Fixed-wing unmanned aerial vehicle flight control method based on three-channel decoupling and dynamic compensation

Abstract

The invention discloses a fixed-wing unmanned aerial vehicle flight control method based on three-channel decoupling and dynamic compensation, and belongs to the technical field of fixed-wing unmanned aerial vehicle flight control. The problem that the attitude control performance is poor due to the fact that aerodynamic interference of the propeller is not considered in an existing method is solved. According to the method, the propeller interference suffered by the fixed-wing unmanned aerial vehicle driven by a propeller propelling system is considered, and the flight control system of the fixed-wing unmanned aerial vehicle is designed and researched under the condition that the propeller interference is considered. Three-channel independent control model design is carried out based on the established kinematics and dynamics models of the fixed-wing unmanned aerial vehicle, a plurality of compensation items are introduced for compensation for the coupling and propeller interference problems, and the simulation analysis is carried out on the designed control system, so that the control precision is improved. The effectiveness of the fixed-wing unmanned aerial vehicle flight control system design method is verified, and the attitude control performance is improved. The method can be applied to flight control of the fixed-wing unmanned aerial vehicle.

Description

Technical Field

[0001] This invention belongs to the field of fixed-wing unmanned aerial vehicle (UAV) flight control technology, specifically relating to a fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation. Background Technology

[0002] Unmanned aerial vehicles (UAVs) are unmanned aircraft that were initially controlled manually via ground-based radio remote control devices or relied on pre-programmed flight procedures for autonomous navigation. In recent years, with breakthroughs in automatic control technology, sensing devices, communication systems, and microelectronics, fixed-wing UAVs have found widespread application in many more fields. For example, they can assist in forest fire early warning through high-altitude monitoring, conduct geographic information collection using aerial surveying technology, perform high-voltage line inspections to ensure power grid safety, cooperate with news organizations to collect aerial images, assist meteorological departments in atmospheric environmental observation, carry out agricultural plant protection spraying and pest control, and undertake critical tasks such as material delivery and disaster assessment in disaster relief. Furthermore, their functions also cover diverse needs such as target positioning and calibration, precision-guided attacks, and electromagnetic spectrum countermeasures.

[0003] Small fixed-wing UAVs represent an important research direction for "miniaturization and lightweighting" and "information intelligence." However, current research on propeller-driven fixed-wing UAVs often neglects the influence of the propeller, which directly affects their attitude control performance. This is especially true for small and micro fixed-wing UAVs, which, due to their small size and light weight, will suffer from poor attitude control performance and difficulty in meeting stability requirements if the current research approach continues to ignore the propeller's impact. Summary of the Invention

[0004] The purpose of this invention is to solve the problem that existing methods ignore the influence of propellers on fixed-wing UAVs, resulting in poor attitude control performance. Therefore, this invention proposes a flight control method for fixed-wing UAVs based on three-channel decoupling and dynamic compensation.

[0005] The technical solution adopted by this invention to solve the above-mentioned technical problems is: a flight control method for fixed-wing unmanned aerial vehicles based on three-channel decoupling and dynamic compensation, the method specifically including the following steps:

[0006] Step 1: Based on the forces and torques acting on the fixed-wing UAV, establish the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV;

[0007] Step 2: Based on the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV, obtain the control model of the fixed-wing UAV, and then design the three-channel control system of pitch, yaw and roll according to the control model of the fixed-wing UAV.

[0008] The beneficial effects of this invention are:

[0009] This invention considers propeller interference experienced by fixed-wing UAVs driven by propeller propulsion systems, and designs and studies the flight control system of fixed-wing UAVs under the consideration of propeller interference. Based on the established kinematic and dynamic models of the fixed-wing UAV, three-channel independent control models are designed. Several compensation terms are introduced to address coupling and propeller interference issues. Simulation analysis of the designed control system verifies the effectiveness of the fixed-wing UAV flight control system design method of this invention, improving attitude control performance. This method is particularly applicable to small fixed-wing UAVs. Attached Figure Description

[0010] Figure 1 This is the pitch channel control block diagram;

[0011] The output is obtained after limiting. ;

[0012] Figure 2 This is a block diagram of the yaw channel control.

[0013] Figure 3 This is the control block diagram for the roll channel;

[0014] Figure 4 These are the pitch channel step response diagrams at feature point 1 and feature point 3;

[0015] Figure 5 These are Bode plots of the pitch channel at feature point 1 and feature point 3;

[0016] Figure 6 This is the pitch channel step response diagram at feature point 2;

[0017] Figure 7 This is the Bode plot of the pitch channel at feature point 2;

[0018] Figure 8 These are the pitch channel step response diagrams at feature points 4 and 6;

[0019] Figure 9 These are Bode plots of the pitch channel at feature point 4 and feature point 6;

[0020] Figure 10 This is the pitch channel step response diagram at feature point 5;

[0021] Figure 11 This is the Bode plot of the pitch channel at feature point 5;

[0022] Figure 12 These are the pitch channel step response diagrams at feature points 7 and 9;

[0023] Figure 13 These are Bode plots of the pitch channel at feature points 7 and 9;

[0024] Figure 14 This is a step response diagram of the pitch channel at feature point 8;

[0025] Figure 15 This is the Bode plot of the pitch channel at feature point 8;

[0026] Figure 16 This is the step response diagram of the compensated pitch channel at feature point 1;

[0027] Figure 17 This is the step response diagram of the yaw channel before compensation at feature point 1;

[0028] Figure 18 This is the step response diagram of the roll channel before compensation at feature point 1;

[0029] Figure 19 This is the step response diagram of the pitch channel after compensation at feature point 1;

[0030] Figure 20 This is the step response diagram of the yaw channel after compensation at feature point 1;

[0031] Figure 21 This is the step response diagram of the roll channel after compensation at feature point 1. Detailed Implementation

[0032] This invention first defines commonly used coordinate systems and provides transformation matrices between them.

[0033] 1. Ground coordinate system

[0034] A ground coordinate system is a coordinate system fixed on the Earth's surface, with its origin at... This is usually set to the location of the drone's ground launch point. The axis can point in any direction to the ground plane. The axis is vertically upward. axis, shaft and The axes form a right-handed coordinate system.

[0035] 2. Projectile coordinate system

[0036] The missile body coordinate system is a coordinate system fixed to the aircraft body, with the UAV's center of mass as the origin. , The axis coincides with the longitudinal axis of the aircraft. The positive axis points towards the head of the drone; The axis is located within the longitudinal symmetry plane of the aircraft, and... The axis is vertical and positive is indicated by pointing upwards; The axial direction is determined according to the right-hand rectangular coordinate system and is perpendicular to the longitudinal symmetry plane of the aircraft.

