180 results about "Aircraft flight control system" patented technology

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 receive one or more signals to control movement of the aircraft, determine at least one flight condition of the aircraft, calculate at least one or more loads associated with the aircraft based on the determined at least one flight condition; determine a flight configuration to alleviate loads on one or more components of the aircraft based on the received one or more signals and the calculated one or more loads by predicting a distribution of loads associated with the aircraft, generate one or more effector commands based on the flight configuration; and actuate one or more aircraft effectors based on the one or more effector commands.

The invention discloses a digital twin platform of an aircraft flight control system, a data processing method of the digital twin platform and a medium, and relates to the technical field of aircraft simulation training, and the platform comprises a parameter setting unit which is used for setting aircraft basic information and current flight state parameters; the control signal acquisition unit is used for acquiring an electric signal generated when a user operates the control device; the calculation unit is used for calculating the initial control surface aerodynamic force according to the control surface aerodynamic force calculation model, calculating the control surface deflection angle according to the control surface aerodynamic-control device force transmission model, and recalculating the control surface aerodynamic force according to the control surface aerodynamic force calculation model; and the force sensing feedback unit is used for calculating joystick feedback force in the current aircraft flight state and the current control state according to the control surface aerodynamic force-control device force transmission model, and feeding back generated corresponding control resistance to the control device. The platform not only saves the economic cost and occupied area for establishing an entity iron bird test bench, but also can test the performance of a flight control system of an airplane in different flight states.

The invention discloses an unmanned aerial vehicle natural language multi-modal navigation method and system based on a multi-dimensional thinking chain. The method mainly comprises the following steps: constructing a software-in-the-loop simulation environment deeply integrated with task planning software, and collecting a live-action image with a time-space stamp; secondly, the natural language task instruction is analyzed, and key landmarks and position coordinates of the key landmarks are recognized in combination with a real scene image; then the recognition result and the structured task sequence are input into a multi-dimensional thinking chain reasoning engine in parallel to be processed; and finally, integrating the output of the multi-dimensional thinking chain inference engine, generating a flight action sequence, and issuing the flight action sequence to a flight control system to complete a navigation task. According to the method, the explicit thinking chain reasoning process is embedded into each step of navigation decision and is deeply integrated with the existing task planning ecology, so that the autonomy, adaptability and task success rate of the unmanned aerial vehicle in an unknown or dynamic environment can be remarkably improved.

An aircraft having a plurality of independent yaw control mechanisms includes an airframe having a central fuselage sized to hold at least one operator or payload and a wing extending from the fuselage. The aircraft includes a distributed thrust array coupled to the airframe, the thrust array having at least a first, second, and third pair of propulsion assemblies. Each propulsion assembly includes a rotor and is operable for at least single-axis thrust vectoring. The aircraft further includes a flight control system operable to independently control and combine each yaw mechanism of the propulsion assemblies. The yaw mechanisms for inducing a yaw moment include: canting at least one pair of propulsion assemblies away from the fuselage, selectively tilting at least one pair of propulsion assemblies forwards and backwards, varying an aerodynamic control surface of at least one propulsion assembly, and varying a rotational speed of at least one propulsion assembly.

A system and method for flight control of an aircraft, the flight control system including a plurality of flight components coupled to an aircraft, wherein the plurality of flight components includes a plurality of redundant control surfaces, a plurality of redundant low voltage buses communicatively connected to the plurality of flight components, wherein a failure in a redundant low voltage bus of the plurality of redundant low voltage busses does not impact the operability of the aircraft.

An unmanned quadcopter aircraft system integrated with a pneumatic paintball projectile launching mechanism for recreational and training applications. The system includes a compressed gas powered single-shot paintball launcher with gravity-fed ammunition hopper, mechanical counterweight recoil compensation, and multiple electronic safety systems. GPS-based geofencing prevents firing in restricted zones, while altitude sensors prevent firing above predetermined heights. The flight control system automatically reduces aircraft speed and altitude limits when the projectile system is armed. A ground control system integrates standard flight controls with firing controls, safety indicators, and emergency stop capability. An alternative embodiment allows a used to adjust the elevation of the firing system while maintaining the orientation of the drone in space and also allowing recoil compensation. The system enables aerial paintball gaming and training scenarios while maintaining safety through redundant control systems and automatic parameter limiting.

