A rotor tilting flight control method for a multi-tilting rotor aircraft in a hovering low-speed state based on redundant variable-distance operation

Abstract

The application discloses a kind of based on redundancy variable-pitch control's multi-inclined rotorcraft hover low-speed state rotor tilt flight control method, belong to the technical field of control, regulation.The application is by the periodic variable-pitch control of multiple groups of rotors to complete the 90° tilt of rotorcraft in hover or low-speed flight state rotor, so that the aircraft is upright, then realizes the quick transition flight from helicopter flight mode to fixed-wing aircraft flight mode in tail stand mode.It specifically includes: redundant actuator control distribution strategy based on the geometric relationship of operating component, parameter self-adaptive proportional-integral-derivative second-order position outer loop controller, acceleration feedback expected attitude solving architecture, and parameter simplified proportional-integral first-order attitude and angular velocity double-loop controller.The application realizes the decoupling of rotor tilt dynamic process and speed dynamic change, expands the selection range of take-off and landing platform, avoids reverse tilt when high-speed flight mode changes to low-speed flight mode, and significantly improves the balance control ability and safety of aircraft.

Description

Technical Field

[0001] This invention relates to aircraft flight control technology, and discloses a flight control method for a multi-tilt rotor aircraft to achieve full rotor tilt during hovering or low-speed flight, belonging to the technical field of control and regulation. Background Technology

[0002] Tiltrotor aircraft combine the vertical takeoff and landing, hovering, and low-speed flight capabilities of rotorcraft with the high-speed cruise and long-range flight capabilities of fixed-wing aircraft, thus showing promising application prospects in emergency rescue, logistics transportation, inspection and reconnaissance, and special operations. With the development of unmanned aerial vehicle (UAV) technology, related application scenarios place higher demands on the flight safety, mission adaptability, and stability control capabilities across the entire flight envelope of tiltrotor aircraft.

[0003] In existing two-tilt rotorcraft technology, rotor tilting requires meeting speed variation requirements, which severely affects model uncertainty and jeopardizes flight safety. Four-tilt and higher-tilt rotorcraft have multiple rotors and multiple sets of cyclic pitch control, offering high control redundancy and diverse control methods. Through the coordinated control of multiple rotors, aircraft attitude adjustment and external force balance can be achieved, thereby reducing the dependence of the tilting process on forward flight speed and making rotor tilting possible in hovering or low-speed flight states.

[0004] In existing I-shaped quad-tilt rotor aircraft technology, the specific structural distribution of the rotors in the longitudinal and lateral directions of the fuselage provides greater force and torque adjustment space during tilting, which is beneficial for achieving balance control during rotor tilting. This specific structural layout of multiple rotors offers the following possibilities: decoupling of the rotor tilting dynamic process and speed dynamic changes, improving flight safety; allowing takeoff and landing on steep slopes, considering that the takeoff and landing points do not need to be level, thus expanding new takeoff and landing platforms; and eliminating the need for reverse rotor tilting when transitioning from a fully tilted high-speed flight mode to a low-speed flight mode. This invention arose against this backdrop. Summary of the Invention

[0005] The purpose of this invention is to address the shortcomings of the aforementioned background technology by providing a rotor tilting flight control method for multi-tilt rotor aircraft in hovering and low-speed states based on redundant variable pitch control. This method solves the technical problem of flight insecurity caused by the coupling of the rotor tilting dynamic process and speed dynamic changes in tilt rotor UAVs. It enables the transition from high-speed flight mode to low-speed flight mode without rotor tilting, thereby improving flight reliability and enhancing the flight safety and mission adaptability of tilt rotor aircraft in hovering and low-speed flight states.

[0006] To achieve the above-mentioned objectives, the present invention employs the following technical solution:

[0007] The first aspect of the present invention provides a method for controlling the tilting flight of a multi-tilt rotorcraft in a hovering low-speed state based on redundant variable-pitch control, comprising:

[0008] S1. Obtain the current position, attitude, angular velocity, acceleration, and rotor tilt angle of the multi-tilt rotor aircraft;

[0009] S2. A parameter-adaptive proportional-integral-derivative second-order position outer loop controller is used to calculate the position outer loop control quantity based on the desired position trajectory.

[0010] S3. Solve for the desired attitude angle and the nominal term of the total force control channel quantity based on the position outer loop control quantity and acceleration feedback. Perform adaptive compensation based on acceleration feedback on the nominal term of the total force control channel quantity to obtain the total force control channel quantity.

[0011] S4. A parameter adaptive proportional-integral first-order attitude and angular velocity dual-loop controller is used to estimate the desired angular velocity and angular velocity control quantity based on the desired attitude angle.

[0012] S5 performs redundant execution control allocation based on the geometric relationship of the control components for the angular velocity control quantity and the total force control channel quantity, and obtains the actual actuation quantity of the rotor control actuator, including the collective pitch control quantity, the lateral cyclic pitch control quantity, and the longitudinal cyclic pitch control quantity.

