A control method of a tail seat type unmanned aerial vehicle

Abstract

This application discloses a control method for a tail-mounted unmanned aerial vehicle (UAV). The method acquires the current pitch angle of the UAV in real time. When the absolute value of the pitch angle exceeds a preset threshold, it enters a large pitch angle control mode. In this mode, on the one hand, coordinated turn control is executed: the lateral translation control channel is closed, while the roll control channel remains open. The desired roll angle is acquired, and a coordinated yaw rate command is calculated based on the desired roll angle, desired pitch angle, and current airspeed to enable the UAV to achieve a coordinated turn flight state and reduce sideslip. On the other hand, coupled pitch altitude control is executed: based on the altitude control input and the current pitch angle, the X-axis vector of the UAV is projected onto a horizontal plane and rotated around the Y-axis by a pitch compensation angle to construct a virtual horizontal plane. The attitude control loop makes the UAV's attitude follow the changes in the virtual horizontal plane, indirectly achieving altitude adjustment by changing the thrust direction, thus compensating for the impact of thrust direction changes on vertical control capability at large pitch angles.

Description

Technical Field

[0001] This application relates to the field of unmanned aerial vehicle (UAV) flight control technology, and more particularly to a control method for a tail-mounted UAV. Background Technology

[0002] Tail-seat unmanned aerial vehicles (UAVs) are a type of fixed-wing aircraft capable of both vertical takeoff and landing (VTOL) and near-horizontal flight. In practical engineering applications, especially on miniaturized and lightweight tail-seat UAV platforms, to reduce system complexity and improve reliability, flight control systems often run on embedded processors or microcontroller platforms with limited computing power and resources. Against this backdrop, quadcopter-based control frameworks are commonly used in engineering implementations to achieve attitude and altitude control for tail-seat UAVs. This control framework typically uses an inner and outer loop structure as its core. The inner loop controls roll, pitch, and yaw angles, while the outer loop adjusts altitude through total thrust or vertical velocity. This approach offers advantages such as simplicity, low computational complexity, and high engineering maturity.

[0003] In research on tail-seat unmanned aerial vehicles (UAVs) and high angle-of-attack flight control, scholars have attempted to address the control difficulties arising from the strong coupling between attitude, thrust, and aerodynamics using various methods. One type of method is based on a complete dynamic and aerodynamic model, achieving full-envelope flight control through optimized control, but this typically requires high model accuracy and online computing resources. Another type of method is based on incremental nonlinear dynamic inversion (INDI), achieving robust control through angular or linear acceleration information measured by sensors, but this is somewhat dependent on sensor accuracy and calibration quality. Both of these methods have relatively complex control structures, limiting their engineering applicability on low-computing-power flight control platforms.

[0004] Therefore, in practical engineering applications, a quadcopter-based control framework is still commonly used to control tail-seat UAVs. This approach implicitly assumes that by adjusting the thrust components caused by yaw and pitch angles, desired control forces can be generated in the lateral and vertical directions, respectively, thus achieving approximate decoupling between lateral motion control and altitude control. This assumption is reasonable in engineering when the tail-seat UAV is flying at a small pitch angle. However, when the UAV enters a large pitch angle flight state, the thrust vector direction rotates significantly, and its effective component in the direction of gravity decreases significantly, resulting in a strong coupling between attitude control and altitude control. Under these conditions, if traditional control strategies are still used, problems such as increased sideslip and discontinuous flight paths are likely to occur during turns, while the efficiency and stability of altitude control decrease significantly.

[0005] In summary, there is a lack of existing technologies that can effectively address the issues of insufficient turn coordination and altitude control performance in tail-mounted UAVs during high pitch angle flight, while maintaining the quadcopter control framework unchanged. Summary of the Invention

[0006] This application provides a control method for a tail-seat unmanned aerial vehicle (UAV) to solve the technical problem in the prior art where the turn control and altitude control based on a quadcopter control frame are strongly coupled and the control performance is reduced when the tail-seat UAV is flying at a large pitch angle.

[0007] In a first aspect, this application provides a control method for a tail-seat unmanned aerial vehicle (UAV), applied to a tail-seat UAV equipped with a flight control system, wherein the flight control system operates based on a quadcopter control frame, and the method includes: acquiring the current pitch angle of the tail-seat UAV; Determine whether the absolute value of the pitch angle exceeds a preset threshold; In response to the absolute value of the pitch angle exceeding the preset threshold, the system enters a large pitch angle control mode; In the large pitch angle control mode, coordinated turn control is performed. The coordinated turn control includes: closing the lateral translation control channel, keeping the roll control channel open, obtaining the desired roll angle and current airspeed, and obtaining a coordinated yaw rate command based on the desired roll angle, the desired pitch angle currently output by the flight control system, and the current airspeed, and controlling the tail-mounted UAV to perform coordinated turn flight based on the desired roll angle and the coordinated yaw rate command. In the large pitch angle control mode, coupled pitch altitude control is performed. The coupled pitch altitude control includes: based on the altitude control input and the current pitch angle, constructing a virtual horizontal plane by projecting the x-axis direction vector of the aircraft onto the horizontal plane and rotating it around the y-axis of the aircraft by a pitch compensation angle determined by the altitude control input; using the pitch attitude of the virtual horizontal plane as the pitch control target; and using the attitude control loop to make the attitude of the tail-mounted UAV follow the changes of the virtual horizontal plane to achieve altitude adjustment.

