Attitude control method for quadrotor unmanned aerial vehicle based on active disturbance rejection fast terminal sliding mode

Abstract

The application discloses a quadrotor unmanned aerial vehicle attitude control method based on a self-disturbance rejection fast terminal sliding mode, constructs a cascade controller to control the attitude of the quadrotor unmanned aerial vehicle, the cascade controller is a two-stage cascade controller of inner and outer loops, the outer loop of the cascade controller performs self-disturbance rejection control, and outputs a desired attitude angle to the inner loop; the inner loop performs fast terminal sliding mode control, and outputs an attitude angle change amount to the aircraft; and the aircraft changes the flight attitude according to the attitude angle change amount. The quadrotor unmanned aerial vehicle attitude control method based on the self-disturbance rejection fast terminal sliding mode has strong anti-interference capability, faster tracking speed and robustness.

Description

TECHNICAL FIELD

[0001] The application relates to a four-rotor unmanned aerial vehicle attitude control method based on a self-disturbance rejection fast terminal sliding mode, and belongs to the technical field of aircraft control. BACKGROUND

[0002] Due to the under-actuated and strong-coupling characteristics of the four-rotor unmanned aerial vehicle system, the system is very susceptible to external environmental disturbances such as gusts and ground effects, and the self-disturbance rejection control method and the fast terminal sliding mode control method are usually used to solve the problem in the prior art.

[0003] However, these methods still have certain defects:

[0004] The self-disturbance rejection algorithm needs to convert the system into an integral series form, and unifies the external disturbance and internal uncertainty of the system as total disturbance, which often wastes effective information of the model and causes phase lag of the system;

[0005] The expression of the fast terminal sliding mode controller is too complex to be implemented, and if the initial value of the system state is not good, singular terms will appear in the expression of the controller, and the output of the controller tends to infinity.

[0006] Therefore, it is necessary to further study the four-rotor unmanned aerial vehicle attitude control method to improve the attitude control stability of the four-rotor unmanned aerial vehicle. SUMMARY

[0007] In order to overcome the above problems, the present application has been researched in depth, and a four-rotor unmanned aerial vehicle attitude control method based on a self-disturbance rejection fast terminal sliding mode is provided, which comprises:

[0008] A cascade controller is constructed to control the attitude of the four-rotor unmanned aerial vehicle, the cascade controller is a two-stage cascade controller of inner and outer loops, the outer loop of the cascade controller performs self-disturbance rejection control, outputs the expected attitude angle to the inner loop, the inner loop performs fast terminal sliding mode control, and outputs the attitude angle change amount to the aircraft, and the aircraft changes the flight attitude according to the attitude angle change amount.

[0009] Further, the input of the outer loop of the cascade controller is the expected angular velocity, and the output is the expected attitude angle, which comprises a nonlinear PD controller and an outer loop state observer,

[0010] The output of the nonlinear PD controller is the attitude angle control amount, the output of the outer loop state observer is the expected angular velocity and actual angular velocity error estimation and attitude angle disturbance compensation amount, and the attitude angle control amount u0 and the attitude angle disturbance compensation amount are combined to obtain the expected attitude angle.

[0011] The input of the nonlinear PD controller is the desired angular velocity, and the desired angular velocity and actual angular velocity error estimation output by the outer loop state observer, the input of which is the attitude angle of the UAV and the attitude angle control amount output by the nonlinear PD controller.

[0012] In one embodiment, the outer loop state observer can be expressed as:

[0013]

[0014]

[0015] wherein subscript i represents different angular velocities, p represents the roll angular velocity, q represents the pitch angular velocity, and r represents the yaw angular velocity; x i1 , x i2 , x i3 , y i are state quantities, wherein is the observation value of x i1 , is the observation value of x i2 , is the observation value of x i3 , is the observation value of the state quantity output by the outer loop extended state observer, β i1 , β i2 , β i3 are design parameters; b i is a compensation factor; f i is the outer loop control disturbance, expressed as:

[0016]

[0017] wherein subscript j represents different attitude angles, φ represents the roll angle, θ represents the pitch angle, and ψ represents the yaw angle.