[0013] As a further optimization of the rotor tilt flight control method for low-speed hovering of multi-tilt rotorcraft based on redundant variable pitch control, in S2, the parameter adaptive proportional-integral-derivative second-order position outer loop controller includes a roll channel outer loop controller, a pitch channel outer loop controller, and a yaw channel outer loop controller. The control law of the roll channel outer loop controller is... , ,in, This is the outer loop control value for the position of the roll channel. For the parameters of the roll controller, The desired position of the roll channel. This represents the actual position of the rolling channel.

[0014] As a further optimization scheme of the rotor tilting flight control method for multi-tilt rotorcraft hovering at low speeds based on redundant variable pitch control, S3 solves for the desired attitude angle and the nominal terms of the total force control channel quantity based on the position outer loop control quantity and acceleration feedback, specifically:

[0015] The desired resultant force in the ground coordinate system is calculated based on the position outer loop control variable and the acceleration feedback compensation term. , , ,in, To represent the desired resultant force in the ground coordinate system, For the mass of the aircraft, This is the outer loop control value for the pitch channel position. This refers to the outer ring control quantity for the yaw channel position. It is the acceleration due to gravity. This is the rotation matrix for transforming from the body coordinate system to the ground coordinate system. The rotation transformation matrix required to achieve the outer loop position control is as follows: To represent the desired resultant force in the body coordinate system, The nominal value for the total force control channel quantity. For the acceleration of the pitch and yaw channels, This refers to the rotor tilt angle;

[0016] Based on the desired heading angle Calculate the rotational transformation matrix required to implement the outer loop control quantity. , , ;

[0017] The desired roll angle of the aircraft is solved in the body coordinate system based on the desired resultant force direction. and desired pitch angle , , .

[0018] As a further optimization scheme for the rotor tilt flight control method of multi-tilt rotorcraft hovering at low speeds based on redundant variable pitch control, S3 performs adaptive compensation based on acceleration feedback for the nominal term of the total force control channel quantity, specifically as follows: ,in, This is an adaptive compensation term for the total force control channel quantity. This is the gain coefficient for compensation in the direction of total force. The acceleration of the rolling channel; the total force obtained controls the channel volume. for .

[0019] As a further optimization of the rotor tilt flight control method for low-speed hovering of multi-tilt rotorcraft based on redundant variable pitch control, S4 incorporates a parameter-adaptive proportional-integral first-order attitude and angular velocity dual-loop controller, including a roll channel dual-loop controller, a pitch channel dual-loop controller, and a yaw channel dual-loop controller. The roll channel dual-loop controller comprises an outer loop for roll channel attitude angle and an inner loop for roll channel angular velocity. The control law for the outer loop of the roll channel attitude angle is... The inner loop control law for the roll channel angular velocity is ,in, For the desired roll rate, This is the actual roll angular velocity. For the desired roll angle, This is the actual roll angle. These are the outer loop control parameters for the roll channel attitude angle. These are the inner loop control parameters for the angular velocity of the roll channel.

[0020] As a further optimization of the rotor tilt flight control method for hovering low-speed states of multi-tilt rotorcraft based on redundant variable pitch control, S5 performs redundant execution control allocation of angular velocity control quantity and total force control channel quantity based on the geometric relationship of the control components. Specifically, the control channel quantity output by the flight controller is allocated as linkage control quantity according to the rotor tilt angle, and the linkage control quantity is allocated to the actual actuation quantity of the rotor control actuator.

[0021] As a further optimization of the rotor tilt flight control method for multi-tilt rotorcraft hovering at low speeds based on redundant variable pitch control, the control channel quantities output by the flight controller are allocated as linkage control quantities according to the rotor tilt angle, specifically: ,in, This is the total distance linkage. This is the total distance difference between the front and rear ends. The total distance difference between the left and right sides is the momentum. This refers to the longitudinal periodic variable pitch linkage. This refers to the longitudinal periodic variable pitch differential momentum. The left and right longitudinal periodic variable pitch differential momentum. This refers to the lateral periodic variable pitch linkage. This refers to the forward and backward lateral periodic pitch differential momentum. It is the left and right lateral periodic variable pitch differential momentum. To control the distribution gain parameters, To control the amount of the rollover channel Distributed to left and right total pitch differential Dynamic gain, , To control the amount of the rollover channel Distributed to left and right longitudinal periodic variable pitch differential Dynamic gain, , To adjust the pitch control channel quantity Distributed to front and rear total pitch differential Dynamic gain, , To adjust the pitch control channel quantity Distributed to longitudinal periodic variable pitch linkage Dynamic gain, , To adjust the pitch control channel quantity Distributed to longitudinal periodic differential momentum Dynamic gain, .

[0022] As a further optimization of the rotor tilting flight control method for multi-tilt rotorcraft hovering at low speeds based on redundant variable pitch control, the linkage control quantity is distributed to the actual actuator momentum of each rotor, specifically as follows: ,in, This refers to the collective pitch control of rotors 1 through 4. This refers to the lateral cyclic pitch control values ​​for rotors 1 through 4. This refers to the longitudinal cyclic pitch control parameters for rotors 1 through 4.