The invention discloses a self-coupling PID (proportion integration differentiation) flight control method of a rudder-controlled transverse-row type rotor unmanned aerial vehicle based on dynamic multi-dimensional heterogeneous fuzzy logic, and belongs to the technical field of flight control of rudder-controlled transverse-row type rotor unmanned aerial vehicles. A flight control system architecture of the self-coupling PID flight control method is composed of three core modules: (1) a dynamic multi-dimensional fuzzy logic coupling PID controller (DMHFL-SCPID); (2) control instruction distribution; and (3) a six-degree-of-freedom dynamical model of the rudder-controlled transverse rotorcraft. According to the control algorithm, input and output dynamic parameters of a system are analyzed in real time, and four-dimensional flight control quantities (roll, pitch, yaw and height instructions) are generated. A control distribution strategy based on the aerodynamic characteristics of the rotor converts the instructions into actuating quantity of an actuating mechanism, and flight attitude adjustment is realized by changing aerodynamic load distribution of a rotor system, so that the unmanned aerial vehicle is guided to realize accurate tracking along a predetermined flight path.

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 application discloses a kind of based on hybrid power eVTOL unmanned aerial vehicle control method and related device, its method includes: flight control computer is based on flight control instruction obtains the flight phase that eVTOL unmanned aerial vehicle currently is in, the flight phase includes vertical take-off and landing phase, overdrive mode phase, cruise phase and emergency mode phase;The flight control computer indexes to the flight control strategy corresponding to the flight phase that eVTOL unmanned aerial vehicle currently is in;The flight control computer controls the turbine power generation system and / or the power battery pack based on the flight control strategy to N cruise electric drive unit and M vertical take-off electric drive unit on the eVTOL unmanned aerial vehicle Power supply processing.In the embodiment of the application, turbine power generation system and power battery pack are connected in parallel to supply power for eVTOL unmanned aerial vehicle, while combining redundancy flight control system, so as to achieve optimization in voyage expansion and safety performance improvement.

System and method for intercepting intruding stratospheric balloons. The method comprises: receive (102, 104) a high-altitude wind profile (6) and a position (P0) of an intruding stratospheric balloon (2); Determine (106, 108), using the high-altitude wind profile (6), a future trajectory (10) of the intruding stratospheric balloon (2) and ascending trajectories (11a, 11b) of an intercepting stratospheric balloon (5) that intercept the future trajectory (10) at different positions (PF1, PF2); select (110) an upward path (11a); to perform a launch (112) of the intercepting stratospheric balloon (5) at a launch location (Li) and time (tL1) associated with the selected upward trajectory (11a); detect (114) deviations of the trajectory followed by the interceptor stratospheric balloon (5) during ascent with respect to the selected ascending trajectory (11a) and correct such deviations by activating a flight control system (12) of the interceptor stratospheric balloon (5); and activate (116) an interception apparatus (13) of the intercepting stratospheric balloon (5) once the interception (15) of the intruding stratospheric balloon (2) occurs.

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.

1. Name of the product in this design: Helicopter Flight Control Module (NEXUS-X). 2. Purpose of this design: This design is used in the flight control system of a model helicopter. 3. The key design feature of this product is its shape. 4. The image or photograph that best illustrates the design's key points: a 3D model.

The invention provides a method for reconstructing a mode conversion control law of an actuator of a large aircraft, and belongs to the technical field of flight control systems, and the method comprises the steps: S1, analyzing the influence caused by a mode conversion fault of each actuator of the large aircraft, and obtaining an analysis result; and 2, determining an actuator mode conversion control law reconstruction principle according to an analysis result, and reconstructing the control law based on the reconstruction principle. According to the method for reconstructing the mode conversion control law of the large aircraft actuator, the control law logic and the parameters after the mode conversion fault of the actuator are reconstructed, so that the control surface deflection logic and the flight quality are ensured, the calculation safety of the control law is ensured, and the fault tolerance of the control law is improved.

A rotorcraft having a main rotor with at least one main rotor actuator; a series actuator arranged upstream of the main rotor actuator; engines for powering the main rotor; and a flight control system for controlling flight operation in an AEO mode or in an AFR mode, comprising: a control logic for pre-conditioning the series actuator upon entering the AFR mode for enabling rotorcraft control upon occurrence of an IFSD, wherein pre-conditioning of the series actuator comprises shifting a reduced nominal stroke range authorized for the series actuator stroke of the series actuator in the AEO mode into a reduced shifted stroke range authorized for the series actuator stroke of the series actuator in the AFR mode to enable authorization of an increased stroke range upon occurrence of an IFSD to provide sufficient control authority for a recovery maneuver.