[0018] Preferably, the nonlinear PD controller can be expressed as:

[0019]

[0020] wherein i represents different angular velocities, i d represents the desired of the corresponding angular velocity, represents the observation value of the corresponding angular velocity, represents the observation value of the corresponding angular velocity; e i1 represents the error of the desired angular velocity and the angular velocity observation value; u i0 represents the control amount of different attitude angles, β i01 , β i02 represents the gain of the nonlinear PD controller; fal is a nonlinear filtering function, αi1 ,α i2 ,δ i are design parameters.

[0021] Preferably, the attitude angle disturbance compensation output by the outer loop state observer is:

[0022]

[0023] The outer loop state observer is simplified as:

[0024]

[0025]

[0026] wherein b i0 is a compensation factor.

[0027] In one embodiment, the input of the inner loop of the cascade controller is the desired attitude angle, and the output is the attitude angle change amount, and the aircraft changes the flight attitude according to the attitude angle change amount.

[0028] The inner loop of the cascade controller comprises a fast non-singular terminal sliding mode controller and an inner loop state observer,

[0029] wherein the output of the fast non-singular terminal sliding mode controller is the attitude angle change amount, and the output of the inner loop state observer is the error estimation of the desired attitude angle and the actual attitude angle.

[0030] The input of the fast non-singular terminal sliding mode controller is the desired attitude angle, and the error estimation of the desired attitude angle and the actual attitude angle output by the inner loop state observer, and the input of the inner loop state observer is the desired attitude angle and the actual attitude angle of the aircraft.

[0031] Preferably, the inner loop state observer can be expressed as:

[0032]

[0033]

[0034] wherein subscript j represents different attitude angles, φ represents the roll angle, θ represents the pitch angle, and ψ represents the yaw angle.

[0035] x j1 , x j2 , x j3 , y j are state quantities, is the observation value of x j1 , is the observation value of x j2 , is the observation value of xj3 the observation value of the state quantity output by the inner loop state observer, the observation value of the state quantity output by the inner loop state observer, β j1 , β j2 , β j3 is a design parameter; b j is a compensation factor; f j is the inner loop control disturbance, which is expressed as:

[0036]

[0037] In a preferred embodiment, the sliding surface s j of the fast non-singular terminal sliding mode controller is set as:

[0038]

[0039] e j = j d -j

[0040] wherein α j , β j are constants, m0, n0 are positive odd numbers, and m0 > n0.

[0041] In a preferred embodiment, the reaching law of the fast non-singular terminal sliding mode controller is set as:

[0042]

[0043] wherein, γ j is a constant, m, n are positive odd numbers, and m > n.

[0044] In a preferred embodiment, the fast non-singular terminal sliding mode controller can be expressed as:

[0045]

[0046]

[0047] wherein the subscript j represents different attitude angles, φ represents the roll angle, θ represents the pitch angle, and ψ represents the yaw angle, u j = τ j represents the attitude angle change amount, h j , k j are constant parameters, is the observation value of f j .

[0048] The present application has the following beneficial effects:

[0049] ​(1) The outer loop active disturbance rejection controller is used to improve the convergence speed of the system state, and to improve the chattering problem in the sliding mode control, reduce the adjustable parameters, and be easy for engineering application;

[0050] (2) The inner loop terminal sliding mode control is used to effectively utilize the known model of the system, and to avoid the waste of model information. BRIEF DESCRIPTION OF DRAWINGS

[0051] Figure 1 A schematic diagram of a cascade controller structure in a quadrotor unmanned aerial vehicle attitude control method based on active disturbance rejection fast terminal sliding mode according to a preferred embodiment of the present application is shown;

[0052] Figure 2 A fal function schematic diagram in a quadrotor unmanned aerial vehicle attitude control method based on active disturbance rejection fast terminal sliding mode according to a preferred embodiment of the present application is shown;

[0053] Figure 3 A roll angle channel response curve diagram in experimental example 1 is shown;

[0054] Figure 4 A pitch angle channel response curve diagram in experimental example 1 is shown;

[0055] Figure 5 A yaw angle channel response curve diagram in experimental example 1 is shown;

[0056] Figure 6 A roll angle channel error response diagram in experimental example 1 is shown;

[0057] Figure 7 A pitch angle channel error response diagram in experimental example 1 is shown;

[0058] Figure 8 A yaw angle channel error response diagram in experimental example 1 is shown. DETAILED DESCRIPTION

[0059] The present application is further described in detail by the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present application will become more apparent.