[0037] 3. Ballistic coordinate system

[0038] Origin of the ballistic coordinate system For the center of mass of the projectile, The positive direction of the axis is consistent with the direction of the instantaneous velocity vector of the aircraft. The axis is located in the area containing In the vertical plane of the axis, Axis perpendicular to The axis is positive when it points upwards. The axis is perpendicular to the other two axes and is determined by the right-hand rule.

[0039] 4. Velocity coordinate system

[0040] The velocity coordinate system has the projectile's center of mass as its origin. , The axis coincides with the instantaneous velocity vector of the aircraft. The axis lies within the longitudinal plane of symmetry of the aircraft and is parallel to... The axis is vertical, with the upward direction considered positive. Axis perpendicular to For planes, direction is determined by the right-hand rule.

[0041] The coordinate system transformation relationships are as follows:

[0042] (1) Ground coordinate system and projectile coordinate system

[0043] The relationship between the UAV's ground coordinate system and the missile's body coordinate system can be established through the pitch angle. Yaw angle and roll angle The description states that, through coordinate transformation, the rotation matrix from the ground coordinate system to the projectile coordinate system can be derived. :

[0044] (1)

[0045] (2) Ground coordinate system and ballistic coordinate system

[0046] The relationship between the ground coordinate system and the ballistic coordinate system can be established through the ballistic inclination angle. and ballistic deflection The description states that, through coordinate transformation, the rotation matrix from the ground coordinate system to the ballistic coordinate system can be derived. :

[0047] (2)

[0048] (3) Velocity coordinate system and projectile coordinate system

[0049] because shaft and The axes are all located in the longitudinal plane of symmetry of the UAV, therefore the transformation matrix between the velocity coordinate system and the projectile coordinate system is related to the angle of attack. and sideslip angle Based on this, the rotation matrix from the velocity coordinate system to the projectile coordinate system can be derived through coordinate transformation. :

[0050] (3)

[0051] (4) Velocity coordinate system and ballistic coordinate system

[0052] because shaft and Since the axes coincide, the transformation relationship between the velocity coordinate system and the ballistic coordinate system can be obtained through the velocity tilt angle. This can be determined by coordinate transformation. The rotation matrix from the ballistic coordinate system to the velocity coordinate system can be derived from this transformation. :

[0053] (4)

[0054] The method of the present invention will be further described in detail below based on the defined coordinate system and coordinate system transformation relationship.

[0055] Specific Implementation Method 1: The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation described in this implementation method specifically includes the following steps:

[0056] Step 1: Based on the forces and torques acting on the fixed-wing UAV, establish the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV;

[0057] Specifically, the forces acting on the fixed-wing UAV include thrust, gravity, and aerodynamics; the torque acting on the fixed-wing UAV includes the reaction torque generated by the propeller. and gyro torque .

[0058] This invention mainly considers two factors affecting fixed-wing UAVs: reaction torque and gyroscopic effect.

[0059] (1) Reaction torque

[0060] When a propeller rotates, it churns and rotates the surrounding air. The torque generated in the opposite direction of this churning is called the propeller's reaction torque. This torque causes the drone to tilt in the opposite direction of the propeller's rotation. According to leaf element theory, the reaction torque generated by the propeller is proportional to the square of the propeller's rotational speed. For example, in a right-turning propeller drone, the drone will tilt to the left (roll). Especially for small fixed-wing drones, because their moment of inertia is smaller, the reaction torque has a greater impact on the fuselage; therefore, the reaction torque is taken into account in the modeling.

[0061] (2) Gyroscope effect

[0062] When a propeller rotates at high speed during flight, if it is subjected to a torque that changes the direction of its axis, it will not rotate exactly in the direction of that torque. Instead, it will produce an additional deflection, a phenomenon known as propeller precession. A gyroscope, during precession, will generate a reaction torque and a gyroscopic torque. For example, in a right-hand propeller aircraft, when the pilot attempts to pitch the aircraft up, the propeller will not only pitch up but also deflect to the right, causing the nose of the aircraft to also deflect to the right. Therefore, the gyroscopic effect generated by the propeller is taken into account when modeling.

[0063] It should be noted that the torques actually acting on fixed-wing UAVs include not only the reaction torque and gyroscopic torque generated by the propeller, but also other torques, such as angular acceleration torque. Since the research object of this invention is to consider the effect of the propeller, this embodiment does not list all the other torques acting on fixed-wing UAVs considered by this invention.

[0064] The aerodynamic forces acting on a small fixed-wing UAV are generated by factors such as the wings, body, and control surfaces. The components of these aerodynamic forces along the three axes of the velocity coordinate system are respectively... , and :

[0065] (5)

[0066] In the formula: Indicates resistance. Indicates lift. Indicates lateral force; This represents the dimensionless drag coefficient. Indicates the dimensionless lift coefficient; Indicates the dimensionless lateral force coefficient; Indicates dynamic pressure; Indicates the reference area of ​​a fixed-wing unmanned aerial vehicle;

[0067] (6)

[0068] In the formula: This indicates the speed of a fixed-wing drone; Indicates air density; aerodynamic coefficients are respectively:

[0069] (7)

[0070] In the formula: Indicates the zero-lift drag coefficient; Indicates the induced drag coefficient; Represents the zero angle-of-attack lift coefficient (given by the UAV relative to the angle of attack). (The asymmetry between the upper and lower parts of the plane is generated). This represents the derivative of the lift coefficient with respect to the angle of attack; This represents the derivative of the lift coefficient with respect to the elevator; This represents the derivative of the lateral force coefficient with respect to the sideslip angle; This represents the derivative of the lateral force coefficient with respect to the rudder. Indicates the angle of attack; Indicates the sideslip angle.

[0071] The magnitude of gravity acting on the fixed-wing UAV in the ground coordinate system. for:

[0072] (8)

[0073] In the formula: Indicates the quality of the drone. Let gravitational acceleration be the acceleration due to gravity; assuming the mass of the drone remains constant and the thrust *r* is related to the propeller speed, then the magnitude of the thrust acting on the fixed-wing drone is... for:

[0074] (9)

[0075] In the formula: Indicates air density; Indicates the diameter of the paddle tray; Indicates the motor speed; Represents dimensionless aerodynamic parameters; the torque of thrust acting on the UAV through the UAV's center of mass along the three axes of the projectile coordinate system. , and They are respectively:

[0076] (10)

[0077] In the formula: Indicates the rolling moment coefficient; Indicates the yaw moment coefficient; Indicates the pitching moment coefficient; This represents the reaction torque generated by the propeller. Indicates angular acceleration torque. This indicates the characteristic length of a fixed-wing unmanned aerial vehicle (UAV). This represents the gyroscopic torque generated by propeller precession along the projectile coordinate system. Axial components; This represents the gyroscopic torque generated by propeller precession along the projectile coordinate system. Shaft components; moment coefficient , and They are respectively:

