Self-adaptive reverse cooperative control method and system for coaxial dual-rotor aircraft

Abstract

The invention belongs to the technical field of aircraft control, and discloses a self-adaptive reverse cooperative control method and system for a coaxial dual-rotor aircraft, and the method comprises the steps: obtaining a target attitude instruction; according to whether pitching and rolling instructions exceed a set threshold value or not, a reverse cooperative control strategy or a conventional same-direction control strategy is intelligently selected, and upper and lower rotor wing control instructions are obtained through calculation; according to the yawing and vertical speed instructions, the target rotating speed of the double motors is calculated; and converting the control instruction into a steering engine PWM signal and synchronously outputting the steering engine PWM signal to an execution mechanism. The system correspondingly comprises an attitude instruction acquisition module, a control instruction calculation module, a motor rotating speed calculation module, a steering engine signal conversion module and a synchronous output module. According to the method, the response capacity under large maneuverability is improved through reverse cooperative control, safe takeover and degraded operation under the steering engine set fault are achieved through a fault detection and redundancy control mechanism, and the control precision, maneuverability and reliability of the aircraft under the complex working condition are remarkably improved.

Description

Technical Field

[0001] This invention belongs to the field of aircraft control technology, specifically relating to an adaptive reverse cooperative control method and system for a coaxial dual-rotor aircraft. Background Technology

[0002] Coaxial twin-rotor aircraft, as a special type of rotorcraft configuration, have shown broad application prospects in various fields such as reconnaissance, inspection, logistics, and agricultural plant protection due to their compact structure, high lift efficiency, strong wind resistance, and unique flight control characteristics. Compared with single-rotor or multi-rotor aircraft, the coaxial twin-rotor configuration effectively cancels out counter-torque by having two rotors rotate in opposite directions along the same axis, simplifying the mechanical transmission structure and providing higher lift density.

[0003] However, coaxial twin-rotor aircraft still face a series of challenges in practical applications. First, their dynamic model is complex, and the aerodynamic coupling effect between the upper and lower rotors is significant, resulting in inherent nonlinear and strongly coupled characteristics in attitude and position control. This makes it difficult for traditional linear model-based control methods to achieve high-precision, wide-range attitude control across the entire flight envelope, especially when performing large-amplitude maneuvers or facing complex external disturbances such as gusts, where the aircraft's responsiveness, stability, and control accuracy are often unsatisfactory.

[0004] Secondly, existing control systems lack maneuverability and adaptability under complex operating conditions. Traditional coaxial twin-rotor aircraft control typically employs a co-directional cooperative control strategy, where the pitch commands of the upper and lower rotors change in the same direction to alter the total lift or attitude. This mode performs adequately in conventional flight, but when high-difficulty maneuvers such as rapid attitude adjustments or high angle-of-attack flight are required, it may fail to achieve ideal results due to aeroelasticity or control margin limitations, and may even lead to flight instability.

[0005] Furthermore, the reliability and fault tolerance of aircraft need improvement. With the increasing complexity of coaxial rotorcraft applications, the probability of failure in critical actuators (such as servo motors or motors) during missions increases. Most existing aircraft control systems lack effective adaptive fault detection and redundant control mechanisms. Once a critical component fails, the entire flight system may rapidly become unstable, leading to mission failure or even catastrophic loss of life and the aircraft, severely limiting its widespread application in missions with high safety requirements.

[0006] Therefore, existing coaxial dual-rotor aircraft control technologies still have significant shortcomings in improving the aircraft's maneuverability, flight stability, and reliability under complex operating conditions. There is an urgent need for a more advanced, efficient, and fault-tolerant cooperative control method and system to overcome the deficiencies of the existing technologies. Summary of the Invention

[0007] To overcome the shortcomings of existing technologies, this invention provides an adaptive reverse cooperative control method and system for coaxial dual-rotor aircraft, aiming to solve the technical problems of insufficient maneuverability, limited control accuracy, and lack of effective fault tolerance in existing coaxial dual-rotor aircraft under complex operating conditions. By introducing a reverse cooperative control strategy, adaptive fault detection, and redundant control mechanisms, the limitations of traditional co-directional control modes are overcome, significantly improving the attitude control accuracy and rapid maneuver response capability of the aircraft in complex disturbance environments, while enhancing flight safety and mission reliability in the event of failure of key actuators.