[0060] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Unless specifically stated otherwise, the drawings are not drawn to scale and are shown for purposes of explanation only.

[0061] The application provides a quadrotor unmanned aerial vehicle attitude control method based on a self-disturbance rejection fast terminal sliding mode. Figure 1

[0062] In the application, the self-disturbance rejection control and the fast terminal sliding mode control are effectively combined by constructing a cascade controller, so that the disadvantages caused by single use of the two methods are solved, and stronger interference and robustness are achieved while fast response tracking is realized.

[0063] Further, the input of the outer loop of the cascade controller is a desired angular velocity, and the output is a desired attitude angle, and the outer loop comprises a nonlinear PD controller and an outer loop state observer,

[0064] The output of the nonlinear PD controller is an attitude angle control amount u0, the output of the outer loop state observer is a desired angular velocity and actual angular velocity error estimation and an attitude angle disturbance compensation amount, and the desired attitude angle is obtained by combining the attitude angle control amount u0 and the attitude angle disturbance compensation amount.

[0065] The input of the nonlinear PD controller is a desired angular velocity, and the desired angular velocity and actual angular velocity error estimation output by the outer loop state observer, and the input of the outer loop state observer is the attitude angle of the unmanned aerial vehicle and the attitude angle control amount u0 output by the nonlinear PD controller.

[0066] In a preferred embodiment, the outer loop state observer can be represented as:

[0067]

[0068]

[0069] Wherein, the subscript i represents different angular velocities, p represents a roll angular velocity, q represents a pitch angular velocity, and r represents a yaw angular velocity. i1 , x i2 , x i3 , y i are state variables, wherein is the observation value of x i1 , is the observation value of x i2 , is the observation value of x i3 , is the observation value of the state variable output by the outer loop state observer, and β i1 ​, β i2 , β i3 is a design parameter; b i is a compensation factor; f i is an outer loop control disturbance, expressed as:

[0070]

[0071] wherein subscript j represents different attitude angles, φ represents a roll angle, θ represents a pitch angle, and ψ represents a yaw angle.

[0072] In this embodiment, the system internal uncertainty and external disturbance are unified as a total disturbance f i of the system, which is taken as a state variable of the system, and then the influence of the disturbance is reduced through an outer loop state observer.

[0073] In a preferred embodiment, the parameter β i1 , β i2 , β i3 is obtained by using a parameter bandwidth method, which is a commonly used parameter determination method in self-disturbance control and is not described herein.

[0074] Preferably, the parameters β i1 , β i2 , β i3 are selected as:

[0075]

[0076] ω0 is the bandwidth of the outer loop state observer.

[0077] The nonlinear PD controller is actually an error feedback of the system, and an error estimate of the desired angular velocity and the actual angular velocity is obtained by the desired angular velocity and the outer loop state observer, and is combined nonlinearly.

[0078] In a preferred embodiment, the nonlinear PD controller can be expressed as:

[0079]

[0080] wherein i represents different angular velocities, i d represents a desired value of the corresponding angular velocity, represents an observed value of the corresponding angular velocity, represents an observed value of the corresponding angular velocity; e i1 represents an error of the desired angular velocity and the observed value of the corresponding angular velocity, and represents an error of the desired angular velocity and the observed value of the corresponding angular velocity; u i0 represents control amounts of different attitude angles, β i01 , β i02represents a nonlinear PD controller gain; fal is a nonlinear filter function, which is any function that can realize small gain when large error and large gain when small error, and α i1 ,α i2 ,δ i are design parameters.