[0008] Preferably, obtaining the coordinated yaw rate command based on the desired roll angle, the desired pitch angle currently output by the flight control system, and the current airspeed includes: The coordinated yaw rate command is obtained according to the following formula: in, For the coordinated yaw rate command, It is the acceleration due to gravity. The airspeed of the tail-mounted UAV is given. The desired roll angle, The desired pitch angle; The desired pitch angle is the pitch angle setting currently output by the flight control system.

[0009] Preferably, closing the lateral translation control channel includes: Set the lateral acceleration control command and lateral velocity control command in the body coordinate system to zero, and pause the closed-loop control of lateral translation.

[0010] Preferably, the construction of the virtual horizontal plane includes: Obtain the direction vector of the tail-mounted UAV's x-axis in the inertial coordinate system. ; Projecting the direction vector onto the horizontal plane yields the horizontal projection vector. ,in: in, is the unit vector perpendicular to the inertial coordinate system; The surrounding body Axis rotation includes: rotating the horizontal projection vector Surrounding the body The pitch compensation angle is rotated along the axis to obtain a longitudinal reference direction vector for the virtual horizontal plane; the longitudinal reference direction vector is then compared with the body... The cross product along the axis is used to obtain the normal vector of the virtual horizontal plane; wherein, the body axis Let be the horizontal axis of the tail-mounted UAV's body coordinate system; wherein, the pitch compensation angle is obtained by proportional mapping from the altitude control input.

[0011] Preferably, the pitch compensation angle is obtained in the following way: Multiply the altitude control input by the pitch limit value to obtain the pitch compensation angle; The value range of the height control input is as follows: The pitch limit value is the upper limit of the preset virtual horizontal plane pitch angle range.

[0012] Preferably, the pitch limit value is 30 degrees.

[0013] Preferably, the method further includes: In response to the absolute value of the pitch angle falling back to the preset threshold range, the large pitch angle control mode is exited; The pitch control target generated based on the virtual horizontal plane is subjected to low-pass filtering attenuation processing; Smoothly transition height control to conventional height control methods.

[0014] Preferably, the step of making the attitude of the tail-mounted UAV follow the changes of the virtual horizontal plane through the attitude control loop includes: the flight control system generates throttle output through a multi-loop PID controller to control the position, speed and acceleration state of the tail-mounted UAV on the virtual horizontal plane; The state error is projected onto the vertical direction of the virtual horizontal plane.

[0015] Preferably, the coordinated turn control is designed under the assumption that the sideslip angle β is zero; under this assumption, the fuselage of the tail-mounted UAV is symmetrical in the XZ plane, and the lateral aerodynamic coefficient is... ,in For the angle of attack, the coordinated turn control maintains the sideslip angle close to zero by coordinating and matching the roll angle with the yaw rate.

[0016] Preferably, the preset threshold value ranges from 70° to 90°; The process of obtaining the current pitch angle of the tail-mounted UAV includes: obtaining the pitch angle through attitude calculation based on the angular velocity data and acceleration data obtained by the inertial measurement unit; The method further includes: acquiring altitude information based on a barometric altimeter and a global navigation satellite system as altitude feedback for the coupled pitch altitude control; and acquiring ground speed and heading information based on the global navigation satellite system as track reference for the coordinated turn control.

[0017] The technical solution provided in this application enables the flight control system to automatically enter a high pitch angle control mode when it detects that the pitch angle exceeds a preset threshold. On one hand, by closing the lateral translation control channel and obtaining a coordinated yaw rate command based on the desired roll angle, desired pitch angle, and current airspeed, the tail-seat UAV achieves a coordinated turning flight state under high pitch angle conditions, thereby reducing the sideslip angle during turns and improving flight path consistency. On the other hand, by constructing a virtual horizontal plane, the altitude control input is converted into pitch attitude changes, indirectly achieving altitude adjustment by changing the thrust direction. This solution effectively improves the altitude response lag problem caused by the thrust direction deviating from the gravity direction under high pitch angle flight conditions without introducing flight mode switching or relying on complex dynamic models. It improves the response speed and stability of altitude control, and solves the technical problems of insufficient turning coordination and altitude control performance of tail-seat UAVs under high pitch angle flight conditions. Attached Figure Description

[0018] Figure 1 This is a schematic diagram of the coordinate system of a tail-mounted unmanned aerial vehicle provided in an embodiment of this application; Figure 2 A flowchart illustrating the control method for a tail-mounted unmanned aerial vehicle provided in an embodiment of this application; Figure 3 This is a schematic diagram of height control provided for an embodiment of this application. Detailed Implementation

[0019] The technical solutions in the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings.

