tangent function coefficient is softened, a chattering phenomenon existing in sliding mode control is further inhibited, and the control performance is improved. Finally, an improved event triggering mechanism is designed at an actuator end, a self-adaptive triggering threshold value is introduced, redundant transmission is avoided, dynamic adaptability is enhanced, and communication resources are further saved. According to the method provided by the invention, the requirement of green flight control is met while the control performance of the system is improved.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a flight control system of an aircraft is disclosed, configured to determine one or more desired commands for the electric aircraft, determine at least one reference command based on the desired commands and one or more aircraft conditions, monitor energy states of the plurality of battery packs, wherein at least a first battery pack of the plurality of battery packs is electrically isolated from at least a second battery pack of the plurality of battery packs, adjust the at least one reference command based on the monitored energy states of the plurality of battery packs, generate control commands for the plurality of effectors based on the adjusted at least one reference command, and control the plurality of effectors according to the generated control commands to meet the one or more desired commands of the electric aircraft.

The invention belongs to the technical field of flight control, particularly relates to a simulation control method and device for a control load system and a storage medium, and aims to solve the problem of limited effect during simulation control of the control load system in the prior art. The method comprises the following steps: acquiring manipulation input data; carrying out filtering processing on the joystick force / torque data, and carrying out dynamic simulation on the control load system through a pre-constructed digital twin system based on the filtered joystick force / torque data and the control position data so as to determine a virtual force / torque track of the control load system for responding to the control input data; acquiring current flight parameters of the flight simulator, and resolving according to a preset aerodynamic model to obtain a reference force / torque trajectory; and taking the reference force / torque track as a control target, performing optimization solution on the virtual force / torque track, and determining an actuator driving instruction. According to the method, the effect is better when the control of the control load system is executed.

The invention discloses an arm two-degree-of-freedom transformation quadrotor flight control method and device. The method comprises the following steps: S1, carrying out structural design on an arm two-degree-of-freedom transformation quadrotor; s2, deducing a kinetic model and a kinematic model of the two-degree-of-freedom transformation quadrotor of the arm according to the two-degree-of-freedom transformation quadrotor structure of the arm; s3, deducing a control distribution model of the two-degree-of-freedom deformation quadrotor of the arm according to the steps S1 and S2; s4, analyzing the singularity problem of a control distribution model according to the control distribution model of the two-degree-of-freedom deformation quadrotor of the arm derived in the step S3, and designing a special control distribution model for the singularity; and S5, designing a pose quaternion nonsingular terminal sliding mode flight controller according to the dynamical model, the kinematic model, the control distribution model and the special control distribution model of the arm two-degree-of-freedom deformation quadrotor. The method has good control and stabilization effects on the two-degree-of-freedom deformation four rotors of the arms.

The invention belongs to the technical field of wind tunnel test and flight control crossing, and discloses a control device and a control method of a multivariable hypersonic variable configuration aircraft model. The control device comprises a main control computer; the main control computer is used for project management and development, data storage and display, flight and control simulation; the system also comprises a real-time simulation machine which is used for simulating the real-time environment of the flight control system, and completing the data collection of the sensor, and the generation and transmission of a control instruction. The control method comprises the steps of establishing a digital simulation model; verifying the digital simulation model; compiling a control law program and testing; performing a virtual flight test; and analyzing and processing data. According to the control device and the control method of the multivariable hypersonic-speed variable-configuration aircraft model, simulation of a multivariable hypersonic-speed variable-configuration aircraft digital model and a control law is realized, and a control algorithm operated in a real-time simulation machine can be used for carrying out a wind tunnel pneumatic, motion and control coupling virtual flight test; the engineering practical value is realized.

The invention relates to the technical field of low-altitude aircrafts, in particular to an emergency landing system and method for an aerocar, and ensures stable operation of a parachute in emergency through accurate opening judgment, starting, unfolding control and wind resistance adjustment functions of a parachute module. Secondly, the low-altitude sensing and obstacle avoidance module realizes real-time monitoring of the surrounding environment of the hovercar and formulation of an obstacle avoidance strategy by integrating various sensors and intelligent algorithms. And the emergency landing safety air bag module is used for rapidly unfolding a safety air bag before landing, so that an effective buffer layer is formed, and the impact force during landing is reduced. And the precise flight control module is responsible for keeping the stable attitude and precise path planning of the hovercar in the landing process. And finally, the power management module is automatically switched to the standby power supply when the main power supply fails, so that the continuous operation of the key system is ensured.