[0078] (11)

[0079] In the formula: Zero sideslip rolling moment coefficient (caused by shape asymmetry due to production errors); Zero sideslip yaw moment coefficient (due to the UAV's shape relative to...) Because it is planar symmetric, its value is 0. The zero-lift pitch moment coefficient (due to the UAV's shape relative to...) (This is caused by planar asymmetry) This is the derivative of the rolling moment coefficient with respect to the sideslip angle; This is the derivative of the yaw moment coefficient with respect to the sideslip angle; This is the derivative of the pitch moment coefficient with respect to the angle of attack; This refers to the roll deflection angle of the rudder. Indicates the yaw angle; Indicates the pitch deflection angle; This is the control torque coefficient for the roll rudder; This is the control torque coefficient for the yaw rudder; This is the pitch control torque coefficient; This is the derivative of the yaw moment with respect to the roll angle; This is the derivative of the rolling moment with respect to the yaw angle; This represents the roll angular velocity in the projectile coordinate system; The yaw rate in the projectile coordinate system; The pitch angular velocity in the projectile coordinate system; The coefficient of dimensionless rolling damping moment; The dimensionless yaw damping moment coefficient; The dimensionless pitch damping moment coefficient;

[0080] Dimensionless roll angular velocity Yaw angular velocity with dimensionless 1 Pitch angular velocity with dimensionless value of 1 ; This is the derivative of the yaw moment with respect to the dimensionless roll velocity; This is the derivative of the rolling torque with respect to the dimensionless yaw rate; express The first derivative; This is the derivative of the pitch moment with respect to the angle of attack; for The first derivative; This is the derivative of the yaw moment with respect to the sideslip angle; express The first derivative; This is the derivative of the pitch moment with respect to the pitch deflection angle; express The first derivative; Let be the derivative of the yaw moment with respect to the yaw angle; According to blade element theory, the reaction torque generated by the propeller has the following formula:

[0081] (12)

[0082] in: This refers to the propeller speed; It is related to propeller-related parameters and air density, and can be treated as a constant in the analysis; The sign is related to the direction of propeller rotation. Viewed from the tail, if the propeller rotates clockwise, then... Take the negative value, and vice versa. If we take the positive reference direction for the propeller, then the reaction torque generated by the propeller can be expressed as:

[0083] (13)

[0084] in: Indicates the propeller speed. express The absolute value; The angular acceleration torque generated by the propeller is constant; it is expressed as:

[0085] (14)

[0086] in: This represents the moment of inertia of the shaft and propeller, equivalent to that of the shaft.

[0087] Gyroscopic torque generated by propeller precession Along the projectile coordinate system shaft and The components of the axis are represented as follows:

[0088] (15)

[0089] (16)

[0090] The establishment of the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV are specifically as follows:

[0091] 1. The system of equations for the center of mass dynamics in the dynamic model of a fixed-wing UAV

[0092] According to Newton's laws of motion, the vector representation of the UAV's center-of-mass dynamic equation is as follows:

[0093] (17)

[0094] In the formula: The velocity of the fixed-wing UAV in the ground coordinate system. express The absolute derivative in the ground coordinate system, express The absolute derivative in a moving coordinate system (projectile body, trajectory, velocity system), For engine thrust, For aerodynamics (synthesized from X, Y, and Z). For gravity, This represents the angular velocity vector of the moving coordinate system relative to the ground coordinate system. Represents the mass of the drone; in ballistic coordinates, the velocity... Represented as:

[0095] (18)

[0096] in: Indicates speed In the ballistic coordinate system Axial direction component; Indicates speed In the ballistic coordinate system Axial direction component; Indicates speed In the ballistic coordinate system Axial direction components; then, in the ballistic coordinate system The absolute derivative is:

[0097] (19)

[0098] in: express The derivative of the angular velocity of the ballistic coordinate system relative to the ground coordinate system; for:

[0099] (20)

[0100] in: Indicates the trajectory deflection angle. express The first derivative; Indicates the trajectory inclination angle, express The first derivative; express In the ballistic coordinate system Axial direction component; express In the ballistic coordinate system Axial direction component; express In the ballistic coordinate system Axial direction component;

[0101] According to equations (18) and (20), we get:

[0102] (twenty one)

[0103] Substituting equations (19) and (21) into equation (17), we get:

[0104] (twenty two)

[0105] in: Indicates the thrust in the ballistic coordinate system Axial direction component; Indicates the thrust in the ballistic coordinate system Axial direction component; Indicates the thrust in the ballistic coordinate system Axial direction component; Represents aerodynamics in the ballistic coordinate system Axial direction component; Represents aerodynamics in the ballistic coordinate system Axial direction component; Represents aerodynamics in the ballistic coordinate system Axial direction component; Represents gravity in the ballistic coordinate system Axial direction component; Represents gravity in the ballistic coordinate system Axial direction component; Represents gravity in the ballistic coordinate system Axial direction component;

[0106] Assuming thrust The direction relative to the projectile coordinate system If the axes coincide, then the expression for the thrust in the ballistic coordinate system is:

[0107] (twenty three)

[0108] Based on the expression for aerodynamics in the velocity coordinate system, a coordinate transformation is performed to obtain the expression for aerodynamics in the ballistic coordinate system:

[0109] (twenty four)

[0110] Based on the expression for gravity in the ground coordinate system, a coordinate transformation is performed to obtain the expression for gravity in the ballistic coordinate system:

[0111] (25)

[0112] Substituting equations (23), (24), and (25) into equation (22), we obtain the dynamic equation for the motion of the UAV's center of mass in the ballistic coordinate system as follows:

[0113] (26)

[0114] In the velocity coordinate system, the velocity Represented as:

[0115] (27)

[0116] in: Indicates speed In the velocity coordinate system Axial direction component; Indicates speed In the velocity coordinate system Axial direction component; Indicates speed In the velocity coordinate system Axial direction component;

[0117] So, in the velocity coordinate system The absolute derivative is:

[0118] (28)

[0119] The rotational angular velocity of the velocity coordinate system relative to the ground coordinate system Represented as:

[0120] (29)

[0121] in: This represents the angular velocity vector of the projectile's coordinate system relative to the ground coordinate system; This represents the angular velocity vector of the projectile coordinate system relative to the velocity coordinate system; express In the velocity coordinate system Axial direction component; express In the velocity coordinate system Axial direction component; express In the velocity coordinate system Axial direction component;

[0122] The angular velocity vector of the projectile's coordinate system relative to the ground coordinate system can be obtained through coordinate transformation. for:

[0123] (30)

[0124] The angular velocity vector of the projectile coordinate system relative to the velocity coordinate system is:

[0125] (31)

[0126] in, In the projectile coordinate system On the axis, In the velocity coordinate system On the axis, according to the coordinate transformation, we get:

[0127] (32)

[0128] Substituting equations (30) and (32) into equation (29), we obtain the rotational angular velocity of the velocity coordinate system relative to the ground coordinate system. for:

[0129] (33)

[0130] According to equations (27) and (33), we get:

[0131] (34)

[0132] Assuming thrust The direction relative to the projectile coordinate system If the axes coincide, then the expression for the thrust in the velocity coordinate system is:

[0133] (35)

[0134] in: Represents thrust in the velocity coordinate system Axial direction component; Represents thrust in the velocity coordinate system Axial direction component; Indicates the thrust in the ballistic coordinate system Axial direction component;

[0135] According to the expression for aerodynamics in the velocity coordinate system, we have:

[0136] (36)

[0137] in: Representing aerodynamics in the velocity coordinate system Axial direction component; Representing aerodynamics in the velocity coordinate system Axial direction component; Representing aerodynamics in the velocity coordinate system Axial component; Based on the expression for gravity in the ground coordinate system, perform a coordinate transformation to obtain the expression for gravity in the velocity coordinate system:

[0138] (37)

[0139] in: Represents gravity in the velocity coordinate system Axial direction component; Represents gravity in the velocity coordinate system Axial direction component; Represents gravity in the ballistic coordinate system Axial direction component;

[0140] Substituting equations (28), (34), (35), (36), and (37) into equation (17), we obtain the dynamic equation for the motion of the UAV's center of mass in the velocity coordinate system as follows:

[0141] (38)

[0142] Rearranging equation (38), we get:

[0143] (39)

[0144] 2. The dynamic equations of rotation about the center of mass in the kinematic model of a fixed-wing UAV

[0145] In the projectile coordinate system, the dynamic equation of the UAV rotating about its center of mass is:

[0146] (40)

[0147] In the formula: Angular momentum express The absolute derivative in the ground coordinate system, express The absolute derivative in a moving coordinate system For aerodynamic torque, The angular velocity of the moving coordinate system relative to the ground coordinate system;

[0148] In the projectile coordinate system, the angular momentum is expressed as:

[0149] (41)

[0150] in: Represents the coordinate system of the projectile. Moment of inertia of the shaft; Represents the coordinate system of the projectile. Moment of inertia of the shaft; Represents the coordinate system of the projectile. Moment of inertia of the shaft; The angular momentum in the projectile coordinate system is represented by Axial direction component; The angular momentum in the projectile coordinate system is represented by Axial direction component; The angular momentum in the projectile coordinate system is represented by Axial direction component; Represents the projectile's position relative to the projectile's coordinate system. shaft and The product of inertia of the axis; Represents the projectile's position relative to the projectile's coordinate system. shaft and The product of inertia of the axis; Represents the projectile's position relative to the projectile's coordinate system. shaft and The product of inertia of the axis;

[0151] For the model of the present invention, ,So, The expression in the projectile coordinate system is:

[0152] (42)

[0153] but The absolute derivative in the moving coordinate system is:

[0154] (43)

[0155] in, express The first derivative, express The first derivative, express The first derivative;

[0156] angular velocity of the projectile coordinate system relative to the ground coordinate system for:

[0157] (44)

[0158] So, The expression is:

[0159] (45)

[0160] Substituting equations (43) and (45) into equation (40), we obtain the dynamic equation for the rotation about the center of mass:

[0161] (46)

[0162] in, , and It is given by the following formula:

[0163] Generally speaking, it is believed that , .

[0164] 3. The kinematic equations of the center of mass in the kinematic equations of a fixed-wing UAV

[0165] Establish the position equation of the body's center of mass relative to the ground coordinate system:

[0166] (47)

[0167] in: Represents the body velocity in the ground coordinate system Axial direction component; Represents the body velocity in the ground coordinate system Axial direction component; Represents the body velocity in the ground coordinate system Axial component; superscript T indicates transpose;

[0168] As defined by the ballistic coordinate system, the velocity direction of the projectile's center of mass is along the ballistic coordinate system. Axis, that is:

[0169] (48)

[0170] Using the transformation relationship between the ballistic coordinate system and the ground coordinate system, we get:

[0171] (49)

[0172] in, Let represent the rotation matrix from the ballistic coordinate system to the ground coordinate system; substituting equation (49) into equation (47), we get:

[0173] (50)

[0174] 4. The kinematic equations of a fixed-wing UAV involving rotation about its center of mass.

[0175] To determine the projectile's attitude in space, it is necessary to establish the kinematic equations for its rotation around its center of mass, i.e., to establish the attitude angles. , and The rate of change is related to the angular velocity component of the projectile relative to the ground coordinate system. , and The relationship between them.

[0176] Based on the transformation relationship between the ground coordinate system and the projectile coordinate system, the angular velocity of the projectile coordinate system relative to the ground coordinate system is:

[0177] (51)

[0178] in, express The first derivative, express The first derivative, express The first derivative;

[0179] With ground coordinate system Axis coincidence, With respect to the projectile coordinate system If the axes coincide, then:

[0180] (52)

[0181] After rearranging equation (52), we get:

[0182] (53)

[0183] 5. Overload Equations for Fixed-Wing Unmanned Aerial Vehicles

[0184] Based on equations (23) and (25), the projection of the overload vector onto the ballistic coordinate system is obtained:

[0185] (54)

[0186] in, For tangential overload in the ballistic coordinate system, and For ballistic coordinate system normal overload;

[0187] Based on equations (35) and (36), the projection of the overload vector onto the velocity coordinate system is obtained:

[0188] (55)

[0189] in, For tangential overload in the velocity coordinate system, and For overload in the velocity coordinate system normal;

[0190] According to equation (54), the projection of the overload vector onto the projectile coordinate system is obtained as follows:

[0191] (56)

[0192] in, For longitudinal overload of the projectile coordinate system, and This refers to the lateral overload of the projectile coordinate system.

[0193] Step 2: Based on the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV, obtain the control model of the fixed-wing UAV, and then design the three-channel control system of pitch, yaw and roll according to the control model of the fixed-wing UAV.

[0194] Specific Implementation Method Two: This implementation method differs from Specific Implementation Method One in that, based on the kinematic and dynamic model of the fixed-wing UAV and its overload equations, the control model of the fixed-wing UAV is obtained, specifically as follows:

[0195] Based on the classic decoupling control strategy, several assumptions are made:

[0196] 1. The UAV motion model is a small-disturbance motion model, meaning that the actual flight trajectory is not much different from the ideal flight trajectory;

[0197] 2. Ignore the influence of the drone's own gravity and add a gravity compensation term to the guidance law design;

[0198] 3. Considering the motion of the UAV within a short period, it is assumed that the state parameters of the UAV change very slowly within a short period;

[0199] 4. Angle of attack during drone flight Sideslip angle As a small approximation, the angle of attack and sideslip angle are both considered as first-order small quantities, and , , , Ignore minor quantities of second order and above, i.e. ;

[0200] 5. Calculate the angular velocity component during the drone's flight. , Consider them as small quantities; small quantities of second order and above can be ignored.