[0023] A second aspect of the present invention provides a flight controller including a memory and a processor, wherein the memory stores a computer program that runs on the processor, and the processor executes the steps of the control method described above when running the computer program.

[0024] A third aspect of the present invention provides a multi-tilt rotor aircraft, comprising: a fuselage and at least four sets of tiltable rotors disposed on the fuselage, rotor control actuators, flight status sensors, and the aforementioned flight controller; the flight status sensors are used to acquire position, attitude, angular velocity, acceleration, and rotor tilt angle information of the multi-tilt rotor aircraft; the flight controller is used to allocate actual actuation force of the rotor control actuators according to the real-time position, attitude, angular velocity, acceleration, and rotor tilt angle of the multi-tilt rotor aircraft; the rotor control actuators are used to drive each set of tiltable rotors to perform collective pitch control, longitudinal cyclic pitch control, lateral cyclic pitch control, and tilt control according to the actual actuation force.

[0025] The present invention, by adopting the above technical solution, has the following beneficial effects:

[0026] 1. Achieves 90° rotor tilt in hover: This invention proposes a flight control method for a four-tilt rotor UAV with redundant cyclic pitch control capabilities to achieve 90° rotor tilt in hover. By utilizing the redundant control capabilities among multiple control variables such as cyclic pitch, collective pitch, rotational speed, and tilt angle, the aircraft can maintain attitude and position stability during rotor tilt, thereby achieving hover tilt flight that is difficult to achieve with traditional tilt rotor UAVs.

[0027] 2. Expanded flight envelope of tilt-rotor UAVs: Many existing four-tilt rotor configurations typically only have variable collective pitch or variable rotation speed capabilities, lacking cyclic pitch control of the rotor, making it difficult to achieve large angles or even 90° tilts under hovering conditions; This invention introduces and fully utilizes cyclic pitch redundancy control, enabling the rotor to maintain the hovering and controllable flight capabilities of the aircraft after completing a 90° tilt, significantly expanding the flight modes and attitude working range of tilt-rotor UAVs.

[0028] 3. Achieves stable hovering after the rotor is fully tilted: This invention considers the aerodynamic and torque distribution characteristics after the rotor is fully tilted to 90°. Through corresponding control distribution algorithms and controller design, the aircraft can still achieve stable and reliable "upright hovering" after the rotor is fully tilted, breaking through the limitation that existing tilt-rotor UAVs usually cannot maintain stable hovering in the state of full rotor tilt.

[0029] 4. Improved flight stability and reliability at different tilt angles: The control allocation algorithm proposed in this invention can allocate control quantities in real time according to the tilt angle of each rotor, so that the aircraft has stable attitude control and anti-disturbance capabilities throughout the entire process of the rotor gradually tilting from a horizontal hovering state to 90°, thereby improving flight safety and control reliability during the tilt transition process.

[0030] 5. Control advantages of multiple flight configurations: After the rotor tilts to 90°, the aircraft can form a flight attitude similar to a tail-seat aircraft, and can perform forward flight, side flight and other flight maneuvers according to the flight control method of the tail-seat configuration. Therefore, this invention is not only applicable to the hovering and tilting control of four-tilt rotor UAVs, but also provides a technical basis for its control in multiple flight modes such as upright hovering, forward flight, side flight and transition flight. Attached Figure Description

[0031] Figure 1 This is a simplified schematic diagram showing the dimensional relationships of the control components of an I-shaped quad tiltrotor aircraft with cyclic pitch control provided in an embodiment of the present invention.

[0032] Figure 2 This is a schematic diagram of the control arm provided in an embodiment of the present invention.

[0033] Figure 3 This is a schematic diagram of the x-channel parameter adaptive proportional-integral-differential second-order position outer loop controller provided in this embodiment of the invention.

[0034] Figure 4 This is a simplified schematic diagram of the proportional-integral first-order attitude and angular velocity dual-loop controller for the roll channel provided in this embodiment of the invention.

[0035] Figure 5This is a schematic diagram of the hovering and tilting control architecture of an I-shaped four-tilt rotor aircraft with rotor periodic pitch control provided in an embodiment of the present invention.

[0036] Figure 6 This is a schematic diagram of the attitude control process of a four-tilt rotor aircraft with an I-shaped layout and rotor periodic pitch control provided in an embodiment of the present invention during a 90° tilt of the rotor in a forward hovering state.

[0037] Figure 7 This is a waveform diagram of the desired instruction designed in an embodiment of the present invention.

[0038] Figures 8(a), 8(b), and 8(c) show the control strategies proposed in this invention and traditional PID control strategies provided in the embodiments of this invention. aisle, Channels and Comparison chart of simulation results for channel position tracking curves.