The invention relates to the technical field of flight control, in particular to an unmanned aerial vehicle real-time obstacle avoidance method based on space-time probability grids, and the method comprises the steps: collecting a three-dimensional visual data flow, and mapping the three-dimensional visual data flow to a probability grid unit of a space-time coordinate system; defining a current moment and a future prediction period; modeling occupancy states of the grid units based on a sparse Gaussian process, and approximating posteriori distribution by using a random variational inference method; acquiring occupancy probability distribution, and constructing and online updating a space-time probability grid; generating a space-time visual cost grid representing a future collision risk; and outputting the cost grid to a flight control system to execute trajectory planning. By adopting the sparse Gaussian process and the random variational inference method, modeling and online updating are carried out on the grid occupancy state covering the current and future periods, the calculation complexity is reduced, the space-time visual cost grid can be generated to evaluate the future collision risk, and real-time obstacle avoidance is supported.

The invention discloses an online obstacle avoidance route planning method based on random scattering, and belongs to the field of unmanned flight route planning, and the method specifically comprises the steps: firstly, obtaining initial position coordinates of a starting point and a target point in a to-be-measured obstacle region, and converting the initial position coordinates into a Mercator plane coordinate system; then, generating a minimum rectangle or square for the obstacle area; and constructing a key point set Sets by adopting a Gaussian bias random point scattering strategy. Thirdly, calculating a comprehensive evaluation function of each key point, and selecting the key point corresponding to the minimum value as an optimal waypoint at the next moment; and judging the feasibility of the minimum value of the comprehensive evaluation, outputting the waypoint to a flight control system when the minimum value of the comprehensive evaluation is feasible, and starting real-time re-planning of the next period as a new starting point. Otherwise, falling into local deadlock or no solution, returning to increase the standard deviation in the Gaussian bias random point scattering strategy, and forcing the search range to diffuse to the region deviating from the target until a feasible solution is found. The task can be completed safely and efficiently.

The utility model provides a gliding guidance kit based on an unmanned aerial vehicle. The gliding guidance kit comprises a shell, a reducing sleeve, a gliding wing, an inertial navigation assembly, a flight control system, a battery and a rudder tail, the reducing sleeve is installed at the front end of the shell, the rudder tail is installed at the tail end of the shell, the hang glider, the inertial navigation assembly, the flight control system and the battery are all installed in the shell, and the inertial navigation assembly and the flight control system are both electrically connected with the battery; under the air-drop scene, the lift-drag ratio is increased by unfolding the large hang glider of the missile, the air-drop strike distance is increased, meanwhile, the deviation rectifying capacity is achieved, and the precise strike effect is achieved.

This application relates to the field of flight control technology, specifically to a high-maneuverability flight control method for fixed-wing aircraft based on curriculum-based reinforcement learning. It constructs a closed-loop learning system comprising a curriculum scheduler, an agent, a simulation environment, and an experience buffer. The agent has a policy network. The state space of the aircraft and the normalized control surface and throttle action space output by the policy network are defined. A curriculum difficulty measurement model is established, quantifying task difficulty through weighted state deviation and envelope penalty terms. Based on the current curriculum level and this model, a safe and difficulty-matched training task set is dynamically generated. In the simulation environment, the policy network is iteratively updated through a two-layer loop training process. After completing all courses, a high-maneuverability flight control strategy is obtained. This strategy is deployed to the flight control system to achieve high-maneuverability flight control based on real-time state. This achieves efficient, safe, and adaptive flight control strategy training and deployment.

The invention discloses a fixed-wing unmanned aerial vehicle flight control method based on active disturbance rejection control, and belongs to the technical field of fixed-wing unmanned aerial vehicle flight control. The problem that the attitude control performance is poor due to the fact that aerodynamic interference of the propeller is not considered in an existing method is solved. According to the method, aerodynamic interference factors of propellers are systematically introduced into modeling and control design, namely, reaction torque and gyroscopic moment are introduced into the modeling and control design, and a flight control system of the fixed-wing unmanned aerial vehicle is designed based on an active disturbance rejection control strategy; and dynamic and external disturbances which are not covered by the model are estimated and compensated in real time, so that the attitude control performance of the fixed-wing unmanned aerial vehicle is improved. Meanwhile, under various given input signals, simulation verification is performed on the performance of the active-disturbance-rejection controller designed by the invention, which shows that the performance of the active-disturbance-rejection controller designed by the invention can meet design indexes. The method can be applied to flight control of the fixed-wing unmanned aerial vehicle.