[0008] To achieve the above objectives, the first aspect of the present invention provides an adaptive reverse cooperative control method for a coaxial dual-rotor aircraft, applicable to a coaxial dual-rotor aircraft having independently driven first and second motors, and independently controlled upper and lower rotor control mechanisms, comprising the following steps:

[0009] S1. Obtain target attitude commands, including pitch angle command θ_cmd and roll angle command. Yaw angle command and vertical speed command ;

[0010] S2. The target attitude command is calculated into an upper rotor control command group and a lower rotor control command group, wherein the upper rotor control command group includes the upper rotor longitudinal cyclic pitch angle. Lateral cyclic pitch angle of the upper rotor and rotor collective pitch angle The lower rotor control command group includes the lower rotor longitudinal periodic pitch angle. Lateral cyclic pitch angle of the lower rotor and the collective pitch angle of the lower rotor The solution strategy is as follows:

[0011] when When, set and ;

[0012] when When, set and ;

[0013] when and At that time, the conventional co-directional control strategy is adopted to make the longitudinal periodic pitch angle of the upper and lower rotors change in the same direction and the lateral periodic pitch angle change in the same direction.

[0014] in and These are the preset pitch maneuver threshold and roll maneuver threshold, respectively. and These are the pitch control gain and roll control gain, respectively, and when and Simultaneously execute the above-mentioned reverse coordination settings for pitch and roll;

[0015] S3, according to the yaw angle command and vertical speed command Calculate the target speed of the first motor Target speed of the second motor The calculation method is as follows:

[0016] Reference speed Speed ​​difference ,but , ,in Yaw control gain;

[0017] S4. Convert the upper rotor control command group and the lower rotor control command group into PWM position control signals for the first servo motor group and the second servo motor group respectively through inverse kinematics calculation;

[0018] S5. Synchronously output the PWM position control signal of the first servo motor group, the PWM position control signal of the second servo motor group, and the target speed command of the first motor. and the target speed command for the second motor .

[0019] As a further embodiment of the control method of the present invention, the conventional co-directional control strategy is to make the longitudinal periodic pitch angles of the upper and lower rotors change in the same direction, and to make the lateral periodic pitch angles of the upper and lower rotors change in the same direction.

[0020] 3. The cooperative control method for a coaxial dual-rotor aircraft according to claim 1, characterized in that, in step S3, the target speed of the first motor is... Second motor target speed The calculation method is as follows:

[0021] Reference speed ;

[0022] Speed ​​difference ;but ;

[0023] in, This is the yaw control gain.

[0024] As a further embodiment of the control method of the present invention, a fault detection step S6 is also included:

[0025] Real-time monitoring of the operating status of each servo motor in the first and second servo motor groups; when any of the following fault judgment criteria are met, the corresponding servo motor group is determined to be faulty:

[0026] (a) The position deviation continuously exceeds the preset deviation threshold and the duration exceeds the preset time threshold;

[0027] (b) Abnormal drive current;

[0028] (c) The temperature exceeds the safety threshold; the fault determination needs to be confirmed through multiple consecutive control cycles.

[0029] As a further embodiment of the control method of the present invention, after fault detection, the following redundant control strategy S7 is further executed:

[0030] S7.1 When a fault is detected in the first servo motor group, the upper rotor control command group is set to a safe value, and the attitude control task is completely assigned to the second servo motor group, and the enhanced lower rotor control command is calculated.

[0031] S7.2 When a fault is detected in the second servo motor group, the downrotor control command group is set to a safe value, the attitude control task is completely assigned to the first servo motor group, and the enhanced uprotor control command is calculated.

[0032] As a further embodiment of the control method of the present invention, the enhanced uprotor control command or downrotor control command is obtained by compensating for gain. The calculation shows that the Based on attitude control error Adaptive adjustment.

[0033] As a further embodiment of the control method of the present invention, the redundancy control strategy S7 also includes a dynamic motor power allocation strategy under fault modes:

[0034] When one side of the servo motor fails, the output power of the motor on the failed side is reduced, while the output power of the motor on the healthy side is increased.

[0035] As a further embodiment of the control method of the present invention, the redundant control strategy S7 also includes a safe return or landing mode S8:

[0036] In redundant control mode, the flight speed and yaw rate are limited, and a path is automatically planned to return to a safe landing point.