[0039] Figures 9(a), 9(b), and 9(c) show the control strategies proposed in this invention and traditional PID control strategies provided in the embodiments of this invention. aisle, Channels and Comparison of simulation results of channel velocity response curves.

[0040] Figures 10(a), 10(b), and 10(c) show the expected roll angles under the control proposed in this invention and the traditional PID control strategy provided in the embodiments of this invention. Desired pitch angle and desired yaw angle Comparison of simulation results of response curves.

[0041] Figures 11(a), 11(b), and 11(c) show the roll angular velocity under the control proposed in this invention and the traditional PID control strategy provided in the embodiments of this invention. Desired pitch angular velocity and desired yaw rate Comparison of simulation results of response curves. Detailed Implementation

[0042] The present invention will be further described below with reference to the accompanying drawings and embodiments. It should be understood that the following embodiments are only used to illustrate the technical solutions of the present invention, and are not intended to limit the scope of protection of the present invention. All equivalent substitutions or modifications made under the concept of the present invention should fall within the scope of protection of the present invention.

[0043] A multi-tilt rotor aircraft comprises an airframe, multiple sets of tiltable rotors, rotor control actuators, flight status sensors, and a flight controller. Each set of tiltable rotors is capable of collective pitch control, longitudinal cyclic pitch control, lateral cyclic pitch control, and tilt control. The flight status sensors acquire flight status information such as position, velocity, attitude angle, angular velocity, and acceleration. The flight controller generates control commands based on the flight status information and drives the tiltable rotors to move through the rotor control actuators.

[0044] In this embodiment, the multi-tilt rotor aircraft is preferably a four-tilt rotor aircraft, with four sets of tiltable rotors mounted on the fuselage, preferably in an I-shape configuration. The four sets of tiltable rotors are respectively installed at corresponding positions on the front, rear, and left and right sides of the fuselage, providing sufficient force and torque adjustment space during rotor tilting. The tilt angles of the four sets of tiltable rotors change synchronously, meaning the tilt angles of the four sets of rotors satisfy the following:

[0045]

[0046] In equation (1), The rotor tilt angle, These represent the tilt angles of the four tiltable rotors.

[0047] In this embodiment, the rotational speed of the four tilting rotors can be kept constant. The aircraft mainly achieves external force balance, attitude adjustment, and torque control during rotor tilting through collective pitch control, longitudinal cyclic pitch control, and lateral cyclic pitch control. By employing these methods, the dependence of the rotor tilting process on forward flight speed can be reduced, enabling the multi-tilt rotor aircraft to complete a 90° rotor tilt in hovering or low-speed flight.

[0048] This embodiment provides a method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering mode based on redundant variable-pitch control. The method includes: establishing a redundant actuator control allocation strategy based on the geometric layout of the control components; establishing a parameter-adaptive proportional-integral-derivative second-order position outer loop controller; designing an acceleration feedback desired attitude solution architecture; and designing a parameter-adaptive proportional-integral first-order attitude and angular velocity dual-loop controller. The control architecture of this method is as follows: Figure 5 As shown.

[0049] Establish a redundant actuator control allocation strategy based on the geometric relationship of the manipulated components:

[0050] Obtain the geometric parameters of the control components of a quad tiltrotor aircraft. For example... Figure 2 As shown, the geometric parameters of the control components of a four-tilt rotor aircraft include: rotor nacelle length. Vertical distance from the rotor tilt axis to the center of gravity of the aircraft Longitudinal distance from the nose rotor to the center of gravity of the aircraft and the longitudinal distance from the rear rotor to the center of gravity of the aircraft .

[0051] A four-tilt rotor aircraft, due to different rotor tilt angles... Figure 2 The change in lever arm leads to a change in control effectiveness. Each rotor group has 5 control parameters. The control parameters of a quadcopter can be written as: collective pitch Longitudinal periodic torque Lateral periodic pitch Rotation speed tilt angle Since the rotational speed is controlled to be constant, each rotor group has four main control parameters, for a total of 16 rotor control parameters. Collective pitch actuation causes... Figure 1 The tension shown Longitudinal periodic torque actuation caused Figure 1 The backforce shown Lateral periodic torque actuation caused Figure 1 The lateral force shown .

[0052] Because the rotor thrust is during the tilting process... Direction and lever arm All vary with tilt angle As the force and torque change, the same control input will produce different forces and torques at different tilt angles. The lever arm that generates torque for the simultaneous longitudinal cyclic torque control of the first and second rotor groups. The lever arm that generates torque for the simultaneous longitudinal cyclic torque control of the third and fourth rotor groups. The lever arm that generates torque for the simultaneous collective pitch control of the first and second rotor groups. This refers to the lever arm that generates torque for the simultaneous collective pitch control of the third and fourth rotor groups. To compensate for changes in control effectiveness during rotor tilting, this embodiment establishes a redundant actuator control allocation strategy related to the rotor tilting angle.