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 invention discloses a complex environment cluster unmanned aerial vehicle intelligent obstacle avoidance flight control system, and belongs to the technical field of unmanned aerial vehicle control. The system comprises a sensing module, a data processing module, an intelligent obstacle avoidance decision module, a flight control module, a communication module and a cluster collaborative management module. The sensing module collects multi-source environment information of static / dynamic obstacles, landforms and environment interference in a complex environment and position, speed, attitude, electric quantity and load state information of unmanned aerial vehicles in a cluster. The data processing module performs preprocessing, data fusion and feature extraction on the multi-source data; according to the method, a sensor cooperation mechanism of'layered triggering + redundant backup 'is adopted, accurate sensing of different distance levels is achieved, and waste of detection resources and a blind area are avoided; according to the dynamic weight fusion algorithm, the weight is adjusted according to the real-time signal-to-noise ratio of the sensor, and the dynamic change of the performance of the sensor in a complex environment is effectively adapted in combination with a Transform-based multi-modal fusion network.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircraft driven by an electric propulsion system, as well as other types of aircraft. In some embodiments, a flight control system of an aircraft is disclosed, the flight control system configured to: determine one or more desired commands for an electric aircraft; determining at least one reference command based on the desired command and one or more aircraft conditions; monitoring an energy state of a 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; adjusting the at least one reference command based on the monitored energy states of the plurality of battery packs; generating a control command for the plurality of actuators based on the adjusted at least one reference command; and controlling the plurality of actuators to satisfy the one or more desired commands of the electric aircraft according to the generated control commands.

The present invention relates to a flight control system for a VTOL aircraft (10) that defines an aircraft reference coordinate system having a roll axis (Xb), a pitch axis (Yb), and a yaw axis (Zb), wherein the flight control system comprises first and second manual control devices for inputting control commands by an operator, and a flight control computer connected to the first and second manual control devices and configured to output flight control commands based on the pivot positions of first and second stick members with respect to first to fourth control axes, wherein the first manual control device comprises a first stick member, the first stick member being pivotable around first and second control axes with respect to a first neutral position, and the second manual control device The system comprises a second stick member, the second stick member being pivotable around third and fourth control axes with respect to a second neutral position, and the flight control computer is configured to derive and output flight control commands for longitudinal motion control based on the pivot position of the first stick member with respect to the first control axis, for lateral motion control based on the pivot position of the first stick member with respect to the second control axis, for vertical motion control based on the pivot position of the second stick member with respect to the third control axis, and for directional motion control based on the pivot position of the second stick member with respect to the fourth control axis, while at least partially eliminating the interconnection between the individual directions of motion. The present invention further relates to a VTOL aircraft (10) having such a flight control system.

The invention belongs to the technical field of aircraft flight control system design, and particularly relates to an aircraft sideslip angle calculation method which comprises the following steps: S1, acquiring parameters such as lateral overload and dynamic pressure; s2, constructing a relation between a sideslip angle and lateral overload; and S3, calculating the sideslip angle of the aircraft according to the calculation relation established in the step S2. According to the method, pneumatic data analysis is started, a theoretical calculation formula is deduced and converted, and a pseudo sideslip angle formula for solving the sideslip angle according to lateral overload, a large derivative of a lateral force coefficient to the sideslip angle, gravitational acceleration and vacuum velocity is obtained. A pseudo sideslip angle signal solved by the formula is introduced into a flight control law, and when an aircraft sideslip angle signal fails, the pseudo sideslip angle signal is used for replacing a sideslip angle signal measured by a sensor, so that the lateral course modal characteristic of the aircraft can be improved, a sideslip angle limiting function is realized, the flight quality is improved, and the flight safety is ensured. And meanwhile, a speed vector symbol driving instruction can be provided for an unmanned aerial vehicle control interface.