[0037] A second aspect of the present invention provides a coaxial dual-rotor aircraft control system for implementing the above-mentioned control method, comprising:

[0038] The attitude command acquisition module is used to acquire the target attitude command;

[0039] The control command calculation module is used to calculate the target attitude command into an upper rotor control command group and a lower rotor control command group, and can switch between the reverse cooperative control strategy and the conventional co-directional control strategy.

[0040] The motor speed calculation module is used to calculate the target speeds of the first motor and the second motor.

[0041] The servo control signal conversion module is used to convert the control command group into the corresponding servo servo group PWM position control signal;

[0042] The synchronous output module is used to simultaneously output the PWM position control signal and the motor target speed command.

[0043] As a further embodiment of the control system of the present invention, it also includes: a fault detection module, used to monitor the working status of the servo motor group in real time and determine the fault according to the preset fault judgment criteria; and a redundancy control module, used to execute a redundancy control strategy when a fault of the servo motor group is detected.

[0044] The beneficial effects of this invention are:

[0045] This invention introduces a reverse cooperative control strategy, enabling a coaxial dual-rotor aircraft to generate attitude control torques far exceeding those of traditional co-directional control under high-maneuverability conditions. This results in faster and more precise attitude adjustments, significantly improving maneuverability and meeting the demands of complex missions. Meanwhile, conventional co-directional control strategies, under low-maneuverability conditions, ensure aircraft stability and endurance by reducing aerodynamic interference and optimizing energy consumption, thereby improving flight stability and efficiency.

[0046] Through a fault detection mechanism, this invention can detect potential problems with the servo motor assembly in real time. In the event of a fault, the redundant control strategy can quickly switch, with the healthy rotor taking over control and adaptively adjusting control gain and motor power to effectively maintain the aircraft's controllability. This enhances fault tolerance and prevents crashes caused by single points of failure. In redundant control mode, a safe return-to-home or landing mode is activated, limiting flight parameters and automatically planning a safe path, greatly ensuring flight safety and improving the aircraft's survivability and mission success rate in unexpected situations.

[0047] Furthermore, by independently calculating the reference speed and the speed difference, this invention effectively decouples vertical speed, total lift, and yaw control, simplifying control design and improving control accuracy. Attached Figure Description

[0048] Figure 1 This is a front view schematic diagram of the coaxial dual-rotor aircraft to which this invention applies.

[0049] Figure 2 This is a three-dimensional structural diagram of the coaxial dual-rotor aircraft to which this invention applies.

[0050] Figure 3 This is a schematic diagram of the overall architecture of the coaxial dual-rotor aircraft control system of the present invention.

[0051] Figure 4 This is a flowchart of the collaborative control method of the present invention.

[0052] Figure 5 This is a schematic diagram of the solution strategy in step S2 of the present invention.

[0053] Figure 6 This is a schematic diagram of the motor speed calculation in step S3 of the present invention.

[0054] Figure 7 This is a schematic diagram of the fault detection principle in step S6 of the present invention.

[0055] Figure 8 This is a schematic diagram of the redundancy control strategy in step S7 of the present invention.

[0056] In the diagram: Upper rotor system 1, Lower rotor system 2, First brushless motor 3, First servo motor group 4, Upper swashplate 5, Second brushless motor 6, Lower swashplate 7, Second servo motor group 8, Flight controller 9. Detailed Implementation

[0057] The specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0058] The coaxial dual-rotor aircraft to which this invention applies, such as Figure 1 , 2 As shown, the system includes an upper rotor system 1 and a lower rotor system 2 arranged coaxially and rotating in opposite directions. The upper rotor system 1 is independently driven by a first brushless motor 3, and its pitch is controlled by an upper swashplate 5 connected to it via a first servo motor group 4. The lower rotor system 2 is independently driven by a second brushless motor 6, and its lower rotor 2-1 rotates in the opposite direction to the upper rotor 1-1 to counteract the anti-torque. Its pitch is controlled by a lower swashplate 7 independently operated by a second servo motor group 8. In this embodiment, there is no mechanical linkage between the two sets of motors and servo motor groups, and they are also isolated from each other in terms of electrical and control signal paths, thus forming the physical basis for achieving independent control and redundancy backup. The core of the flight controller 9 uses a high-performance microprocessor, which is responsible for running the control algorithm detailed in this invention and simultaneously managing tasks such as sensor data acquisition, communication, and fault monitoring.