[0020] The tail-seat UAV control system provided in this application includes a tail-seat UAV platform and a ground control terminal. The tail-seat UAV platform adopts a tail-seat vertical takeoff and landing (VTOL) layout. During takeoff and landing, the aircraft operates in a near-vertical attitude. During flight, it can maintain a continuous flight attitude within a wide pitch angle range, possessing both hovering and near-horizontal flight capabilities, covering the complete attitude change process from low to high pitch angles. The tail-seat UAV platform is equipped with a flight control system, which operates based on a quadcopter control framework. This framework is centered on an inner and outer loop structure. The inner loop controls roll, pitch, and yaw angles, while the outer loop adjusts altitude through total thrust or vertical velocity. In a preferred embodiment, the flight control system is implemented based on the PX4 open-source flight control architecture. This architecture employs a multi-level cascaded control structure consisting of a position loop, velocity loop, acceleration loop, attitude loop, and angular rate loop: the outer loop position controller outputs the desired velocity based on the position error; the velocity controller outputs the desired acceleration based on the velocity error and converts it into desired tilt angle and thrust commands; the inner loop attitude controller converts the desired attitude into angular rate commands; and the angular rate controller drives the actuators. The aforementioned attitude controller and angular rate controller together constitute an attitude control loop. This loop receives the desired roll angle, desired pitch angle, and desired yaw rate as inputs, and generates actuator control commands through cascaded PID control to achieve attitude following control of the aircraft. In conventional control mode, lateral translation control is achieved by calculating the desired roll angle through the outer loop position controller and velocity controller. Those skilled in the art will understand that the above control architecture is not limited to the PX4 system; any quadcopter control framework employing a similar multi-level cascaded PID structure is applicable to the methods described in this application. The tail-seat UAV platform features a lightweight airframe with good overall rigidity. Its power system uses a multi-rotor propeller layout, and its thrust direction is strongly correlated with the airframe attitude, enabling it to accurately reflect thrust component changes and attitude-motion coupling characteristics during high pitch angle flight. The tail-seat UAV platform also carries onboard sensor components, including an inertial measurement unit (IMU), a Global Navigation Satellite System (GNSS) module, a barometric altimeter, and a magnetometer. The ground control terminal is a handheld remote controller, which establishes a two-way data communication connection with the tail-seat UAV platform via a wireless communication link. The remote controller includes a pitch / forward joystick and an altitude joystick. The pitch / forward joystick is used to send pitch control commands and desired roll angles to the tail-mounted UAV platform, while the altitude joystick is used to send altitude control inputs. The ground control unit simultaneously receives flight status data transmitted back from the tail-mounted UAV platform.

[0021] like Figure 1 As shown, the tail-mounted UAV described in this embodiment adopts a body coordinate system, where the xb axis points forward of the nose, the yb axis points to the right wing, and the zb axis is determined by the right-hand rule and points to the belly.

[0022] It should be noted that the inertial coordinate system described in this application can adopt the North-East-Earth (NED) convention, where the x-axis points due north, the y-axis points due east, and the z-axis points towards the Earth's center. In the body coordinate system, the xb-axis strictly points forward of the nose, the yb-axis strictly points towards the right wing, and the zb-axis is determined by the right-hand rule, pointing towards the belly of the aircraft. The attitude angles described in this application adopt the ZYX rotation sequence (i.e., first rotate the yaw angle around the z-axis of the inertial frame). Then rotate the pitch angle θ around the rotated y-axis, and finally rotate the roll angle around the rotated x-axis. The pitch angle θ is defined as follows: [-90°, 90°]. It should be noted that the attitude calculation within the flight control system uses quaternions to avoid the gimbal lock singularity problem of Euler angles when the pitch angle is close to ±90°. In this application, the pitch angle θ is used as a scalar criterion, obtained by extracting from the quaternion using inverse trigonometric functions. It is only used for threshold judgment and coordinating yaw rate formula calculation, and is not used as an internal state representation of the attitude control loop. It should be noted that in the engineering implementation, the flight control system rotates and adapts the body coordinate system so that the pitch angle corresponding to the hovering state is zero or close to zero. Therefore, the pitch angle in this application gradually increases as the aircraft transitions from hovering to level flight. When the absolute value of the pitch angle is large, the aircraft is in a transitional flight state from hovering to level flight, the thrust direction approaches horizontal, and its effective component in the direction of gravity decreases accordingly. The term "large pitch angle" in this application refers to the case where the absolute value of the pitch angle is large under this convention. The term "hovering" as used in this application refers to the flight state in which a tail-seat UAV maintains its position in the air at a near-vertical attitude, with its nose pointing upwards, its x-axis approximately pointing towards the zenith, and its thrust direction approximately opposite to the direction of gravity. The term "level flight" as used in this application refers to the flight state in which the tail-seat UAV transitions from a near-vertical hovering attitude to a flight state where its x-axis is nearly horizontal. In this state, the aircraft's thrust direction is approximately horizontal, and the aircraft primarily relies on aerodynamic lift to maintain altitude; its flight characteristics are similar to those of a conventional fixed-wing aircraft.

[0023] The aerodynamic forces acting on an aircraft during flight include lift L (perpendicular to the airspeed vector Va), drag D (opposite to the airspeed vector Va), and side force Y (along the yb axis). The airspeed vector Va is the velocity vector of the aircraft relative to the air. The angle of attack α is the angle between the projection of the airspeed vector onto the xb-zb plane of the aircraft and the xb axis of the aircraft, representing the longitudinal deviation of the airflow relative to the longitudinal axis of the aircraft. The sideslip angle β is the angle between the airspeed vector and the xb-zb plane of the aircraft, representing the lateral deviation of the airflow relative to the plane of symmetry of the aircraft.

[0024] like Figure 2 As shown, the control method for a tail-mounted unmanned aerial vehicle provided in this application includes the following steps: S201, obtain the current pitch angle of the tail-mounted UAV.