The invention discloses a flight control system structure of an electric tilting six-rotor unmanned aerial vehicle in a propeller aircraft mode. The flight control system structure is composed of a layered closed-loop architecture composed of outer ring attitude / height (or climbing rate) control and inner ring attitude angular rate control, and an airspeed single-loop control, trajectory guidance and actuator distribution / constraint management module. The flight control system structure of the electric tilting six-rotor unmanned aerial vehicle in the propeller aircraft mode has the remarkable advantages that under the influence of the control surface efficiency changing along with the air speed / attack angle, external disturbance and coupling, the rolling and pitching channel adjusting time is remarkably shortened, and overshoot and steady-state errors can be neglected; the airspeed is kept without overshoot, and the steady-state error is small; the yaw rate response is fast, and the coupling inhibition effect is good; the system structure calculation load is low, parameter calibration is friendly, and robustness is high; the method is suitable for attitude, airspeed and height / climbing rate cooperative control and route autonomous tracking of the electric tilt six-rotor unmanned aerial vehicle in a propeller aircraft mode.

The invention relates to the technical field of flight control, in particular to a light aircraft redundant flight control system, which comprises an assembly health evaluation module, a flight control module and a flight control module, and is characterized in that the assembly health evaluation module is used for calculating and generating a continuously quantified health state index; the control weight reconstruction module is used for reducing the weights of the components with performance degradation, and calculating and compensating the total control efficiency loss caused by the weights so as to generate a target control authority distribution matrix; the weight smooth application module is used for generating an actual application weight matrix finally applied in the current period; the control instruction distribution module is used for distributing an original control instruction to each control surface actuator by adopting the actual application weight matrix subjected to smooth processing to generate a final driving signal; according to the method, it is ensured that the aircraft always maintains the expected control response capability in the key control axial direction, and the fault-tolerant capability of the system is remarkably enhanced.

The invention discloses a precise plant protection unmanned aerial vehicle system based on machine vision, and relates to the technical field of unmanned aerial vehicles. The unmanned aerial vehicle has the beneficial effects that the unmanned aerial vehicle body, the precise spraying holder carried below the unmanned aerial vehicle body, the flight control system and the visual recognition system are arranged, precise recognition and fixed-point spraying are achieved during use, machine vision and the two-degree-of-freedom spraying holder are combined, pests and diseases can be precisely attacked, and the spraying efficiency is improved. The traditional large-area and undifferentiated spraying mode is thoroughly changed; as only the identified pest and disease damage area is sprayed, the environmental pollution caused by pesticide abuse is effectively reduced, and the farmland ecology is protected; a distributed architecture is adopted, flight control and holder control are separated, mutual interference between system tasks is avoided, and the flight stability of the unmanned aerial vehicle during execution of complex recognition and spraying tasks is guaranteed.

The invention belongs to a flight control system test technology, and provides a semi-physical simulation and test method and system for a redundancy flight control system, and the method comprises the steps: constructing a semi-physical simulation test environment, constructing a redundancy sensor simulation model integrated with a high-fidelity error model, an execution mechanism model and an aircraft dynamics model in a simulation computer; the simulation computer generates an instruction and drives the model to operate, performs instruction-response-feedback closed-loop data interaction with flight control hardware, and sequentially executes model correction, integration test, scene test, robustness test and stepped fault injection test; quantitative indicators are calculated based on the test data to assess control performance and redundancy management capability of the flight control system. According to the method, the problems of unreal modeling, incomplete test coverage and weak fault verification capability of a traditional simulation method for a redundancy system are solved, and efficient and high-fidelity ground verification of the redundancy flight control system from functions to reliability is realized.

This invention discloses a closed-loop flight control system and method for unmanned aerial vehicles (UAVs) based on multimodal dynamic gesture recognition, belonging to the field of UAV flight control. The method includes: acquiring multimodal gesture video streams; extracting spatial and frequency features using a model combining residual networks and FreqFormer, then fusing the extracted gesture prediction categories; quantizing the model using INT8 and deploying it on an airborne NPU based on a producer-consumer model; using a recognition queue to perform mode smoothing on the gesture prediction categories to obtain valid gesture commands, which are then mapped to flight control targets; combining positioning information with cascaded PID control to calculate rotational speed signals, and introducing a first-order inertial link model for delay compensation; improving recognition accuracy in complex environments through multimodal fusion and frequency domain attention mechanisms; ensuring airborne real-time performance through quantization and parallel processing; and significantly enhancing the smoothness and closed-loop stability of UAV gesture control by combining smoothing and delay compensation techniques.