The invention discloses a bionic butterfly based on a flapping wing structure, and belongs to the field of flapping wing type aircrafts, a butterfly main body comprises a rack, two motors assembled with speed reduction groups, two couplings, two wings and two wing connecting pieces; the wing connecting piece comprises connecting parts and a transmission shaft which are connected with each other, and the two wings are respectively fixed on the two connecting parts; the rack is T-shaped, longitudinal mounting holes matched with a transmission shaft in shape are formed in the two ends of the transverse part of the rack, and semicircular fixing grooves matched with a motor in shape are formed in the two sides of the tail end of the longitudinal part of the rack; the two transmission shafts are respectively mounted in the two mounting holes; the two motors are respectively bonded in the two fixing grooves; one ends of the two couplings are connected with the output ends of the two speed reduction sets respectively, and the other ends of the two couplings are connected with the two transmission shafts respectively. The flapping wing type bionic butterfly provided by the invention is small in size and light in weight, and more efficient flapping wing driving and more stable flight control can be realized.

The invention discloses a semi-open source control system for unmanned aerial vehicles and a design method thereof which is constructed single-chip multi-core processor and shared register on a UAV. A main processing core of the single-chip multi-core processor is built with a semi-open source control module configured to control the UAV. The semi-open source control module includes a closed source control function module and an open source control program module. The closed source control function module includes data processing function module, communication function module, flight control function module and hardware setting function module, enabling the semi-open source control module to perform power management, ground communication, aviation sensing, flight control, navigation calculation and flight record information processing of the UAV. A secondary processing core of the single-chip multi-core processor is built with user application programs enabling users to program working task application programs according to their required tasks, and the working task application programs are composed with the open source control program module through the shared register. Thereby, an autopilot program is capable to be built in the main processing core and flight status and waypoints saved in shared registers, users can develop other application programs through other processing cores for navigation or AI computing, then achieve the aims of increasing the stability of system software and hardware, reducing space and weight, and reducing current consumption and cost.

The invention provides a modular free-combination rotor unmanned aerial vehicle and a combination method, and the unmanned aerial vehicle comprises a plurality of independent rotor unmanned aerial vehicle units, and each rotor unmanned aerial vehicle unit comprises a fuselage and a rotor; a battery, a motor, a communication module and a power control module are arranged in the machine body; a standard socket is arranged outside the fuselage and used for being fixedly connected with other rotor wing unmanned aerial vehicle units or rods with the same interfaces, and sharing of power supply and communication is achieved through the standard socket. The power control module dynamically adjusts rotor power distribution in real time according to the form of the combined unmanned aerial vehicle, and the communication module automatically establishes a master-slave control relation after combination to achieve unified coordination of flight control. According to the technical scheme provided by the invention, through free combination of the modular rotor unmanned aerial vehicle units, a specific unmanned aerial vehicle can be quickly formed, the purchase and use cost is reduced, and the flight stability and safety are improved.

The invention discloses an aircraft attack angle protection control method based on dynamic instruction correction, belongs to the field of flight control methods, and is used for solving the tracking control problem under the condition that the aircraft attack angle is limited. Firstly, based on an aircraft dynamics model, a backstepping method is utilized to design an adaptive controller to ensure stable tracking of a system, a virtual attack angle instruction part needs to be limited by adopting a saturation function, and a limited boundary value is provided by a subsequent protection system. And then, designing a prediction model representing correlation between the attack angle and the attack angle instruction and dynamics based on the designed closed-loop control system, reversely deducing a complex function relationship between the attack angle instruction and the system state and dynamics through the prediction model, and constructing an attack angle protection system. And finally, calculating an attack angle instruction boundary value through the prediction model and a system motion state trend under the current input, and correcting an original virtual attack angle instruction boundary so as to realize attack angle protection. According to the method, stable tracking of the system under attack angle constraint can be ensured, the residual attack angle in the envelope can be fully utilized, the maneuverability of the aircraft is improved, and the method has an engineering application reference value.