[0025] In this step, the flight control system acquires the aircraft's attitude information in real time through onboard sensor components. Specifically, the inertial measurement unit (IMU) includes a three-axis accelerometer and a three-axis gyroscope, used to measure the linear acceleration and angular velocity of the aircraft in real time, providing basic data support for attitude estimation. Based on the angular velocity and acceleration data acquired by the IMU, the flight control system obtains the current pitch angle of the tail-mounted UAV through attitude calculation. Preferably, the method further includes: acquiring relative altitude information based on a barometric altimeter and altitude information based on a global navigation satellite system (GNSS), as altitude feedback for subsequent coupled pitch altitude control; that is, acquiring barometric altimeter data and combining it with altitude data provided by the GNSS, and obtaining altitude information after data fusion processing, as altitude feedback for subsequent coupled pitch altitude control. Among these technologies, the barometric altimeter offers the advantage of sensitive response to short-term altitude changes, while the Global Navigation Satellite System (GNSS) provides an absolute altitude reference. The fusion of these two technologies ensures both real-time accuracy and long-term precision in altitude measurement. Ground speed and heading information acquired through GNSS can be used to analyze track changes and speed characteristics during turns, serving as a track reference for coordinated turn control. The magnetometer acquires the aircraft's heading angle information, assisting in attitude calculation and heading analysis. All of these sensor data are synchronized and fused by the flight control system to form complete flight status information.

[0026] It should be noted that the tail-seat UAV platform described in this embodiment is not equipped with an airspeed indicator. The airspeed information required in the coordinated yaw rate formula is approximately obtained by measuring the ground speed using a global navigation satellite system under typical flight conditions. Since the tail-seat UAV described in this application mainly performs flight missions in low to medium wind speed environments, the difference between ground speed and airspeed is within an acceptable range in engineering. Under conditions of higher wind speeds, the ground speed can be corrected by introducing wind speed estimation to obtain a more accurate airspeed estimate. Furthermore, when the airspeed is lower than a preset minimum airspeed threshold, the coordinated yaw rate command is set to zero or maintains the output value of the previous effective control cycle to avoid division by zero anomalies. In a preferred embodiment, the preset minimum airspeed threshold is 3 m / s.

[0027] S202, determine whether the absolute value of the pitch angle exceeds the preset threshold.

[0028] In this step, the flight control system compares the absolute value of the currently acquired pitch angle with a preset threshold. When the absolute value of the pitch angle exceeds the preset threshold, the aircraft is determined to be in a high pitch angle flight state, and step S203 is executed to enter the high pitch angle control mode. When the absolute value of the pitch angle does not exceed the preset threshold, the flight control system maintains the conventional quadcopter control mode, that is, lateral motion control and altitude control are achieved through yaw channel and thrust adjustment, respectively. Furthermore, the preset threshold ranges from 70° to 90°. When the pitch angle of the tail-seat UAV is within this range, there is a significant deflection between the thrust vector direction and the gravity direction, and the performance of the traditional quadcopter-based control method in turning and altitude control is significantly reduced, requiring entry into the high pitch angle control mode to improve the control effect. This threshold triggering mechanism does not require explicit flight mode switching or state machine management, ensuring the continuity of control and system stability.

[0029] S203, in response to the absolute value of the pitch angle exceeding the preset threshold, enters the large pitch angle control mode.

[0030] In this step, after detecting that the pitch angle exceeds a preset threshold, the flight control system automatically switches from the conventional quadcopter control mode to the high pitch angle control mode. This switching process is completed within a unified quadcopter control framework, without the need for independent flight mode switching logic. In the high pitch angle control mode, the flight control system simultaneously performs coordinated turn control and coupled pitch altitude control. It should be noted that the necessity of entering the high pitch angle control mode is as follows: when the aircraft transitions from hovering to level flight, traditional control methods still use the control logic under hovering or small attitude angle conditions, achieving turning through yaw (around the fuselage's zb axis). Yaw significantly changes the thrust direction, which does not conform to the aerodynamic characteristics of the aircraft in this flight state. During the turn, problems such as increased sideslip, discontinuous flight path curvature, or unpredictable turning radius are likely to occur; at the same time, the effective component of thrust in the direction of gravity is significantly reduced, and altitude control efficiency is significantly decreased.

[0031] It should be noted that steps S204 and S205 are executed in parallel by the flight control system in the large pitch angle control mode, and the order of the step numbers does not represent the order of execution time.

[0032] S204, in the large pitch angle control mode, performs coordinated turning control.

[0033] In this step, the flight control system performs coordinated turn control, which specifically includes the following sub-steps: S2041, Close the lateral translation control channel.

[0034] The flight control system suspends closed-loop control of lateral acceleration or lateral velocity in the body coordinate system to avoid unreasonable lateral thrust distribution under high pitch angle conditions. Specifically, closing the lateral translation control channel means disconnecting the output paths of the outer loop position controller and velocity controller for roll angle commands, so that the roll angle is no longer driven by the outer loop lateral position or velocity error. In traditional quadcopter control, the lateral translation control channel calculates the desired roll angle through the outer loop controller to generate lateral control force, achieving lateral motion control; while turning mainly relies on the yaw channel to change the heading. However, under high pitch angle conditions, on the one hand, directly generating lateral control force in this way not only introduces a significant lateral motion component but also affects the vertical thrust distribution; on the other hand, yaw turns will significantly change the thrust direction, which does not conform to the aerodynamic characteristics of the aircraft at high pitch angles and is prone to sideslip. Therefore, it is necessary to close this channel and change the turning strategy from yaw-dominated to a method that coordinates roll and tilt with yaw rate. After closing the lateral translation control channel, the desired roll angle is directly input from the outside, including remote control joystick commands or the output of the automatic navigation system.