[0059] Example 1

[0060] like Figure 3As shown, this invention provides an adaptive reverse cooperative control system for a coaxial dual-rotor aircraft, which mainly comprises: an attitude command acquisition module, a control command calculation module, a motor speed calculation module, a servo control signal conversion module, and a synchronization output module. In addition, the system also includes a fault detection module and a redundant control module.

[0061] Specifically, the attitude command acquisition module is used to receive target attitude commands for the aircraft from the flight control host computer, remote controller, or autonomous mission planning system. These commands include at least a pitch angle command. ), roll angle command ( ), yaw angle command ( ) and vertical speed command ( (), representing the desired motion state of the aircraft.

[0062] The control command calculation module is connected to the attitude command acquisition module and is responsible for calculating the target attitude command into an upper rotor control command group and a lower rotor control command group based on the target attitude command and the current state of the aircraft. This module is the core of the invention and is responsible for realizing the intelligent switching between the reverse control strategy and the conventional unidirectional control strategy. The upper rotor control command group consists of the upper rotor longitudinal periodic pitch angle (…). ), Upper rotor lateral cyclic pitch angle ( ) and collective pitch angle of the upper rotor ( The lower rotor control command group consists of the lower rotor longitudinal periodic pitch angle ( ). ), Lower rotor lateral cyclic pitch angle ( ) and the collective pitch angle of the lower rotor ( ).

[0063] The motor speed calculation module is connected to the control command calculation module and the attitude command acquisition module. Its function is to calculate the motor speed based on the yaw angle command ( Vertical speed command ( ) and the average collective pitch angle of the upper and lower rotors, calculate the target speeds of the first and second motors ( and ).

[0064] The servo control signal conversion module is connected to the control command calculation module. It is used to calculate and convert the upper rotor control command group and the lower rotor control command group into PWM position control signals of the first servo servo group and the second servo servo group through their respective independent inverse kinematic models.

[0065] The synchronous output module connects the motor speed calculation module and the servo control signal conversion module, and is responsible for synchronously outputting the PWM position control signal of the first servo servo group, the PWM position control signal of the second servo servo group, and the target speed command of the first motor. ) and the target speed command for the second motor ( Each actuator is assigned to its respective actuator to drive the aircraft to fly according to instructions.

[0066] To enhance the safety and reliability of the aircraft, this system also includes a fault detection module and a redundancy control module. The fault detection module monitors the operating status of each servo motor in the first and second servo motor groups in real time by connecting to the servo motor groups (usually via sensors), and determines whether a fault exists based on preset fault judgment criteria. Once a servo motor group fault is detected, the redundancy control module is activated and executes redundancy control strategies, such as reallocating attitude control tasks and adjusting motor and servo motor control parameters, to maintain the controllability of the aircraft.

[0067] Example 2

[0068] like Figure 4 As shown, the present invention provides an adaptive reverse cooperative control method for a coaxial dual-rotor aircraft, comprising the following sequential steps:

[0069] S1. Obtaining Target Attitude Commands: This method first obtains the aircraft's target attitude commands through the flight control system. These commands typically originate from the flight control mission planner, remote controller, or autopilot, and specifically include pitch angle commands ( ), roll angle command ( ), yaw angle command ( ) and vertical speed command ( These commands represent the desired motion state of the aircraft and, after attitude calculation and control loop, are converted into internal attitude error or attitude angular velocity commands, which serve as inputs for subsequent calculations.

[0070] S2. The target attitude command is resolved into an upper rotor control command group and a lower rotor control command group: This step is one of the core control functions of this invention, aiming to convert high-level attitude commands into specific control quantities for the upper and lower rotors. For example... Figure 4 As shown, the solution strategy intelligently switches between a reverse cooperative control strategy and a conventional co-directional control strategy based on the amplitude of the target attitude command.

[0071] The upper rotor control command group consists of the upper rotor longitudinal periodic pitch angle ( ), Upper rotor lateral cyclic pitch angle ( ) and collective pitch angle of the upper rotor ( The lower rotor control command group consists of the lower rotor longitudinal cyclic pitch angle ( ). ), Lower rotor lateral cyclic pitch angle ( ) and the collective pitch angle of the lower rotor ( ).