[0201] 6. Without considering the interference caused by the drone propellers, the impact of the propellers will be compensated by introducing a compensation term later.

[0202] Based on the kinematic and dynamic model and overload equations of the fixed-wing UAV, and neglecting coupling terms, the control model of the fixed-wing UAV is obtained:

[0203] (57)

[0204] in, Indicates angle of attack. express The first derivative, This represents the roll angular velocity in the projectile coordinate system; The yaw rate in the projectile coordinate system; The pitch angular velocity in the projectile coordinate system; Indicates the sideslip angle. for The first derivative, Indicates dynamic pressure. This indicates the magnitude of the thrust acting on a fixed-wing drone. This represents the reference area of ​​a fixed-wing UAV. Indicates the quality of the drone. This indicates the speed of a fixed-wing drone. This represents the derivative of the lift coefficient with respect to the angle of attack. This represents the derivative of the lift coefficient with respect to the elevator. Indicates the pitch deflection angle. This represents the derivative of the lateral force coefficient with respect to the sideslip angle. This represents the derivative of the lateral force coefficient with respect to the rudder. Indicates the yaw angle. This indicates the characteristic length of a fixed-wing unmanned aerial vehicle (UAV). This is the derivative of the rolling moment coefficient with respect to the sideslip angle. Represents the coordinate system of the projectile. moment of inertia of the shaft Let be the derivative of the rolling torque with respect to the dimensionless yaw rate. Let be the derivative of the yaw moment with respect to the dimensionless roll velocity. Let be the derivative of the rolling moment with respect to the yaw angle. This refers to the roll deflection angle of the rudder. Indicates the yaw angle. Represents the coordinate system of the projectile. moment of inertia of the shaft Represents the coordinate system of the projectile. moment of inertia of the shaft express The first derivative, express The first derivative, express The first derivative, The dimensionless yaw damping moment coefficient, Let be the derivative of the yaw moment with respect to the roll angle. This is the control torque coefficient of the yaw rudder. This represents the derivative of the yaw moment with respect to the dimensionless pitch velocity. The dimensionless pitch damping moment coefficient, Let be the derivative of the pitch moment coefficient with respect to the angle of attack. This is the pitch control torque coefficient. This represents the dimensionless drag coefficient. The dimensionless rolling damping moment coefficient, This is the control torque coefficient for the roll rudder. This is the derivative of the yaw moment coefficient with respect to the sideslip angle. The derivative of the roll angle, Indicates the quality of the drone. It is the acceleration due to gravity;

[0205] For ease of representation, the aerodynamic parameters of each part of the above equation can be defined as follows:

[0206] (58)

[0207] in, All are aerodynamic parameters;

[0208] Substituting the aerodynamic parameters in equation (58) into equation (57), we obtain the control model for the fixed-wing UAV:

[0209] (59)

[0210] in, and This is a normal overload in the ground coordinate system.

[0211] The other steps and parameters are the same as in Specific Implementation Method 1.

[0212] Specific implementation method three: Combining Figure 1This embodiment is described below. The difference between this embodiment and specific embodiments one or two is that the design method of the pitch channel control system is as follows:

[0213] After processing the coupling terms in equation (59), a control model for the pitch channel is established:

[0214] (60)

[0215] Taking the Laplace transform of equation (60), we obtain the transfer function of equation (61):

[0216] (61)

[0217] in, Represents the Laplace variable. The Laplace transform of the pitch angular velocity signal. The Laplace transform of the pitch deflection signal. The Laplace transform of the angle-of-attack signal. Represents the Laplace transform of the normal overload signal;

[0218] According to equation (57), the coupling term of the pitch channel is: , Meanwhile, the pitch channel is affected by the gyro torque. Therefore, the following compensation term is introduced for the pitch channel. For example... Figure 1 As shown, the rudder deflection command for the pitch channel is then obtained according to equation (61). for:

[0219] (62)

[0220] in, For pitch channel compensation, This represents the propeller throttle opening (value range: 0~1). , and These are the PID control parameters for the pitch channel. , , , , and These are all design parameters for the pitch channel; Here is the transfer function of the rate gyroscope. The transfer function of the accelerometer. This indicates a normal overload command for the pitch channel.

[0221] Other steps and parameters are the same as in specific implementation method one or two.

[0222] In this embodiment, the outer loop controller employs PID control. The proportional control stage directly responds to the overload signal, improving the system's dynamic response speed. The integral control stage continuously accumulates system errors, effectively eliminating the steady-state error between the input signal and the actual output, thus improving control accuracy. The derivative control, by introducing pitch rate feedback, significantly enhances the system's damping characteristics, effectively suppressing overshoot and significantly reducing system oscillations. The introduced angular rate and angle-of-attack feedback mechanisms, through real-time compensation for aerodynamic parameter changes, enable the control system to possess excellent flight speed adaptability. This composite control architecture, through the synergistic effect of each stage, achieves fast and accurate overload tracking performance while ensuring system stability, enabling it to adapt to complex and ever-changing flight environments. The invention introduces several compensation terms for the rudder deflection command to eliminate interference from coupling terms and the propeller, thereby improving system performance.

[0223] Specific Implementation Method Four: This implementation method differs from Specific Implementation Methods One to Three in that it simplifies the rate gyroscope system into a second-order system, and the transfer function of the rate gyroscope is:

[0224] (63)

[0225] In the formula, The rate gyroscope damping ratio, is the undamped oscillation frequency of the rate gyroscope.

[0226] The other steps and parameters are the same as those in one of the specific implementation methods one to three.

[0227] In this invention, , Therefore, this invention yields the transfer function of the rate gyroscope:

[0228] (64)

[0229] Specific Implementation Method Five: This implementation method differs from Specific Implementation Methods One through Four in that it simplifies the accelerometer system into a second-order system, and the accelerometer's transfer function is:

[0230] (65)

[0231] In the formula, For the accelerometer damping ratio, is the undamped oscillation frequency of the accelerometer.

[0232] The other steps and parameters are the same as those in one of the specific implementation methods one to four.