[0035] S2042, keep the roll control channel open and obtain the desired roll angle and current airspeed.

[0036] The flight control system retains the roll control channel open. The roll control channel refers to the inner loop control path within the attitude control loop that receives the desired roll angle and drives the actuators via the angular rate controller to achieve roll attitude following. After closing the lateral translation control channel, the desired roll angle of the roll control channel is no longer generated by the outer loop lateral position or velocity error, but is given by the externally input roll control command. Simultaneously, the flight control system acquires the current airspeed, as described in step S201. The desired roll angle generates aerodynamic centripetal force, thereby guiding the flight trajectory to change direction. Under large pitch angle conditions, the UAV's heading change is mainly guided by the desired roll angle rather than directly driven by the yaw channel, which is consistent with the control mechanism of fixed-wing aircraft introducing centripetal acceleration through body tilting under large angles of attack conditions.

[0037] S2043, obtain a coordinated yaw rate command based on the desired roll angle, desired pitch angle and current airspeed.

[0038] The flight control system calculates the coordinated yaw rate command based on the current desired roll angle, desired pitch angle, and current airspeed, thus obtaining a matching coordinated yaw rate command. Preferably, the formula for calculating the coordinated yaw rate command is: in, For the coordinated yaw rate command, It is the acceleration due to gravity. The airspeed of the tail-mounted UAV is given. The desired roll angle, The desired pitch angle is the pitch angle setpoint currently output by the flight control system. In the large pitch angle control mode, the coordinated turn control and coupled pitch altitude control are executed in parallel within each control cycle. In the calculation process of a single control cycle, the flight control system first determines the pitch control target on the virtual horizontal plane based on the altitude control input and the current pitch angle, and uses it as the desired pitch angle for the current cycle. Subsequently, the coordinated yaw rate formula is used. Calculate the yaw rate command. Therefore, Within each control cycle, the virtual horizontal plane is determined first, and the coordinated yaw rate formula is used later; there is no cyclic dependency between the two.

[0039] Through the above relationships, the yaw rate and the desired roll angle can be reasonably matched during turning, maintaining the flight speed direction basically consistent with the longitudinal axis of the aircraft, suppressing the tendency to sideslip, and enabling the UAV to achieve a coordinated turning flight state. It should be noted that when the desired pitch angle... When approaching 90°, As the yaw rate approaches zero, the coordinated yaw rate command decreases accordingly. Physically, this behavior corresponds to the aircraft approaching level flight, where it primarily relies on aerodynamic lift generated by the wings to maintain flight, and turns are mainly achieved through the aerodynamic centripetal force introduced by roll. The required coordinated yaw rate is inherently small, therefore the natural decay of the formula is consistent with the physical characteristics of flight. In engineering implementation, this can be achieved by... Apply a lower limit clamp to the item (e.g.) ≥ 0.1) to ensure a certain coordinated yaw rate output even at maximum pitch angles. Furthermore, in hovering or low-speed flight (small pitch angle), the aircraft's airspeed may fall below the preset minimum airspeed threshold; in this case, the following airspeed protection mechanism is used. The coordinated turn refers to the UAV, during a turn, appropriately matching the roll angle and yaw rate to ensure that the flight speed direction is basically consistent with the longitudinal axis of the aircraft, thereby avoiding significant sideslip. It should be noted that the calculation of the coordinated yaw rate command requires a minimum airspeed threshold. The command is executed when the airspeed exceeds a preset minimum airspeed threshold (e.g., 3 m / s); when the airspeed is below the minimum airspeed threshold, the coordinated yaw rate command is either set to zero or maintains the output value of the previous effective control cycle to avoid division by zero anomalies. Simultaneously, the desired roll angle... The maximum roll angle is limited (e.g., 45°) to ensure that the calculation results are bounded.

[0040] S2044, based on the desired roll angle and the coordinated yaw rate command, control the tail-mounted UAV to perform coordinated turning flight.

[0041] The flight control system outputs the desired roll angle and coordinated yaw rate commands as roll control commands and yaw control commands to the attitude control loop, which then drives the actuators to complete the coordinated turn. Through this control method, the UAV's control strategy at high pitch angles changes from the traditional yaw-dominated turn (around the ZB axis) to a coordinated turn control that combines roll (around the XB axis) with coordinated yaw rate, exhibiting flight behavior similar to that of a fixed-wing aircraft at high angles of attack. This coordinated turn control effectively reduces sideslip during turns, improves flight path consistency, and enhances the predictability of the turn trajectory.

[0042] Furthermore, the coordinated turning control is designed under the assumption that the sideslip angle β is zero. This coordinated turning control maintains the sideslip angle close to zero by coordinating and matching the roll angle with the yaw rate. Specifically, as... Figure 1 As shown, in the dynamic model of the tail-mounted UAV, the aerodynamic forces in the body coordinate system From the angle of attack and sideslip angle The lift coefficient is determined. drag coefficient and lateral force coefficient Both are functions of angle of attack and sideslip angle. Assuming the aircraft is symmetrical in the XZ plane, then... , , .