[0053] The flight controller outputs four control channel values, which are denoted as follows: ,in, To control the amount of rolling control channels, To control the pitch channel volume, Yaw control channel quantity. The amount of total force or lift is controlled by the channel quantity.

[0054] In order to utilize the redundant control capabilities of multiple rotors, the control distribution is divided into two layers.

[0055] The first level of control allocation is used based on the tilt angle. The four control channels output by the flight controller It is allocated into 9 linkage control variables. These 9 linkage control variables include: total distance linkage variable. Total distance differential Left and right total distance differential momentum Longitudinal periodic variable pitch linkage longitudinal periodic variable pitch differential momentum 1. Left and right longitudinal periodic variable pitch differential momentum Lateral periodic variable pitch linkage Lateral lateral periodic differential momentum and the left and right lateral periodic variable pitch differential momentum .

[0056] In the first-level control allocation, a dynamic gain parameter that varies with the rotor tilt angle is introduced. The dynamic gain parameter may include... and Functional terms of the form are used to describe the changes in control effectiveness caused by changes in rotor tilt angle. Through dynamic gain parameters, the flight controller can adjust the allocation ratio of each control channel in real time as the rotor tilts from a vertical lift state to a 90° state, so as to keep the aircraft's attitude stable and force balanced.

[0057] The mathematical expression for the first-level control allocation is as follows:

[0058]

[0059] In equation (2), To control the distribution gain parameter, it is set to a constant value; To control the amount of the rollover channel Distributed to left and right total pitch differential Dynamic gain, ; To control the amount of the rollover channel Distributed to left and right longitudinal periodic variable pitch differential Dynamic gain, ; To adjust the pitch control channel quantity Distributed to front and rear total pitch differential Dynamic gain, ; To adjust the pitch control channel quantity Distributed to longitudinal periodic variable pitch linkage Dynamic gain, ; To adjust the pitch control channel quantity Distributed to longitudinal periodic differential momentum Dynamic gain, .

[0060] The second-level control allocation is used to further distribute the aforementioned multiple linkage control quantities to the actual actuator motions of each rotor group. Specifically, the second-level control allocation will allocate the collective pitch linkage quantity... and total distance differential momentum , Collective pitch control allocated to the four rotors ; to adjust the longitudinal periodic pitch linkage and longitudinal periodic pitch differential , Longitudinal cyclic pitch control values ​​allocated to the four rotors ; to adjust the lateral periodic pitch linkage and transverse periodic differential momentum , Lateral cyclic pitch control values ​​allocated to the four rotors .

[0061] The mathematical expression for the second-level control allocation is as follows:

[0062]

[0063] By decoupling the rotor tilting dynamic process from the flight speed dynamic change through the above two-layer control distribution, the aircraft can make full use of the redundant control capabilities of each rotor group during the rotor tilting process, so that collective pitch control, longitudinal cyclic pitch control and lateral cyclic pitch control are coordinated with each other, thereby achieving force and torque balance during the tilting process.

[0064] Establish a parameter-adaptive proportional-integral-derivative second-order position outer-loop controller:

[0065] After completing the design of the redundant actuator control allocation strategy, a position outer loop controller is established. The position outer loop controller is used to generate the position outer loop control quantity based on the aircraft's desired position trajectory and actual position feedback.

[0066] Let the desired location trajectory be:

[0067]

[0068] In equation (4), These represent the desired positions of the aircraft in the roll, pitch, and yaw directions, respectively. Desired position trajectory. It is a second-order differentiable trajectory.

[0069] Assume the actual position of the aircraft for:

[0070]

[0071] Equation (5), These represent the actual positions of the aircraft in the three control channels: roll, pitch, and yaw.

[0072] Position outer loop controllers are respectively for aisle, Channels and The channel establishes a control law. Taking the channel as an example, the controller according to Generation of desired position, actual position, integral term of position error, and differential term of position error Directional position outer loop control quantity . Channels and Channels are generated using the same structure and .

[0073] The channel control law is designed as follows:

[0074]

[0075]

[0076] In equations (6) and (7), yes Similarly, the channel controller parameters also include... Channels and The channel controllers, they constitute as follows Figure 3 The outer loop controller is shown.

[0077] In this embodiment, the position outer loop controller is a parameter adaptive proportional-integral-derivative second-order position outer loop controller. An error integral term is introduced into its control law. Sum of error differential terms Construct adaptive parameter variables. By using adaptive parameter variables, the controller can adjust the control parameters in real time according to changes in flight status, thereby improving the robustness of position control and reducing the impact of model uncertainty, external disturbances, and rotor tilting processes on position control accuracy.

[0078] Position outer loop controller output position outer loop control quantity Used for subsequent desired attitude calculation.

[0079] Establish an architecture for solving the desired attitude based on acceleration feedback:

[0080] During rotor tilting, solving for the desired attitude solely based on position error is susceptible to model uncertainties and changes in tilt angle. To improve the accuracy of attitude calculation, this embodiment further introduces aircraft acceleration feedback, establishing a desired attitude solution architecture based on acceleration feedback.