[0233] In this invention, , Thus, the transfer function of the accelerometer is obtained:

[0234] (66)

[0235] Specific Implementation Method Six: Combination Figure 2 This embodiment is described below. The difference between this embodiment and one of the specific embodiments one to five is that the design method of the yaw channel control system is as follows:

[0236] After processing the coupling terms in equation (59), a control model for the yaw channel is established:

[0237] (67)

[0238] Applying a Laplace transform to equation (67) yields the transfer function of equation (68):

[0239] (68)

[0240] in, This is the Laplace transform of the yaw rate signal. This is the Laplace transform of the yaw angle signal. This is the Laplace transform of the sideslip angle signal. This is the Laplace transform form of the transverse overload signal;

[0241] According to equation (57), the coupling term of the yaw channel is: , Meanwhile, the yaw channel is affected by the gyro torque. Therefore, the following compensation terms are introduced for the yaw channel. For example... Figure 2 As shown, the rudder deflection angle command for the yaw channel is then obtained according to equation (68). for:

[0242] (69)

[0243] in, This is a compensation item for deviation from the designated route. , and These are all PID control parameters for the yaw channel. , , , , and These are all design parameters for the yaw channel. This is a lateral overload command for the yaw channel.

[0244] The other steps and parameters are the same as those in one of the specific implementation methods one to five.

[0245] In this embodiment, the outer loop controller uses PID control, and angular rate feedback and sideslip angle feedback mechanisms are introduced.

[0246] Specific implementation method seven: Combination Figure 3 This embodiment is described below. The difference between this embodiment and one of the specific embodiments one through six is ​​that the design method of the roll channel control system is as follows:

[0247] After processing the coupling terms in equation (59), a control model for the roll channel is established:

[0248] (70)

[0249] in, Indicates roll angle The first derivative;

[0250] Performing a Laplace transform on equation (70) yields the transfer function of equation (71):

[0251] (71)

[0252] in, This is the Laplace transform of the roll angular velocity signal. This is the Laplace transform of the roll deflection angle signal. The Laplace transform of the roll angle signal;

[0253] According to equation (57), the coupling term of the rolling channel is: Meanwhile, the effects of the reaction torque and angular acceleration torque on the rolling channel are expressed as follows: Therefore, the following compensation term is introduced for the roll path. For example... Figure 3 As shown, the rudder deflection angle command for the roll channel is obtained according to equation (71):

[0254] (72)

[0255] in, For the roll-through compensation item, and These are all PD controller parameters. , , and These are all design parameters for the roll channel. This is the command roll angle for the roll channel.

[0256] The other steps and parameters are the same as those in one of the specific implementation methods one to six.

[0257] In this embodiment, the outer loop controller uses PD control, and roll angle feedback and roll rate feedback mechanisms are introduced to enable the control system to meet the performance requirements.

[0258] In this invention, the control commands for all three channels are used as inputs to the transfer function of the rudder system. While a typical rudder system is a third-order nonlinear system, this invention simplifies the rudder system to a second-order system. Therefore, the transfer function of the rudder system used in this invention is:

[0259] (73)

[0260] In the formula, The damping ratio of the rudder system, This is the undamped oscillation frequency of the rudder system. In this invention, , Thus, the transfer function of the rudder system is obtained:

[0261] (74)

[0262] Before verifying the performance of the method of the present invention, the present invention first selected nine drone states for subsequent verification, and each drone state was referred to as a feature point:

[0263] Feature point 1 represents the drone descending with a slight leftward glide; feature point 2 represents the drone descending in a straight line; feature point 3 represents the drone descending with a slight rightward glide; feature point 4 represents the drone flying horizontally with a slight leftward glide; feature point 5 represents the drone flying horizontally without a sideslip and in symmetrical flight; feature point 6 represents the drone flying horizontally with a slight rightward glide; feature point 7 represents the drone climbing with a slight leftward glide; feature point 8 represents the drone climbing in a straight line; feature point 9 represents the drone climbing with a slight rightward glide. In this invention, the angle of attack changes for all states need to cover three longitudinal states: descent (-4°), level flight (0°), and climb (4°). The sideslip angle changes for all states need to cover three lateral states: left sideslip (-4°), no sideslip (0°), and right sideslip (4°). Nine feature points cover the main attitude combinations of the UAV within its low-altitude flight envelope. Symmetrical flight includes level flight, straight climb, and straight descent; asymmetrical flight includes turns with sideslip; and extreme states include high angle-of-attack climbs with sideslip and high angle-of-attack descents with sideslip. During verification, the control systems of each channel were used to control the states corresponding to all feature points.

[0264] The specific experimental details are as follows:

[0265] The fixed-wing UAV studied in this invention exhibits a good linear relationship between aerodynamic parameters and related physical quantities when the angle of attack and sideslip angle are both small. Data from the UAV at corresponding feature points in a real-world scenario are selected to calculate the aerodynamic parameters. A control system is designed and simulated at these selected feature points to obtain the corresponding controller parameters and system performance, providing a data foundation for real-time interpolation between feature points. Subsequently, interpolation processing is performed on the data between feature points to obtain control system parameters within a certain range. The parameters for each feature point are shown in Table 1.

[0266] Table 1 Feature Point Value Table

[0267]

[0268] The following is a three-channel simulation:

[0269] 1. Pitch Channel Simulation

[0270] The pitch channel control system designed in this invention is as follows: Figure 1 As shown, simulation debugging was performed at each feature point. First, the PID coefficients and corresponding parameters were coarsely adjusted. After obtaining indicators that met the design requirements, fine-tuning was performed to finally obtain the control system parameters that met the design requirements.

[0271] It should be noted that the coarse-tuning process involves: calculating control parameters based on control theory and the UAV dynamics model; substituting these parameters into the control system for step response or frequency response simulation; adjusting the parameter range if the system is unstable or the response is too slow until the system basically meets the design requirements (rise time, settling time, overshoot, etc.); and then completing the coarse-tuning process. The fine-tuning process involves: after obtaining parameters that better meet the design requirements, fine-tuning is performed by finely adjusting each parameter according to specific performance indicators (phase margin, crossover frequency, gain margin, etc.) until the control system parameters that meet the design requirements are obtained.

[0272] The limiting stage restricts the rudder deflection angle at... Taking feature point 1 and feature point 3 as examples, substituting the corresponding aerodynamic data at feature point 1 and feature point 3 into equation (61) yields the transfer function of the corresponding link:

[0273] (75)

[0274] By substituting the transfer function into the simulation program and adjusting the parameters, a set of parameters with better system performance can be obtained. At feature points 1 and 3, the step response and Bode plot of the pitch channel are as follows: Figure 4 and Figure 5 As shown.

[0275] The simulation results show that the system has no steady-state error at feature point 1 and feature point 3, the rise time to reach 90% of the steady-state value is 1.1s, the settling time under the 2% error band is 1.5s, the system has no overshoot, the phase margin is 65.3°, the crossover frequency is 1.62rad / s, and the gain margin is 6.99dB, which meets the design requirements and performance indicators.