[0081] In a preferred embodiment, the nonlinear filter function fal is:

[0082]

[0083] wherein α, δ are parameters.

[0084] A schematic diagram of the function is shown in Fig. 2, which has a good convergence characteristic and can improve the anti-interference of the system. Figure 2

[0085] In a preferred embodiment, the attitude angle disturbance compensation quantity output by the outer loop state observer is:

[0086]

[0087] The outer loop state observer is simplified as:

[0088]

[0089]

[0090] wherein b i0 is a compensation factor, the value of which is related to b i .

[0091] The setting of the attitude angle disturbance compensation quantity makes the outer loop calculation process become a calculation of a series type pure integral element, greatly simplifying the calculation amount and improving the response speed.

[0092] The traditional fast terminal sliding mode control has poor robustness, and in the present application, a state observer is additionally added on the basis of the fast terminal sliding mode control to improve the stability.

[0093] Further, the input of the inner loop of the cascade controller is the desired attitude angle, and the output is the attitude angle change amount, and the aircraft changes the flight attitude according to the attitude angle change amount.

[0094] The inner loop of the cascade controller comprises a fast non-singular terminal sliding mode controller and an inner loop state observer,

[0095] wherein the output of the fast non-singular terminal sliding mode controller is the attitude angle change amount, and the output of the inner loop state observer is the error estimation of the desired attitude angle and the actual attitude angle. ​

[0096] The input of the fast non-singular terminal sliding mode controller is the desired attitude angle, and the estimation error of the desired attitude angle and the actual attitude angle of the output of the inner loop state observer, the input of the inner loop state observer is the desired attitude angle and the actual attitude angle of the aircraft.

[0097] In a preferred embodiment, the inner loop state observer can be expressed as:

[0098]

[0099]

[0100] Wherein, the subscript j represents different attitude angles, φ represents the roll angle, θ represents the pitch angle, and ψ represents the yaw angle.

[0101] x j1 , x j2 , x j3 , y j is a state variable, is the observed value of the state variable of the output of the inner loop state observer, wherein is the observed value of x j1 , is the observed value of x j2 , is the observed value of x j3 , β j1 , β j2 , β j3 is a design parameter; b j is a compensation factor; f j is the inner loop control disturbance, expressed as:

[0102]

[0103] The specific values of the design parameters β j1 , β j2 , β j3 can be obtained by experience or multiple tests by those skilled in the art, which are not limited in the present application.

[0104] In a preferred embodiment, the sliding surface s j of the fast non-singular terminal sliding mode controller is set as:

[0105]

[0106] e j = j d -j

[0107] Wherein, α j , β jare constants, m0, n0 are positive odd numbers, and m0 > n0.

[0108] In a preferred embodiment, the reaching law of the fast non-singular terminal sliding mode controller can be expressed as: is set as:

[0109]

[0110] wherein, γ j are constants, m, n are positive odd numbers, and m > n.

[0111] In a preferred embodiment, the fast non-singular terminal sliding mode controller can be expressed as:

[0112]

[0113]

[0114] wherein, subscript j represents different attitude angles, φ represents a roll angle, θ represents a pitch angle, ψ represents a yaw angle, u j = τ j represents an attitude angle change amount, h j , k j are constant parameters, is an observation value of f j .

[0115] Embodiment

[0116] Embodiment 1

[0117] The simulation software is used to simulate the attitude control of the quadrotor unmanned aerial vehicle, and the parameter values of the quadrotor unmanned aerial vehicle are as follows:

[0118] T m = 0.0914, C R = 881.44,

[0119] d = 0.1875, c T = 0.938 * 10 -5 , c M = 1.479 * 10 -7 ,

[0120] m = 1.5, J = [0.021029 0.021633 0.039424],

[0121] J RP = 0.0000819, C d = 0.0746

[0122] wherein, T mrepresents the transition time required for the motor to reach the steady state speed from 0, C R represents the product of the battery capacity and the total voltage, represents the steady state speed of the motor, d represents the distance from the center of the body to the center of the propeller, c T represents the propeller drag coefficient, c M represents the propeller torque coefficient, m represents the mass of the unmanned aerial vehicle, J represents the three-axis rotational inertia of the unmanned aerial vehicle, J RP represents the total rotational inertia of the entire motor and propeller, C d represents the average drag coefficient of the propeller blade section airfoil.