[0081] Let the acceleration feedback quantity of the aircraft be... Based on the location, the outer loop control quantity Combining the gravitational acceleration term and the acceleration feedback compensation term, the desired resultant force is calculated in the ground coordinate system. The desired resultant force represents the direction and magnitude of the force required for the aircraft to achieve position control at the current tilt angle. The desired resultant force in the ground coordinate system is expressed as follows:

[0082]

[0083]

[0084] In equations (8) and (9), To represent the desired resultant force in the ground coordinate system, For the mass of the aircraft, It is the acceleration due to gravity. This is the rotation matrix for transforming from the body coordinate system to the ground coordinate system. The rotation transformation matrix required to achieve the outer loop position control is as follows: To represent the desired resultant force in the body coordinate system, The nominal item for the total force or lift control channel quantity.

[0085] Total force or lift control channel quantity includes nominal items and adaptive compensation terms ,

[0086]

[0087]

[0088]

[0089] In equation (11), This is the gain coefficient for compensation in the direction of total force.

[0090] Subsequently, the desired resultant force is transformed from the ground coordinate system to the body coordinate system using a coordinate transformation matrix. The coordinate transformation matrix required to achieve the outer loop position control is determined based on the aircraft's current attitude angles. , Used to describe the transformation relationship between the aircraft coordinate system and the ground coordinate system. Based on the desired heading angle. Calculated Determine the rotation transformation matrix required to implement the outer loop control quantity. :

[0091]

[0092]

[0093] In the body coordinate system, the desired roll angle of the aircraft is solved based on the desired resultant force direction. and desired pitch angle :

[0094]

[0095]

[0096] In this embodiment, nominal item Determined based on the desired resultant force modulus, coordinate transformation matrix, and rotor tilt angle; adaptive compensation term. Based on the total force direction compensation gain coefficient and Channel acceleration feedback integral Determined. By introducing an adaptive compensation term. Calculate the total force or lift control channel quantity This can reduce the deviation between the actual acceleration and the desired acceleration, making the desired attitude solution more adaptable to the dynamic changes during the rotor tilting process.

[0097] Therefore, when the aircraft tilts its rotors 90° while hovering or flying at low speed, it can still correct the desired attitude in real time based on acceleration feedback, thus avoiding attitude divergence or position drift caused by tilting.

[0098] Establish a parameter-adaptive proportional-integral first-order attitude and angular velocity dual-loop controller:

[0099] To obtain the desired roll angle and desired pitch angle Then, combined with the desired yaw angle A dual-loop controller for attitude and angular velocity is established. The dual-loop controller consists of an outer loop for attitude angle and an inner loop for angular velocity.

[0100] The attitude angle outer loop is used to generate the desired angular velocity based on the error between the desired attitude angle and the actual attitude angle. Specifically, the desired attitude angle includes the desired roll angle. Desired pitch angle and desired yaw angle Actual attitude angles include actual roll angles. Actual pitch angle and actual yaw angle .

[0101] Taking the roll channel as an example, the attitude angle outer ring is based on the desired roll angle. Compared with the actual roll angle The error between the two values, and the integral term of that error, are used to generate the desired roll rate. The pitch and yaw channels generate the desired pitch angular velocity, respectively. and desired yaw rate The outer loop control law for the roll channel attitude angle is:

[0102]

[0103] In equation (17), These are the outer loop control parameters for the roll channel attitude angle.

[0104] The inner angular velocity loop is used to generate the angular velocity control value based on the error between the desired angular velocity and the actual angular velocity. Taking the roll channel as an example, the inner angular velocity loop generates the control value based on the desired roll angular velocity. Compared with actual roll angular velocity The error between them, and the integral term of that error, are used to generate the roll control quantity. The pitch and yaw channels use the same control structure to generate corresponding control variables. The inner loop control law for the roll channel angular velocity is:

[0105]

[0106] In equation (18), These are the inner loop control parameters for the angular velocity of the roll channel.

[0107] like Figure 4 As shown, the attitude and angular velocity dual-loop controller adopts a parameter-adaptive proportional-integral first-order control structure. This structure improves the controller's adaptability to rotor tilting, aerodynamic parameter changes, and external disturbances while ensuring simple implementation and low computational complexity.

[0108] The dual-loop controller ultimately outputs attitude control and angular velocity control quantities, which are used as inputs for control allocation to drive each group of rotors to perform collective pitch control, longitudinal cyclic pitch control, and lateral cyclic pitch control.

[0109] In one embodiment of the present invention, a method for controlling the 90° tilt of a rotor in a low-speed hovering state of a multi-tilt rotorcraft based on redundant pitch control is provided, which specifically includes the following four processes.

[0110] When the aircraft is hovering or flying at low speed, the flight controller first obtains the aircraft's current position. The flight controller then collects status information such as attitude, angular velocity, acceleration, and rotor tilt angle. Following this, the flight controller determines the desired position and trajectory. Calculate the outer loop control quantity of the position. And based on acceleration feedback Solve for the desired attitude angle .