[0276] (2) Feature point 2

[0277] Substituting the corresponding aerodynamic data at feature point 2 into the pitch channel transfer function yields the transfer function for the corresponding component. Substituting this transfer function into the simulation program and adjusting the parameters results in a set of parameters that provide better system performance. At feature point 2, the step response and Bode plot of the pitch channel are as follows: Figure 6 and Figure 7 As shown.

[0278] The simulation results show that the system has no steady-state error at feature point 2, the rise time to 90% of the steady-state value is 0.9s, the peak time is 1.4s, the settling time under the 2% error band is 1.8s, the system overshoot is 3.3%, the phase margin is 56.2°, the crossover frequency is 2.30rad / s, and the gain margin is 4.03dB, which meets the design requirements and performance indicators.

[0279] (3) Feature point 4 and feature point 6

[0280] Substituting the corresponding aerodynamic data at feature points 4 and 6 into the pitch channel transfer function yields the transfer function for the corresponding component. Substituting this transfer function into the simulation program and adjusting the parameters results in a set of parameters that provide better system performance. At feature points 4 and 6, the step response and Bode plot of the pitch channel are as follows: Figure 8 and Figure 9 As shown.

[0281] Simulation results show that at feature points 4 and 6, the system has no steady-state error, the rise time to 90% of the steady-state value is 1.0s, the settling time under the 2% error band is 1.4s, the system has no overshoot, the phase margin is 64.4°, the crossover frequency is 1.99rad / s, and the gain margin is 4.09dB, which meets the design requirements and performance indicators.

[0282] (4) Feature point 5

[0283] Substituting the corresponding aerodynamic data at feature point 5 into the pitch channel transfer function yields the transfer function for the corresponding component. Substituting this transfer function into the simulation program and adjusting the parameters results in a set of parameters that provide better system performance. At feature point 5, the step response and Bode plot of the pitch channel are as follows: Figure 10 and Figure 11 As shown.

[0284] According to the simulation results, at feature point 5, the system has no steady-state error, the rise time to 90% steady-state value is 0.9s, the peak time is 1.4s, the settling time under the 2% error band is 1.9s, the system overshoot is 3.5%, the phase margin is 55.9°, the crossover frequency is 2.29rad / s, and the gain margin is 3.79dB, which meets the design requirements and performance indicators.

[0285] (5) Feature point 7 and feature point 9

[0286] Substituting the corresponding aerodynamic data at feature points 7 and 9 into the pitch channel transfer function yields the transfer function for each component. Substituting this transfer function into the simulation program and adjusting the parameters results in a set of parameters that provide better system performance. At feature points 7 and 9, the step response and Bode plot of the pitch channel are as follows: Figure 12 and Figure 13 As shown.

[0287] Simulation results show that the system has no steady-state error at feature points 7 and 9, the rise time to 90% of the steady-state value is 0.8s, the settling time under the 2% error band is 1.6s, the peak time is 1.3s, the system overshoot is 2.8%, the phase margin is 57.6°, the crossover frequency is 2.38rad / s, and the gain margin is 4.98dB, which meets the design requirements and performance indicators.

[0288] (6) Feature point 8

[0289] Substituting the corresponding aerodynamic data at the 8 feature points into the pitch channel transfer function yields the transfer function for each component. Substituting this transfer function into the simulation program and adjusting the parameters results in a set of parameters that provide better system performance. At feature point 8, the step response and Bode plot of the pitch channel are as follows: Figure 14 and Figure 15 As shown.

[0290] Simulation results show that the system has no steady-state error at feature point 8, the rise time to 90% of the steady-state value is 1.0s, the settling time under the 2% error band is 1.4s, the system has no overshoot, the phase margin is 65.7°, the crossover frequency is 1.94rad / s, and the gain margin is 5.15dB, which meets the design requirements and performance indicators.

[0291] All control parameters for the pitch channel are shown in Table 2:

[0292] Table 2 Pitch Channel Controller Parameter Values

[0293]

[0294] 2. Yaw Channel Simulation

[0295] The yaw channel control system designed in this invention is as follows: Figure 2 As shown, simulation debugging was performed at each feature point. First, the PID coefficients and corresponding parameters were coarsely adjusted. After obtaining indicators that met the design requirements, fine-tuning was performed to finally obtain the control system parameters that met the design requirements.

[0296] The limiting stage restricts the rudder deflection angle at... Taking feature point 1 and feature point 3 as examples, substituting the corresponding aerodynamic data at feature point 1 and feature point 3 into equation (68), the transfer functions of the corresponding links can be obtained as follows:

[0297] (76)

[0298] By substituting the transfer function into the simulation program and adjusting the parameters, a set of parameters with good system performance was obtained. Similarly, the parameters corresponding to each characteristic point were obtained. All control parameters for the yaw channel are shown in Table 3.

[0299] Table 3. Yaw Channel Controller Parameter Values

[0300]

[0301] 3. Rolling channel simulation

[0302] The rolling channel control system designed in this invention is as follows: Figure 3 As shown, simulation debugging was performed at each feature point. First, the PD coefficient and corresponding parameters were coarsely adjusted. After obtaining indicators that met the design requirements, fine-tuning was performed to finally obtain the control system parameters that met the design requirements.

[0303] The limiting stage restricts the rudder deflection angle at... Taking feature point 1 and feature point 3 as examples, substituting the corresponding aerodynamic data at feature point 1 and feature point 3 into equation (71), the transfer functions of the corresponding links can be obtained as follows:

[0304] (77)

[0305] By substituting the transfer function into the simulation program and adjusting the parameters, a set of parameters with good system performance was obtained. Similarly, the parameters corresponding to each feature point were obtained. All control parameters of the roll channel are shown in Table 4.

[0306] Table 4. Parameter Values ​​for Rolling Channel Controller

[0307]

[0308] Before compensation, the designed pitch, yaw, and roll three-channel control system was simulated on an actual model to obtain the response curves of the corresponding channels in the actual model. Taking feature point 1 as an example, the step response curves of the three channels are as follows: Figure 16 , Figure 17 and Figure 18 As shown.

[0309] Simulation results show that the control performance of the pitch, yaw, and roll channels decreases compared to the theoretical values. This is due to the coupling between the three channels and the influence of the propeller. The rise time of the pitch and yaw channels increases, the response speed slows down, the overshoot increases, the settling time lengthens, and the system performance deteriorates. The roll channel is significantly affected by the propeller, resulting in a substantial increase in steady-state error caused by the propeller's counter-torque and angular acceleration torque. To reduce or even eliminate the coupling between the three channels and the interference from the propeller, this invention introduces several compensation terms into the corresponding rudder deflection commands. By introducing compensation terms into the rudder deflection commands, corresponding aerodynamic forces and torques are generated, thereby improving the accuracy and performance of the control system.