[0123] Because in actual situation, the size and direction of the gust received by the quadrotor unmanned aerial vehicle during flight cannot be determined, in order to be more in line with the actual situation, the disturbance added in the simulation is set as a random white noise and a sine function of the simulation time, and the expression is as follows:

[0124] d = 3sin(t) + WGN

[0125] Wherein, WGN represents random white noise.

[0126] A cascade controller is constructed to control the attitude of the quadrotor unmanned aerial vehicle, the cascade controller is a two-stage cascade controller of inner and outer loops, the outer loop of the cascade controller performs active disturbance rejection control, and outputs the desired attitude angle to the inner loop, the inner loop performs fast terminal sliding mode control, and outputs the attitude angle change to the aircraft, and the aircraft changes the flight attitude according to the attitude angle change.

[0127] The input of the outer loop of the cascade controller is the desired angular velocity, and the output is the desired attitude angle, which includes a nonlinear PD controller and an outer loop state observer,

[0128] Wherein, the output of the nonlinear PD controller is the attitude angle control amount u0, and the output of the outer loop state observer is the error estimation of the desired angular velocity and the actual angular velocity, the attitude angle disturbance compensation amount, the attitude angle control amount u0 and the attitude angle disturbance compensation amount are combined to obtain the desired attitude angle;

[0129] The input of the nonlinear PD controller is the desired angular velocity, and the error estimation of the desired angular velocity and the actual angular velocity output by the outer loop state observer, and the input of the outer loop state observer is the attitude angle of the unmanned aerial vehicle and the attitude angle control amount u0 output by the nonlinear PD controller.

[0130] The outer loop state observer can be represented as:

[0131]

[0132]

[0133] wherein,

[0134] The nonlinear PD controller can be expressed as:

[0135]

[0136] The nonlinear filter function fal is:

[0137]

[0138] The attitude angle disturbance compensation output by the outer loop state observer is:

[0139]

[0140] The outer loop state observer is simplified as:

[0141]

[0142]

[0143] The input of the inner loop of the cascade controller is the desired attitude angle, and the output is the attitude angle change amount, and the aircraft changes the flight attitude according to the attitude angle change amount;

[0144] The inner loop of the cascade controller includes a fast non-singular terminal sliding mode controller and an inner loop state observer,

[0145] The output of the fast non-singular terminal sliding mode controller is the attitude angle change amount, and the output of the inner loop state observer is the error estimation of the desired attitude angle and the actual attitude angle;

[0146] The input of the fast non-singular terminal sliding mode controller is the desired attitude angle, and the estimation error of the desired attitude angle and the actual attitude angle of the output of the inner loop state observer, and the input of the inner loop state observer is the desired attitude angle and the actual attitude angle of the aircraft.

[0147] The inner loop state observer can be expressed as:

[0148]

[0149]

[0150]

[0151] The fast non-singular terminal sliding mode controller can be expressed as:

[0152]

[0153]

[0154] wherein the relevant parameters of the outer loop are set as:

[0155] β p1 ,β p2 = [65; 55]; a p1 , a p2 = [0.13; 0.13];

[0156] δ p = 0.02; b p = 11;

[0157] β q1 ,β q2 = [65; 55]; a q1 , a q2 = [0.13; 0.13];

[0158] δ q = 0.02; b q = 5;

[0159] β r1 ,β r2 = [65; 80]; a r1 , a r2 = [0.13; 0.13];

[0160] δ r = 0.02; b r = 8; w0= 10

[0161] wherein the relevant parameters of the inner loop are set as:

[0162] b φ = 11; a φ = 2;

[0163] β φ = 9; h φ = 10; k φ = 10;

[0164] b θ = 11; a θ = 2;

[0165] β θ = 9; h θ = 10; k θ = 10;

[0166] b ψ = 11; a ψ = 2;