[0072] In order to balance high maneuverability and flight efficiency, such as Figure 5 As shown, this invention employs an adaptive solution strategy: when the target pitch angle command amplitude... Greater than the preset pitch maneuver threshold At that time, or when the target roll angle command magnitude Greater than the preset roll maneuver threshold At this time, the system will determine that the aircraft is in a high-maneuverability demand condition. Under this condition, the system executes a reverse cooperative control strategy, that is, sets the cyclic pitch angles of the upper and lower rotors to change in opposite directions to generate enhanced control torque. Specifically, the longitudinal cyclic pitch angle satisfies the following relationship:

[0073]

[0074]

[0075] The transverse periodic pitch angle satisfies the following relationship:

[0076]

[0077]

[0078] in, and These are the pitch control gain and roll control gain, respectively. These gains can be optimized using the aircraft's dynamic model, combined with the desired control response and stability margin. For example, by analyzing the aerodynamic forces and moments acting on the aircraft, a model relating attitude angles to cyclic pitch angles can be established, and then suitable control gains can be obtained. and By using counter-cyclic pitch changes in the upper and lower rotors, large attitude control moments can be efficiently generated, significantly improving the aircraft's maneuverability, especially suitable for high-dynamic flight missions such as rapid attitude adjustments or obstacle avoidance. Simultaneously, because the cyclic pitch changes of the upper and lower rotors are in opposite directions, the resulting additional lift changes typically cancel each other out, thus avoiding large fluctuations in total lift. The pitch maneuver threshold... and roll maneuver threshold It is set based on the aircraft's design parameters, flight performance requirements, and test data, for example, It can be set to 5°-10°. It can be set from 5° to 10°. When and At the same time, both longitudinal and lateral periodic pitches are set according to the above-mentioned reverse coordination relationship.

[0079] Conversely, when and At this time, the system determines that the aircraft is in normal or stable flight conditions (such as hovering, slow cruise, etc.). In this case, the method employs a conventional unidirectional control strategy, causing the longitudinal cyclic pitch angles of the upper and lower rotors to change in the same direction, and the lateral cyclic pitch angles of the upper and lower rotors to change in the same direction. Specifically, the cyclic pitch angles satisfy the following relationship:

[0080]

[0081]

[0082]

[0083]

[0084] in, and This is a co-directional control gain. This strategy helps reduce aerodynamic interference between the upper and lower rotors, improves flight stability, and reduces system energy consumption. It is suitable for flight conditions where maneuverability requirements are not high and stability and efficiency are more important.

[0085] S3, according to the yaw angle command and vertical speed command Calculate the target speed of the first motor. Target speed of the second motor :

[0086] This step is primarily responsible for the aircraft's overall lift control and yaw control. For example... Figure 6 As shown, the target speed is calculated as follows: First, calculate the reference speed. Its expression is The function here It is a function that maps the vertical speed command and the average collective pitch angle of the upper and lower rotors to the required reference rotational speed. A typical implementation is based on the total lift of the aircraft. With rotational speed collective distance angle Relationship:

[0087]

[0088] in, This is the rotor thrust coefficient. air density, For rotor area, This refers to the rotor radius. It's used to achieve the vertical speed command. It requires providing corresponding lift. To balance gravity and the expected acceleration in the vertical direction ,Right now Therefore, the reference speed It can be derived from the following formula:

[0089]

[0090] in, For the mass of the aircraft, It is the acceleration due to gravity. Indicates vertical speed command The corresponding expected vertical acceleration.

[0091] Next, the speed difference used to generate the yaw control torque is calculated. Its expression is:

[0092]

[0093] in The yaw control gain is determined through modeling and parameter identification of the relationship between the aircraft's counter-torque and the speed difference. Finally, by combining the reference speed and the speed difference, the target speed n_1 of the first motor and the target speed of the second motor are calculated. :

[0094]

[0095]

[0096] This differential adjustment method ensures that while generating the required yaw moment, the aircraft's total lift is primarily generated by... This decision decouples yaw and lift control.

[0097] In engineering implementation, , , The unit can be ; and on , , Limiting is performed to meet the operating range of the motor and electronic speed controller, for example... It can impose a rate of change limit on the speed command to suppress transient shocks.