[0111] During rotor tilting, the tilt angle The angle gradually changes from the initial angle to 90°. The redundancy actuator control allocation strategy adjusts the collective pitch control of each rotor in real time according to the changes. Longitudinal periodic pitch control amount and lateral periodic pitch control amount Because a dynamic gain parameter related to the tilt angle is introduced into the control allocation, it is possible to compensate for changes in control effectiveness at different tilt angles.

[0112] As the rotor gradually tilts, the direction of rotor thrust changes from primarily providing lift along the vertical direction to gradually shifting to thrust with a larger horizontal component. At this point, the aircraft generates additional force and torque through the periodic pitch control of multiple rotors to maintain the aircraft's attitude stability and external force balance.

[0113] Throughout the tilting process, the outer loop position controller is used to suppress the aircraft's position deviation; the acceleration feedback desired attitude solution architecture is used to correct the desired attitude; the attitude and angular velocity dual-loop controller is used to ensure the stability of the aircraft's attitude angle and angular velocity; and the redundant actuator control allocation strategy is used to rationally distribute the control quantity to each rotor control actuator.

[0114] Through the aforementioned control process, the aircraft can complete a 90° rotor tilt while hovering or in low-speed flight without relying on high forward speed, avoiding reverse rotor tilting during the transition from high-speed to low-speed flight modes. After completing the 90° rotor tilt, the multi-tilt rotor aircraft transitions from helicopter flight mode to fixed-wing flight mode in a tail-seat manner.

[0115] The following is combined Figures 7 to 1 1. This invention demonstrates the effectiveness and superiority of the control scheme proposed in this invention. A design was developed... Figure 7 Expected instructions, including: expected position trajectory Desired attitude angle Desired rotor tilt angle ,exist During this period, the rotor tilt angle changes from 90° to 0°, as... Figure 6 This is a schematic diagram of forward hovering and tilting, time. During this period, the rotor tilts in the opposite direction, and the aircraft remains hovering throughout this dynamic process. The design was tested during a hovering flight maneuver after the rotor completed a 90° tilt, demonstrating good tracking response even in the "upright" position of the four tiltrotor aircraft, compared with a PID controller. The designed controller was compared with a classic PID controller under the same simulation conditions, and the control response results are as follows: Figures 8(a) to 8(c) , Figures 9(a) to 9(c) , Figures 10(a) to 10(c) , Figures 11(a) to 11(c) The figures shown are the position tracking curve, velocity response curve, attitude response curve, and angular velocity response curve, respectively. Simulation results show that, throughout the entire flight envelope, the controller designed in this invention can achieve better desired state tracking performance: its position and attitude tracking errors are smaller, and the response process is smoother. Figures 9(a) to 9(c) , Figures 10(a) to 10(c) The results further demonstrate that the proposed controller can achieve a continuous change in rotor tilt angle from 90° to 0° without causing significant speed fluctuations, with a stable transition process and no obvious overshoot or oscillation. Even when the tilt angle reaches 0°, the system can still effectively track the desired position and attitude commands, indicating that the controller has good adaptability and robustness to dynamic changes caused by configuration switching. In summary, the designed control strategy can achieve higher tracking accuracy and better transition smoothness throughout the entire flight simulation.

Claims

1. A method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering mode based on redundant variable-pitch control, characterized in that, include: S1. Obtain the current position, attitude, angular velocity, acceleration, and rotor tilt angle of the multi-tilt rotor aircraft; S2. A parameter-adaptive proportional-integral-derivative second-order position outer loop controller is used to calculate the position outer loop control quantity based on the desired position trajectory. S3. Solve for the desired attitude angle and the nominal term of the total force control channel quantity based on the outer loop control quantity of the position and the acceleration feedback, and perform adaptive compensation based on the acceleration feedback on the nominal term of the total force control channel quantity to obtain the total force control channel quantity. S4. A parameter adaptive proportional-integral first-order attitude and angular velocity dual-loop controller is used to estimate the desired angular velocity and angular velocity control quantity based on the desired attitude angle; S5. Redundant execution control allocation is performed on the angular velocity control quantity and total force control channel quantity based on the geometric relationship of the control components to obtain the actual actuation quantity of the rotor control actuator, including collective pitch control quantity, lateral cyclic pitch control quantity, and longitudinal cyclic pitch control quantity.

2. The method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 1, characterized in that, In S2, the parameter adaptive proportional-integral-derivative second-order position outer loop controller includes a roll channel outer loop controller, a pitch channel outer loop controller, and a yaw channel outer loop controller. The control law of the roll channel outer loop controller is: , ,in, This is the outer loop control value for the position of the roll channel. For the parameters of the roll controller, The desired position of the roll-off channel. This represents the actual position of the rolling channel.