[0167] β ψ = 9; h ψ = 10; k ψ = 10;

[0168] m0= 9; n0= 5; m = 3; n = 1;

[0169] Comparative Example 1

[0170] The attitude of the quadrotor unmanned aerial vehicle was simulated using simulation software, the parameter values of the quadrotor unmanned aerial vehicle were the same as those of Example 1, and the control mode was the mode introduced in the paper K.J. Astrom and A.B. Ostberg, "A Teaching Laboratory for Process Control," 1985 American Control Conference, 1985, pp. 1380-1385, doi: 10.23919 / ACC.1985.4788834.

[0171] Experimental Example 1

[0172] The simulation results of Example 1 and Comparative Example 1 are shown in Figures 3-8 .

[0173] Among them, Figure 3 shows the roll angle channel response curve; Figure 4 shows the pitch angle channel response curve; Figure 5 shows the yaw angle channel response curve; Figure 6 shows the roll angle channel error response graph; Figure 7 shows the yaw angle channel error response graph; Figure 8 shows the yaw angle channel error response graph.

[0174] Table 1

[0175] IAE Example 1 Comparative Example 1 Roll angle 0.0086 0.0124 Pitch angle 0.0184 0.0278 Yaw angle 0.0082 0.0122

[0176] Table 1 shows the integral of the absolute error amplitude (IAE) on the three channels in Example 1 and Comparative Example 1, from Figures 3-8 It can be seen from Table 1 and Table 1 that Example 1 has better anti-interference ability, has faster tracking speed and robustness in response to the sine signal, and Example 1 has good decoupling performance, which is due to the extended state observer that can estimate and compensate "total disturbance" in real time.

[0177] In the description of the present application, it should be noted that the terms "upper", "lower", "inner", "outer", "front", "back" and the like indicate the positional or location relationship based on the working state of the present application, and are only for the purpose of facilitating the description of the present application and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application. In addition, the terms "first", "second", "third", "fourth" are only for the purpose of description, and cannot be understood as indicating or implying relative importance.

[0178] In the description of the present application, it should be noted that unless otherwise explicitly specified and limited, the terms "mounting", "connecting", "connecting" should be understood broadly, for example, it can be fixed connection, or detachable connection, or integral connection; it can be mechanical connection, or electrical connection; it can be direct connection, or indirect connection through intermediate medium; it can be the communication inside two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0179] The above describes the present application in combination with the preferred embodiments, but these embodiments are only exemplary and serve only to illustrate. On this basis, various substitutions and improvements can be made to the present application, which all fall within the scope of protection of the present application.