[0098] In addition, the function It can also be achieved through calibration lookup table method: pre-calibrate different... with average total distance angle Calibration under certain conditions to obtain the reference speed A two-dimensional mapping table is used, and interpolation calculation is performed within the control cycle. Thus, in the rotor thrust coefficient Stable control can still be achieved even when parameters are uncertain.

[0099] S4. Convert the upper rotor control command group and the lower rotor control command group into PWM position control signals for the first servo motor group and the second servo motor group respectively through inverse kinematics calculation:

[0100] This step converts the desired periodic pitch angle and collective pitch angle obtained from S2 calculations into physical angles that the servo motors can execute. Each servo motor group corresponds to a specific rotor control mechanism, and its physical configuration is determined. The inverse kinematics model is based on the desired rotor blade angles (e.g., the collective pitch of the upper rotor blades). and periodic pitch components along the longitudinal and transverse directions. , ), calculate the target position of each servo motor required to drive these blades to the desired angle.

[0101] For a typical three-servo swashplate control mechanism, the first Output angle / displacement of each servo motor The solution can be obtained through geometric relationships:

[0102]

[0103] in, The length of the link. The length of the rocker arm. For the first Installation location parameters for each servo motor. The total distance angle, The longitudinal periodic pitch angle, This is the lateral periodic pitch angle; this function typically involves geometric and trigonometric calculations, and its specific form depends on the mechanical design of the swashplate mechanism. Since the control mechanisms of the upper and lower rotors may differ, there are usually two independent inverse kinematic models, one for the first servo motor group and the other for the second servo motor group.

[0104] In one feasible approach, for a three-servo swashplate mechanism, the desired collective pitch and periodic pitch components can be expressed as a linear combination of servo motor outputs, denoted as the... The target output of each servo motor is Then we have:

[0105]

[0106]

[0107]

[0108] in and The parameters are obtained from the geometric parameters of the swashplate mechanism and the installation calibration; independent parameter sets are used for the upper rotor and the lower rotor to form two independent inverse kinematic models.

[0109] Will When converting to a PWM position control signal, a linear conversion relationship can be used. and to Upper and lower limits and rate of change limits are imposed to meet the servo motor travel and dynamic response requirements.

[0110] S5. Synchronously output the PWM position control signal of the first servo motor group, the PWM position control signal of the second servo motor group, and the target speed command of the first motor. and the target speed command for the second motor :

[0111] During each control cycle, the flight controller simultaneously sends all calculated control commands (servo PWM signals and motor speed commands) to their respective actuators. Synchronous output ensures time consistency between the control variables, which is crucial for stable aircraft control. Motor speed commands are typically sent to the electronic speed controller (ESC) via pulse signals or serial communication protocols, while servos are directly driven via PWM signals.

[0112] S6. Fault Detection Steps: To improve the safety and reliability of the aircraft, this invention also includes a real-time fault detection mechanism. For example... Figure 7 As shown, this step monitors the operating status of each servo motor in the first and second servo motor groups in real time. When any of the following fault judgment criteria are met and confirmed after multiple consecutive control cycles, the system will determine that the corresponding servo motor group has failed, thereby avoiding misjudgments caused by transient interference or sensor noise.

[0113] Specific fault diagnosis criteria include: First, the position deviation continuously exceeds a preset deviation threshold and the duration exceeds a preset time threshold. This is typically determined by monitoring the deviation between the actual position and the commanded position of the servo motor. If the position deviation of a certain servo motor... In continuous Continuously satisfy within each control cycle And the total duration exceeds If the fault is detected, it is determined to be faulty. Secondly, abnormal drive current reflects the workload of the servo motor. Abnormal situations, such as within a preset time window (e.g., 100 ms), will result in an actual current... Overload (e.g.) ), or long-term idle time (such as when there is no corresponding position command). Continue to exceed ), or rate of change of current Mutations (e.g.) These factors can all indicate mechanical or electrical malfunctions. Thirdly, the temperature exceeding the safe threshold, i.e., the internal temperature of the servo motor... Exceeding the preset temperature threshold (e.g., 80°C), or the rate of temperature rise per unit time. Exceeding the preset threshold (e.g., 2°C / s) Overheating is usually a sign of overload, stall, or internal failure.

[0114] S7. Execute Redundancy Control Strategy: Once a fault is detected in any servo motor group in step S6, the system will immediately activate the redundancy control module and execute a series of redundancy control strategies, such as... Figure 8 As shown.