[0043] Under the condition that the sideslip angle is zero, the lateral aerodynamic coefficient This means that lateral aerodynamic forces can be approximately ignored. This assumption simplifies the design of coordinated turn control, allowing the control system to focus only on the coordinated matching of the desired roll angle and yaw rate without having to deal with the effects of lateral aerodynamic forces.

[0044] S205, in large pitch angle control mode, performs coupled pitch altitude control.

[0045] In this step, the flight control system performs coupled pitch altitude control, transforming altitude control from a direct adjustment of total thrust to a virtual horizontal plane control method based on pitch attitude. It should be noted that in existing flight control systems based on quadrotor control frameworks, altitude control and attitude control are not perfectly symmetrical in engineering implementation: during upward altitude control, the system allows adjustment of pitch attitude to assist in generating the desired vertical thrust component; while during downward altitude control, the focus is usually on reducing total thrust, and the introduction of pitch-down attitude changes is deliberately avoided. This asymmetric strategy is engineeringly reasonable under small pitch angle conditions, but it easily leads to altitude response lag under large pitch angle flight conditions. This application effectively improves this problem by introducing a coupled pitch altitude control method. In a specific implementation, the following sub-steps are included: S2051, Based on the altitude control input and the current pitch angle, a virtual horizontal plane is constructed by projecting the x-axis direction vector of the aircraft onto the horizontal plane and rotating it around the y-axis of the aircraft by a pitch compensation angle determined by the altitude control input.

[0046] Figure 3 The upper part of the diagram illustrates the initial relationship between the UAV's X-axis and the virtual horizontal plane when the UAV is flying at a large pitch angle. The lower part shows the effect of the virtual horizontal plane tilting relative to its initial horizontal position after the operator inputs altitude control commands via the altitude joystick. This virtual horizontal plane is not a horizontal plane in the actual geographical sense, but rather a reference attitude plane used to characterize the aircraft's altitude control target. The pitch attitude of the virtual horizontal plane is determined by the current aircraft pitch angle and a pitch compensation amount, which is obtained by proportionally mapping the altitude control input. When the altitude control input is upward, the virtual horizontal plane pitch angle increases; when the altitude control input is downward, the virtual horizontal plane pitch angle decreases.

[0047] Specifically, the process of constructing the virtual horizontal plane includes: firstly, obtaining the direction vector of the tail-mounted UAV's x-axis in the inertial coordinate system. ,like Figure 1 As shown, this vector corresponds to the body coordinate system. The axis is represented in an inertial coordinate system; then the direction vector is projected onto the horizontal plane to obtain the horizontal projection vector. Its calculation formula is ,in This is the unit vector in the vertical direction of the inertial coordinate system. It should be noted that when entering the large pitch angle control mode, the introduction of roll will not change the direction of the aircraft's x-axis projection onto the horizontal plane. This represents the projection of the aircraft's x-axis onto the horizontal plane. Finally, the horizontal projection vector... Rotation around the y-axis of the aircraft by a pitch compensation angle Obtain the longitudinal reference direction vector of the virtual horizontal plane. Let the unit vector of the body's y-axis in the inertial coordinate system be... ,because Perpendicular to The formula for calculating the rotation is: .Will and The cross product is used to obtain the normal vector of the virtual horizontal plane. The virtual horizontal plane is composed of and The common span, the normal vector n, is used to determine the attitude direction of the virtual horizontal plane. For example... Figure 3 As shown, the rotated virtual horizontal plane is tilted relative to its initial horizontal position, and the direction and angle of this tilt are determined by the height control input. The pitch compensation angle is obtained by proportional mapping from the height control input.

[0048] Furthermore, the pitch compensation angle is obtained by multiplying the altitude control input by the pitch limit value. The altitude control input has a value range of [range missing]. The pitch limit value corresponds to the travel range of the altitude control stick from its lowest to its highest position on the ground control terminal, and is the upper limit of the preset virtual horizontal plane pitch angle range. Preferably, the pitch limit value is 30 degrees, meaning the virtual horizontal plane range corresponding to the full altitude stick control is ±30 degrees. This limiting design ensures system stability and avoids excessive pitch compensation leading to abnormal flight attitude.

[0049] S2052, the pitch attitude of the virtual horizontal plane is used as the pitch control target. In the coupled pitch altitude control mode, altitude control no longer directly generates vertical thrust correction, but is indirectly achieved through the virtual horizontal plane. Specifically, the flight control system extracts the corresponding pitch angle value based on the normal vector of the virtual horizontal plane, and uses this pitch angle value as the desired pitch angle of the pitch channel in the attitude control loop. This replaces the pitch angle setpoint directly given by the flight mission or control commands in the conventional mode. The flight control system uses the pitch attitude of the virtual horizontal plane as the pitch control target of the aircraft.