3. The method for controlling the tilt-rotor flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 2, characterized in that, In step S3, the desired attitude angle and the nominal term of the total force control channel quantity are solved based on the outer loop control quantity of the position and the acceleration feedback. Specifically: The desired resultant force in the ground coordinate system is calculated based on the outer loop control quantity of the position and the acceleration feedback compensation term. , , ,in, To represent the desired resultant force in the ground coordinate system, For the mass of the aircraft, This is the outer loop control value for the pitch channel position. This refers to the outer ring control quantity for the yaw channel position. It is the acceleration due to gravity. This is the rotation matrix for transforming from the body coordinate system to the ground coordinate system. The rotation transformation matrix required to achieve the outer loop position control is as follows: To represent the desired resultant force in the body coordinate system, The nominal value for the total force control channel quantity. For the acceleration of the pitch and yaw channels, This refers to the rotor tilt angle; Based on the desired heading angle Calculate the rotational transformation matrix required to implement the outer loop control quantity. , , ; The desired roll angle of the aircraft is solved in the body coordinate system based on the desired resultant force direction. and desired pitch angle , , .

4. The method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 3, characterized in that, In step S3, adaptive compensation based on acceleration feedback is performed on the nominal term of the total force control channel quantity, specifically as follows: ,in, This is an adaptive compensation term for the total force control channel quantity. This is the gain coefficient for compensation in the direction of total force. The acceleration of the rolling channel; the total force obtained controls the channel volume. for .

5. The method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 4, characterized in that, In S4, the parameter adaptive proportional-integral first-order attitude and angular velocity dual-loop controller includes a roll channel dual-loop controller, a pitch channel dual-loop controller, and a yaw channel dual-loop controller. The roll channel dual-loop controller includes a roll channel attitude angle outer loop and a roll channel angular velocity inner loop. The control law for the roll channel attitude angle outer loop is as follows: The inner loop control law for the roll channel angular velocity is: ,in, For the desired roll rate, This is the actual roll angular velocity. For the desired roll angle, This is the actual roll angle. These are the outer loop control parameters for the roll channel attitude angle. These are the inner loop control parameters for the angular velocity of the roll channel.

6. The method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 5, characterized in that, In step S5, redundant execution control allocation is performed on the angular velocity control quantity and the total force control channel quantity based on the geometric relationship of the control components. Specifically, the control channel quantity output by the flight controller is allocated as a linkage control quantity according to the rotor tilt angle, and the linkage control quantity is allocated to the actual actuation quantity of the rotor control actuator.

7. The method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 6, characterized in that, Based on the rotor tilt angle, the control channel quantity output by the flight controller is allocated as a linkage control quantity, specifically: ,in, This is the total distance linkage. This is the total distance difference between the front and rear ends. The total distance difference between the left and right sides is the momentum. This refers to the longitudinal periodic variable pitch linkage. This refers to the longitudinal periodic variable pitch differential momentum. The left and right longitudinal periodic variable pitch differential momentum. This refers to the lateral periodic variable pitch linkage. This refers to the forward and backward lateral periodic pitch differential momentum. It is the left and right lateral periodic variable pitch differential momentum. To control the distribution gain parameters, To control the amount of the rollover channel Distributed to left and right total pitch differential Dynamic gain, , To control the amount of the rollover channel Distributed to left and right longitudinal periodic variable pitch differential Dynamic gain, , To adjust the pitch control channel quantity Distributed to front and rear total pitch differential Dynamic gain, , To adjust the pitch control channel quantity Distributed to longitudinal periodic variable pitch linkage Dynamic gain, , To adjust the pitch control channel quantity Distributed to longitudinal periodic differential momentum Dynamic gain, .

8. The method for controlling the tilting flight of a multi-tilt rotorcraft in low-speed hovering state based on redundant variable-pitch control according to claim 7, characterized in that, The linkage control quantity is distributed to the actual actuator momentum of each rotor, specifically as follows: ,in, This refers to the collective pitch control of rotors 1 through 4. This refers to the lateral cyclic pitch control values ​​for rotors 1 through 4. This refers to the longitudinal cyclic pitch control parameters for rotors 1 through 4.

9. A flight controller comprising a memory and a processor, wherein the memory stores a computer program that runs on the processor, characterized in that, When the processor runs a computer program, it executes the steps of the control method of claim 1.

10. A multi-tilt rotor aircraft, characterized in that, The system includes: an airframe and at least four sets of tilt-rotor aircraft mounted on the airframe, rotor control actuators, flight status sensors, and the flight controller as described in claim 9; the flight status sensors are used to acquire the position, attitude, angular velocity, acceleration, and rotor tilt angle information of the multi-tilt rotor aircraft; the flight controller as described in claim 9 is used to allocate the actual actuation force of the rotor control actuators according to the real-time position, attitude, angular velocity, acceleration, and rotor tilt angle of the multi-tilt rotor aircraft; the rotor control actuators are used to drive each set of tilt-rotor aircraft to perform collective pitch control, longitudinal cyclic pitch control, lateral cyclic pitch control, and tilt control according to the actual actuation force.

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