[0115] When a malfunction is detected in the first servo servo group (i.e., a malfunction in the upper rotor control system), the system will set the upper rotor control command group to safe values. For example, the collective pitch angle and cyclic pitch angle of the upper rotor will both be set to preset safe fixed values ​​(e.g., collective pitch is at the minimum allowable value, and cyclic pitch is zero) to prevent the malfunctioning rotor from generating uncontrollable torque or drag. Simultaneously, attitude control tasks will be completely assigned to the second servo servo group, with the healthy lower rotor independently performing all attitude control (pitch, roll) of the aircraft. At this time, enhanced lower rotor control commands need to be calculated because the lower rotor needs to bear a greater control load. The enhanced lower rotor control commands are calculated through gain compensation. Calculations show that Adaptive adjustment based on attitude control error is used to effectively convert the desired attitude control torque into a single rotor control input. For example, if the desired attitude control torque is... The control torque generated by the healthy rotor is ,but . It can be based on attitude error The adaptive gain can be adjusted, for example, designed as follows:

[0116]

[0117] in, As a reference compensation gain, These are adaptive coefficients. When the attitude error... When it is large, Increase to provide a stronger compensating torque; when the attitude approaches stability, Reduce to prevent overshoot. It can be attitude angle error or attitude angular velocity error. Updated in each control cycle, and amplitude limiting is performed, for example Low-pass filtering or setting a dead zone can be used to suppress gain jitter caused by sensor noise.

[0118] Similarly, when a malfunction is detected in the second servo servo group (i.e., a malfunction in the lower rotor control system), the system will set the lower rotor control command group to a safe value and completely assign the attitude control task to the first servo servo group, which will then be independently handled by the healthy upper rotor. The enhanced upper rotor control commands will also be adaptively compensated for gain. Calculated.

[0119] Furthermore, in redundant control mode, this invention also includes a dynamic motor power allocation strategy under fault mode. When one servo motor fails, to avoid motor overload, save energy, and prevent further damage to the failed component, the system will reduce the output power of the failed motor, for example, by reducing the maximum allowable speed of the failed motor. Less than the maximum speed during normal operation To ensure total lift and maintain attitude control of the aircraft, the healthy side motors will increase their output power, for example, by increasing the maximum permissible speed of the healthy side motors. greater than or equal to the normal maximum speed Under total power constraints, more available power is allocated to the motors on the healthy side. This strategy ensures that, under power-limited conditions, the controllability of the healthy side is prioritized, maintaining the controllability of the aircraft.

[0120] Finally, the redundancy control strategy also includes a safe return-to-home or landing mode. In redundancy control mode, since the aircraft is in a damaged but still controllable state, to maximize safety, the system will limit flight speed and yaw rate, reducing the requirements on the control system's response speed and torque. For example, the target horizontal speed command... Limit to a preset safety threshold and the target yaw rate command Limit to a preset safety threshold Meanwhile, the flight control system will automatically plan a path to guide the aircraft back to a preset safe landing point or the nearest open area for an emergency landing. This process usually takes into account the aircraft's current position, remaining battery power, and terrain information, thereby improving the mission success rate and equipment recovery rate.

[0121] This invention is not limited to the embodiments described above. Those skilled in the art can make various modifications or variations without departing from the spirit and scope of this invention. For example, the specific threshold, gain, and function form can be adjusted according to different aircraft models and mission requirements. The scope of protection of this invention is defined by the appended claims.