[0050] S2053, the attitude control loop enables the tail-mounted UAV to adjust its altitude by following changes in the virtual horizontal plane. The flight control system uses inner-loop attitude control to make the aircraft's attitude follow changes in the virtual horizontal plane. Figure 3As shown in the lower half, when the altitude control input changes the tilt angle of the virtual horizontal plane, the aircraft attitude follows the change in the virtual horizontal plane, thereby achieving altitude adjustment by changing the thrust direction. Specifically, in high pitch angle flight, the aircraft's thrust direction is along the x-axis of the fuselage, with a large angle between it and the direction of gravity. When the altitude control input tilts the virtual horizontal plane upward, the pitch control target drives the aircraft's nose to pitch up, the thrust direction deflects in the opposite direction of gravity, the effective component of thrust in the direction of gravity increases, and the aircraft gains net upward acceleration, thus climbing. When the altitude control input tilts the virtual horizontal plane downward, the nose pitches down, the effective component of thrust in the direction of gravity decreases, and the aircraft descends. Simultaneously, the flight control system uses a multi-loop PID controller to perform closed-loop control of position, velocity, and acceleration in the normal direction of the virtual horizontal plane and generates throttle output, working in conjunction with the aforementioned pitch attitude adjustment to achieve precise altitude adjustment. This avoids the problem of decreased altitude control efficiency caused by changes in thrust direction under high pitch angle conditions.

[0051] Preferably, in the coupled pitch altitude control, the flight control system generates throttle output through a multi-loop PID controller to control the position, velocity, and acceleration states of the tail-seat UAV on the virtual horizontal plane, with state errors projected onto the vertical direction of the virtual horizontal plane. In conventional control methods, throttle output is used to control the position, velocity, and acceleration states of the aircraft on the horizontal plane. With the introduction of a virtual horizontal plane, the control logic is similar, but the reference plane is replaced by a virtual horizontal plane instead of the actual horizontal plane, and all state errors are calculated and projected onto the vertical direction of the virtual horizontal plane. It should be noted that "controlling the position, velocity, and acceleration states of the tail-seat UAV on the virtual horizontal plane" refers to closed-loop control of the aircraft's position, velocity, and acceleration in the altitude direction relative to the virtual horizontal plane, using the virtual horizontal plane as a reference. Specifically, the flight control system projects the position error, velocity error, and acceleration error of the aircraft in the inertial coordinate system onto the normal vector direction of the virtual horizontal plane, obtaining the components of each state in the vertical direction of the virtual horizontal plane, and uses these as the input to the multi-loop PID controller. The outer loop position controller outputs the desired normal velocity based on the normal position error, the velocity controller outputs the desired normal acceleration based on the normal velocity error, and the acceleration controller converts the desired normal acceleration into a throttle output command. In this way, the reference plane for altitude control is replaced by a virtual horizontal plane, allowing the control direction of the throttle control loop to adaptively follow changes in the virtual horizontal plane under large pitch angles, thus maintaining consistent altitude control response. This method provides better response consistency for altitude control under large pitch angles and effectively compensates for the impact of thrust direction changes on vertical control capability.

[0052] Furthermore, when the absolute value of the aircraft's pitch angle falls back to within the preset threshold range, the flight control system exits the high pitch angle control mode. During the exit process, the flight control system performs low-pass filtering attenuation processing on the pitch control target generated based on the virtual horizontal plane, smoothly transitioning altitude control to the conventional altitude control mode to avoid flight instability caused by sudden changes in control commands. This smooth transition mechanism ensures the continuity of control, enabling the aircraft to maintain flight stability during the switch between the high pitch angle control mode and the conventional control mode.

[0053] It should be further noted that the dynamic model of the tail-mounted UAV described in this application is based on the following kinematic formula (inertial coordinate system): in This indicates the position of the UAV in the inertial coordinate system. This represents the flight speed of the drone in the inertial coordinate system. The vector of gravitational acceleration. For thrust acceleration, Let be the rotation matrix from the body coordinate system to the inertial coordinate system. For the mass of the aircraft, This refers to the aerodynamic forces in the body coordinate system. For example... Figure 1 As shown, the aerodynamic forces in the body coordinate system are determined by the angle of attack. and sideslip angle The decision, among which, Angle of Attack , Sideslip angle The modulus of airspeed (denoted as in the coordinated yaw rate formula of this application) is the airspeed modulus. )), Let W be the airspeed vector and W be the wind speed. Transform the airspeed vector from the inertial coordinate system to the body coordinate system to obtain the airspeed vector in the body coordinate system. in This is the rotation matrix from the body coordinate system to the inertial coordinate system.

[0054] Lift , resistance , lateral force , in air density, For wing area, , ... These are the lift coefficient, drag coefficient, and lateral force coefficient, respectively, which are generally obtained through wind tunnel testing.

[0055] The above-mentioned lift ,resistance and lateral force Defined in the aerodynamic coordinate system, it needs to be transformed into the aerodynamic force in the body coordinate system through coordinate transformation. Specifically, the transformation relationship from the aerodynamic coordinate system to the body coordinate system is as follows: in, For the angle of attack, , , These are the body coordinate systems. , , The aerodynamic components along the axial direction. This transformation matrix indicates the longitudinal aerodynamic components under the machine system. and normal aerodynamics By lift and resistance Generated through rotational coupling at the angle of attack, while lateral aerodynamic forces... Directly equal to lateral force Under coordinated flight conditions (sideslip angle) ), due to the aforementioned symmetry conditions It can be known Therefore, The lateral aerodynamic forces under the engine system are negligible, which is the aerodynamic basis of the coordinated turning control method of this application.

[0056] The aforementioned dynamic model is used to analyze the rationality of the coordinated flight assumption and the derivation basis of the coordinated turn control formula. However, the control method described in this application does not rely on the precise identification of the above model parameters in the actual flight control process and is suitable for deployment on a low-computing-power flight control platform with a microcontroller as the core.