Claims

1. A quadrotor unmanned aerial vehicle (UAV) attitude control method based on a fast terminal sliding mode with active disturbance rejection, characterized in that, a cascade controller is constructed to control the attitude of the quadrotor UAV, the cascade controller is a two-stage cascade controller with an inner loop and an outer loop, the outer loop of the cascade controller performs active disturbance rejection control and outputs a desired attitude angle to the inner loop, the inner loop performs fast terminal sliding mode control and outputs an attitude angle change to the aircraft, and the aircraft changes the flight attitude according to the attitude angle change; the input of the outer loop of the cascade controller is a desired angular velocity, and the output is a desired attitude angle, and the outer loop includes a nonlinear PD controller and an outer loop state observer, wherein the output of the nonlinear PD controller is an attitude angle control quantity, the output of the outer loop state observer is a desired angular velocity and actual angular velocity error estimation, and an attitude angle disturbance compensation quantity, and the attitude angle control quantity u0 and the attitude angle disturbance compensation quantity are combined to obtain the desired attitude angle; the input of the nonlinear PD controller is the desired angular velocity, and the desired angular velocity and actual angular velocity error estimation output by the outer loop state observer, and the input of the outer loop state observer is the attitude angle of the UAV and the attitude angle control quantity output by the nonlinear PD controller; the input of the inner loop of the cascade controller is the desired attitude angle, and the output is the attitude angle change, and the aircraft changes the flight attitude according to the attitude angle change; the inner loop of the cascade controller includes a fast non-singular terminal sliding mode controller and an inner loop state observer, wherein the output of the fast non-singular terminal sliding mode controller is the attitude angle change, and the output of the inner loop state observer is the error estimation of the desired attitude angle and the actual attitude angle; the input of the fast non-singular terminal sliding mode controller is the desired attitude angle, and the estimation error of the desired attitude angle and the actual attitude angle output by the inner loop state observer, and the input of the inner loop state observer is the desired attitude angle and the actual attitude angle of the aircraft. 2.The quadrotor UAV attitude control method based on a fast terminal sliding mode with active disturbance rejection according to claim 1, characterized in that, the outer loop state observer is expressed as: i e [p, q, r] i e [p, q, r] wherein subscript i represents different angular velocities, p represents a roll angular velocity, q represents a pitch angular velocity, and r represents a yaw angular velocity; x i1 , x i2 , x i3 , y i are state quantities, wherein is an observation value of x i1 , is an observation value of x i2 , is an observation value of x i3 , is a state quantity observation value output by an outer loop extended state observer, β i1 , β i2 , β i3 are design parameters; b i is a compensation factor; f i is an outer loop control disturbance, expressed as: wherein subscript j represents different attitude angles, φ represents a roll angle, θ represents a pitch angle, and ψ represents a yaw angle. 3.The quadrotor UAV attitude control method based on a fast terminal sliding mode with active disturbance rejection according to claim 2, characterized in that, the nonlinear PD controller is expressed as: i = p, q, r where i denotes different angular velocities, i d denotes the desired value of the corresponding angular velocity, denotes the observed value of the corresponding angular velocity, denotes the observed value of the corresponding angular velocity; e i1 denotes the error of the desired angular velocity and the angular velocity observation; u i0 denotes the control quantity of different attitude angles, β i01 , β i02 denotes the gain of the nonlinear PD controller; fal is a nonlinear filtering function, α i1 , α i2 , δ i is a design parameter. 4.The quadrotor UAV attitude control method based on a fast terminal sliding mode with active disturbance rejection according to claim 3, characterized in that, the attitude angle disturbance compensation quantity output by the outer loop state observer is: i e [p, q, r] then the outer loop state observer is simplified as: i e [p, q, r] i e [p, q, r] where b i0 is a compensation factor. 5.The quadrotor UAV attitude control method based on a fast terminal sliding mode with active disturbance rejection according to claim 1, characterized in that, the inner loop state observer is expressed as: j e [φ, θ, ψ] wherein subscript j represents different attitude angles, φ represents a roll angle, θ represents a pitch angle, and ψ represents a yaw angle. x j1 , x j2 , x j3 , y j is a state variable, is an observation value of x j1 , is an observation value of x j2 , is an observation value of x j3 , is a state variable observation value output by the inner loop state observer, β j1 , β j2 , β j3 is a design parameter; b j is a compensation factor; f j is an inner loop control disturbance, expressed as: j e [φ, θ, ψ]. 6.The quadrotor UAV attitude control method based on a fast terminal sliding mode with active disturbance rejection according to claim 1, characterized in that, The sliding surface s of the fast non-singular terminal sliding mode controller j is set to: j e [φ, θ, ψ] e j = j d - j wherein a j , β j are constants, m0, n0 are positive odd numbers, and m0>n0.

7. The active disturbance rejection based fast terminal sliding mode attitude control method for quadrotor UAV according to claim 6, wherein The reaching law of the fast non-singular terminal sliding mode controller is set to: j e [φ, θ, ψ] wherein γ j is a constant, m, n are positive odd numbers, and m > n.

8. The active disturbance rejection based fast terminal sliding mode attitude control method for quadrotor UAV according to claim 7, wherein The fast non-singular terminal sliding mode controller is expressed as: j e [φ, θ, ψ] wherein subscript j represents different attitude angles, φ represents a roll angle, θ represents a pitch angle, and ψ represents a yaw angle, u j = τ j represents a change in attitude angle, h j , k j are constant parameters, is an observation value of f j .

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