Claims

1. An adaptive reverse cooperative control method for a coaxial dual-rotor aircraft, characterized in that, Suitable for coaxial twin-rotor aircraft with independently driven first and second motors, and independently controlled upper and lower rotor control mechanisms, including: S1. Obtain target attitude commands, including pitch angle command θ_cmd and roll angle command. Yaw angle command and vertical speed command ; S2. The target attitude command is calculated into an upper rotor control command group and a lower rotor control command group, wherein the upper rotor control command group includes the upper rotor longitudinal cyclic pitch angle. Lateral cyclic pitch angle of the upper rotor and rotor collective pitch angle The lower rotor control command group includes the lower rotor longitudinal periodic pitch angle. Lateral cyclic pitch angle of the lower rotor and the collective pitch angle of the lower rotor The solution strategy is as follows: when When, set and ; when When, set and ; when and At that time, the conventional co-directional control strategy is adopted to make the longitudinal periodic pitch angle of the upper and lower rotors change in the same direction and the lateral periodic pitch angle change in the same direction. in and These are the preset pitch maneuver threshold and roll maneuver threshold, respectively. and These are the pitch control gain and roll control gain, respectively, and when and Simultaneously execute the above-mentioned reverse coordination settings for pitch and roll; S3, according to the yaw angle command and vertical speed command Calculate the target speed of the first motor Target speed of the second motor The calculation method is as follows: Reference speed Speed ​​difference ,but , ,in Yaw control gain; S4. Convert the upper rotor control command group and the lower rotor control command group into PWM position control signals for the first servo motor group and the second servo motor group respectively through inverse kinematics calculation; S5. Synchronously output the PWM position control signal of the first servo motor group, the PWM position control signal of the second servo motor group, and the target speed command of the first motor. and the target speed command for the second motor .

2. The cooperative control method for a coaxial dual-rotor aircraft according to claim 1, characterized in that, The conventional co-directional control strategy is to make the longitudinal periodic pitch angles of the upper and lower rotors change in the same direction, and also to make the lateral periodic pitch angles of the upper and lower rotors change in the same direction.

3. The cooperative control method for a coaxial dual-rotor aircraft according to claim 1, characterized in that, In step S3, the target speed of the first motor is... Second motor target speed The calculation method is as follows: Reference speed ; Speed ​​difference ;but ; in, This is the yaw control gain.

4. The cooperative control method for a coaxial dual-rotor aircraft according to claim 1, characterized in that, It also includes fault detection step S6: Real-time monitoring of the operating status of each servo motor in the first and second servo motor groups; when any of the following fault judgment criteria are met, the corresponding servo motor group is determined to be faulty: (a) The position deviation continuously exceeds the preset deviation threshold and the duration exceeds the preset time threshold; (b) Abnormal drive current; (c) The temperature exceeds the safety threshold; the fault determination needs to be confirmed through multiple consecutive control cycles.

5. The cooperative control method for a coaxial dual-rotor aircraft according to claim 4, characterized in that, After the fault detection, the following redundancy control strategy S7 is also executed: S7.1 When a fault is detected in the first servo motor group, the upper rotor control command group is set to a safe value, and the attitude control task is completely assigned to the second servo motor group, and the enhanced lower rotor control command is calculated. S7.2 When a fault is detected in the second servo motor group, the downrotor control command group is set to a safe value, the attitude control task is completely assigned to the first servo motor group, and the enhanced uprotor control command is calculated.

6. The cooperative control method for a coaxial dual-rotor aircraft according to claim 5, characterized in that, The enhanced uprotor control command or downrotor control command is achieved through gain compensation. The calculation shows that the Based on attitude control error Adaptive adjustment.

7. The cooperative control method for a coaxial dual-rotor aircraft according to claim 5, characterized in that, The redundancy control strategy S7 also includes a dynamic motor power allocation strategy under fault modes: When one side of the servo motor fails, the output power of the motor on the failed side is reduced, while the output power of the motor on the healthy side is increased.

8. The cooperative control method for a coaxial dual-rotor aircraft according to claim 5, characterized in that, The redundant control strategy S7 also includes a safe return or landing mode S8: In redundant control mode, the flight speed and yaw rate are limited, and a path is automatically planned to return to a safe landing point.

9. A control system for a coaxial dual-rotor aircraft, characterized in that, For implementing the control method according to any one of claims 1 to 8, comprising: The attitude command acquisition module is used to acquire the target attitude command; The control command calculation module is used to calculate the target attitude command into an upper rotor control command group and a lower rotor control command group, and can switch between the reverse cooperative control strategy and the conventional co-directional control strategy. The motor speed calculation module is used to calculate the target speeds of the first motor and the second motor. The servo control signal conversion module is used to convert the control command group into the corresponding servo servo group PWM position control signal; The synchronous output module is used to simultaneously output the PWM position control signal and the motor target speed command.

10. The coaxial dual-rotor aircraft control system according to claim 9, characterized in that, Also includes: The fault detection module is used to monitor the working status of the servo motor assembly in real time and determine faults according to preset fault judgment criteria. A redundancy control module is used to execute a redundancy control strategy when a servo motor failure is detected.

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