[0057] In summary, the control method for tail-seat UAVs provided in this application achieves coordinated turn control and coupled pitch altitude control within a unified quadcopter control framework through a control mode switching mechanism triggered by a pitch angle threshold. This method balances aerodynamic rationality and engineering feasibility, making it particularly suitable for application in small, low-power tail-seat UAV flight control systems.

Claims

1. A control method for a tail-seat unmanned aerial vehicle (UAV), applied to a tail-seat UAV equipped with a flight control system, wherein the flight control system operates based on a quadcopter control frame, characterized in that, The method includes: Obtain the current pitch angle of the tail-mounted UAV; Determine whether the absolute value of the pitch angle exceeds a preset threshold; In response to the absolute value of the pitch angle exceeding the preset threshold, the system enters a large pitch angle control mode; In the large pitch angle control mode, coordinated turn control is performed. The coordinated turn control includes: closing the lateral translation control channel, keeping the roll control channel open, obtaining the desired roll angle and current airspeed, and obtaining a coordinated yaw rate command based on the desired roll angle, the desired pitch angle currently output by the flight control system, and the current airspeed, and controlling the tail-mounted UAV to perform coordinated turn flight based on the desired roll angle and the coordinated yaw rate command. In the large pitch angle control mode, coupled pitch altitude control is performed. The coupled pitch altitude control includes: based on the altitude control input and the current pitch angle, constructing a virtual horizontal plane by projecting the x-axis direction vector of the aircraft onto the horizontal plane and rotating it around the y-axis of the aircraft by a pitch compensation angle determined by the altitude control input; using the pitch attitude of the virtual horizontal plane as the pitch control target; and using the attitude control loop to make the attitude of the tail-mounted UAV follow the changes of the virtual horizontal plane to achieve altitude adjustment.

2. The method according to claim 1, characterized in that, The process of obtaining a coordinated yaw rate command based on the desired roll angle, the desired pitch angle currently output by the flight control system, and the current airspeed includes: The coordinated yaw rate command is obtained according to the following formula: in, For the coordinated yaw rate command, It is the acceleration due to gravity. The airspeed of the tail-mounted UAV is given. The desired roll angle, The desired pitch angle; The desired pitch angle is the pitch angle setting currently output by the flight control system.

3. The method according to claim 1, characterized in that, The closing of the lateral translation control channel includes: Set the lateral acceleration control command and lateral velocity control command in the body coordinate system to zero, and pause the closed-loop control of lateral translation.

4. The method according to claim 1, characterized in that, The construction of the virtual horizontal plane includes: Obtain the direction vector of the tail-mounted UAV's x-axis in the inertial coordinate system. ; Projecting the direction vector onto the horizontal plane yields the horizontal projection vector. ,in: in, is the unit vector perpendicular to the inertial coordinate system; The surrounding body Axis rotation includes: rotating the horizontal projection vector Surrounding the body The pitch compensation angle is rotated along the axis to obtain a longitudinal reference direction vector for the virtual horizontal plane; the longitudinal reference direction vector is then compared with the body... The cross product along the axis is used to obtain the normal vector of the virtual horizontal plane; wherein, the body axis Let be the horizontal axis of the tail-mounted UAV's body coordinate system; wherein, the pitch compensation angle is obtained by proportional mapping from the altitude control input.

5. The method according to claim 4, characterized in that, The pitch compensation angle is obtained as follows: Multiply the altitude control input by the pitch limit value to obtain the pitch compensation angle; The altitude control input ranges from [-1, 1], and the pitch limit value is the upper limit of the preset virtual horizontal plane pitch angle range.

6. The method according to claim 5, characterized in that, The pitch limit is 30 degrees.

7. The method according to claim 1, characterized in that, Also includes: In response to the absolute value of the pitch angle falling back to the preset threshold range, the large pitch angle control mode is exited; The pitch control target generated based on the virtual horizontal plane is subjected to low-pass filtering attenuation processing; Smoothly transition height control to conventional height control methods.

8. The method according to claim 1, characterized in that, The method of making the attitude of the tail-mounted UAV follow the changes of the virtual horizontal plane through the attitude control loop includes: the flight control system generates throttle output through a multi-loop PID controller to control the position, speed and acceleration state of the tail-mounted UAV on the virtual horizontal plane; The state error is projected onto the vertical direction of the virtual horizontal plane.

9. The method according to claim 1, characterized in that, The coordinated turning control is at the sideslip angle Designed under the assumption of zero; under this assumption, the fuselage of the tail-mounted UAV is symmetrical in the XZ plane, and the lateral aerodynamic coefficient is... ,in For the angle of attack, the coordinated turn control maintains the sideslip angle close to zero by coordinating and matching the roll angle with the yaw rate.

10. The method according to claim 1, characterized in that, The preset threshold value ranges from 70° to 90°. The process of obtaining the current pitch angle of the tail-mounted UAV includes: obtaining the pitch angle through attitude calculation based on the angular velocity data and acceleration data obtained by the inertial measurement unit; The method further includes: acquiring altitude information based on a barometric altimeter and a global navigation satellite system as altitude feedback for the coupled pitch altitude control; and acquiring ground speed and heading information based on the global navigation satellite system as track reference for the coordinated turn control.

Images