[0310] Next, we will perform a simulation analysis of the compensated control system: Based on the aerodynamic parameters at the feature points, the corresponding compensation parameters are set as shown in Table 5.

[0311] Table 5 Compensation Item Parameter Value Table

[0312]

[0313] Using the method of this invention to compensate for the rudder deflection command, taking feature point 1 as an example, the step response diagrams before and after compensation are as follows: Figure 19 , Figure 20 and Figure 21 As shown in the simulation results, the step response overshoot of the pitch and yaw channels before compensation is large, the adjustment time is long, and the response speed is slow. After introducing the compensation term, the system step response overshoot is reduced, the response speed is faster, and the introduced compensation term improves the dynamic performance of the system. The steady-state error of the roll channel before compensation is large, while the system after compensation eliminates the steady-state error, effectively compensating for the propeller's influence on the airframe and meeting the design requirements of the control system.

Claims

1. A flight control method for fixed-wing unmanned aerial vehicles based on three-channel decoupling and dynamic compensation, characterized in that, The method specifically includes the following steps: Step 1: Based on the forces and torques acting on the fixed-wing UAV, establish the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV; Step 2: Based on the kinematic and dynamic model of the fixed-wing UAV and the overload equation of the fixed-wing UAV, obtain the control model of the fixed-wing UAV, and then design the three-channel control system of pitch, yaw and roll according to the control model of the fixed-wing UAV.

2. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 1, characterized in that, Based on the kinematic and dynamic model of the fixed-wing UAV and its overload equations, the control model of the fixed-wing UAV is obtained, specifically as follows: (57) in, Indicates angle of attack. express The first derivative, This represents the roll angular velocity in the projectile coordinate system; The yaw rate in the projectile coordinate system; The pitch angular velocity in the projectile coordinate system; Indicates the sideslip angle. for The first derivative, Indicates dynamic pressure. This indicates the magnitude of the thrust acting on a fixed-wing drone. This represents the reference area of ​​a fixed-wing UAV. Indicates the quality of the drone. This indicates the speed of a fixed-wing drone. This represents the derivative of the lift coefficient with respect to the angle of attack. This represents the derivative of the lift coefficient with respect to the elevator. Indicates the pitch deflection angle. This represents the derivative of the lateral force coefficient with respect to the sideslip angle. This represents the derivative of the lateral force coefficient with respect to the rudder. Indicates the yaw angle. This indicates the characteristic length of a fixed-wing unmanned aerial vehicle (UAV). This is the derivative of the rolling moment coefficient with respect to the sideslip angle. Represents the coordinate system of the projectile. moment of inertia of the shaft Let be the derivative of the rolling torque with respect to the dimensionless yaw rate. Let be the derivative of the yaw moment with respect to the dimensionless roll velocity. Let be the derivative of the rolling moment with respect to the yaw angle. This refers to the roll deflection angle of the rudder. Represents the coordinate system of the projectile. moment of inertia of the shaft Represents the coordinate system of the projectile. moment of inertia of the shaft express The first derivative, The dimensionless rolling damping moment coefficient, This is the control torque coefficient for the roll rudder. This is the derivative of the yaw moment coefficient with respect to the sideslip angle. The dimensionless yaw damping moment coefficient, This represents the derivative of the yaw moment with respect to the dimensionless pitch velocity. Let be the derivative of the yaw moment with respect to the roll angle. This is the control torque coefficient of the yaw rudder. The dimensionless pitch damping moment coefficient, Let be the derivative of the pitch moment coefficient with respect to the angle of attack. This is the pitch control torque coefficient. This represents the dimensionless drag coefficient. express The first derivative, express The first derivative, The derivative of the roll angle, Indicates the quality of the drone. It is the acceleration due to gravity; (58) in, All are aerodynamic parameters; Substituting the aerodynamic parameters in equation (58) into equation (57), we obtain the control model for the fixed-wing UAV: (59) in, and This is a normal overload in the ground coordinate system.

3. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 2, characterized in that, The design method for the pitch channel control system is as follows: Establish a control model for the pitch channel: (60) Applying a Laplace transform to equation (60) yields the transfer function of equation (61): (61) in, Represents the Laplace variable. The Laplace transform of the pitch angular velocity signal. The Laplace transform of the pitch deflection signal. The Laplace transform of the angle-of-attack signal. Represents the Laplace transform form of the normal overload signal; The rudder deflection command for the pitch channel is obtained according to equation (61). .

4. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 3, characterized in that, The pitch channel's rudder angle command for: (62) in, For pitch channel compensation, This refers to the propeller throttle opening. , and These are the PID control parameters for the pitch channel. , , , , and These are all design parameters for the pitch channel; Here is the transfer function of the rate gyroscope. The transfer function of the accelerometer. This indicates a normal overload command for the pitch channel.

5. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 4, characterized in that, The transfer function of the rate gyroscope is: (63) In the formula, The rate gyroscope damping ratio, is the undamped oscillation frequency of the rate gyroscope.

6. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 5, characterized in that, The transfer function of the accelerometer is: (65) In the formula, For the accelerometer damping ratio, is the undamped oscillation frequency of the accelerometer.

7. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 6, characterized in that, The design method for the yaw channel control system is as follows: Establish a control model for the yaw channel: (67) Applying a Laplace transform to equation (67) yields the transfer function of equation (68): (68) in, This is the Laplace transform of the yaw rate signal. This is the Laplace transform of the yaw angle signal. This is the Laplace transform of the sideslip angle signal. This is the Laplace transform form of the transverse overload signal; The rudder deflection angle command for the yaw channel is obtained according to equation (68). .

8. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 7, characterized in that, The yaw channel rudder deflection command for: (69) in, This is a compensation item for deviation from the designated route. , and These are all PID control parameters for the yaw channel. , , , , and These are all design parameters for the yaw channel. This is a lateral overload command for the yaw channel.

9. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 8, characterized in that, The design method for the roll channel control system is as follows: Establish a control model for the roll channel: (70) in, Indicates roll angle The first derivative; Performing a Laplace transform on equation (70) yields the transfer function of equation (71): (71) in, This is the Laplace transform of the roll angular velocity signal. This is the Laplace transform of the roll deflection angle signal. The Laplace transform of the roll angle signal; The rudder deflection command for the roll channel is obtained according to equation (71). .

10. The fixed-wing UAV flight control method based on three-channel decoupling and dynamic compensation according to claim 9, characterized in that, The rudder deflection angle command for the roll channel is: (72) in, For the roll-through compensation item, and These are all PD controller parameters. , , and These are all design parameters for the roll channel. This is the command roll angle for the roll channel.

Images