Attitude control method for spin detection satellite

Abstract

An attitude control method for a spin detection satellite, which method belongs to the field of satellite attitude control. The method comprises: S10, establishing a satellite dynamic equation and converting same into a state-space equation; S20, designing a two-layer linear extended state observer, wherein the first-layer linear state observer is used for observing and acquiring a current angular velocity estimate and a current disturbance estimate of a satellite, and the second-layer state observer observes a tracking error of the first-layer linear state observer on the basis of the observation result of the first-layer linear state observer; and S30, then performing calculation on the basis of the current disturbance estimate and the tracking error thereof, so as to obtain the total internal and external disturbance torque estimate of the satellite, using attitude parameters representing the direction of a spin axis to design a PD control law, and compensating for the total internal and external disturbance torque estimate of the satellite, so as to calculate a control torque, which is used for satellite attitude control within a current cycle. The method has the advantages of fewer tuning parameters, ease of adjustment and stable performance, thereby greatly improving the observation precision.

Description

Attitude Control Method of Spin Detection Satellite Technical Field

[0001] The invention relates to a method for controlling the attitude of a spin detection satellite, and belongs to the technical field of satellite attitude control. Background Art

[0002] Due to the needs of scientific exploration missions, some exploration satellites use spin-stabilized attitude control. In satellites that explore the Earth's outer radiation belt environment, the normal operation of the payload requires the satellite to provide a high spin angular velocity, and in order to remove the satellite's residual magnetic interference, a flexible extension rod is required to extend the payload 1 meter away from the satellite. Considering that satellites need to repeatedly cross the outer radiation belt to detect the external magnetic field environment and the requirements for data transmission to the ground, this type of exploration satellite needs to have the ability to repeatedly and reliably spin up and despin and be able to maintain the pointing accuracy of the spin axis during the rotation process. The existing method uses angular momentum exchange and transposition to achieve whole-satellite spin-up and despinning and attitude control. This spin-up method has many advantages, such as saving fuel, improving control accuracy, and flexible and reliable operation.

[0003] However, during a satellite's high-speed spin, some high-precision sensors onboard, such as the star sensor, fail, making it difficult to simply and accurately obtain complete three-axis attitude data. Only spin-axis orientation information can be obtained through sensors such as the sun sensor. Because each set of quaternions corresponds to a specific attitude, traditional methods of controlling satellite spin using quaternions require designing and optimizing the desired quaternion trajectory, placing high demands on onboard computing power.

[0004] Extending the payload 1 meter from the satellite via a flexible extension rod increases the satellite's moment of inertia. During the satellite's high-speed spin, the cross term ω×Jω related to the satellite's moment of inertia and the vibration of the flexible extension rod significantly interfere with the satellite's attitude control, reducing attitude control accuracy. (Where ω is the angular velocity and J is the actual moment of inertia, a state observer, part of active disturbance rejection technology, can be used to estimate the disturbance torques from both inside and outside the satellite, further improving satellite control accuracy.) Traditional nonlinear extended state observers offer good estimation results but are difficult to adjust, while linearized state observers are easier to adjust but suffer from reduced estimation accuracy. Summary of the Invention

[0005] Aiming at the problems that it is difficult to obtain complete attitude parameters during the high-speed spinning of a satellite and the existing interference torque affects the satellite attitude control accuracy, the present invention provides an attitude control method for a spinning detection satellite.

[0006] A method for attitude control of a spin detection satellite according to the present invention includes: establishing a satellite dynamics equation based on rigid body dynamics and converting it into a state space equation; designing a two-layer linear cascade state observer based on the state space equation, wherein the first-layer linear state observer is used to observe and obtain the current angular velocity estimate and the current interference amount estimate of the satellite; the second-layer linear state observer observes the tracking error of the first-layer linear state observer based on the observation results of the first-layer linear state observer; the tracking error is the difference between the current angular velocity estimate and the current interference amount estimate and the current true angular velocity and the current true interference amount; then calculating the total interference torque estimate inside and outside the satellite from the current interference amount estimate and the tracking error; designing a PD control law using an attitude parameter w representing the direction of the spin axis, and compensating the total interference torque estimate inside and outside the satellite to calculate a control torque for satellite attitude control in the current cycle; then inputting the calculated control torque into the linear cascade state observer to observe the interference amount of the next cycle and calculate the control torque of the next cycle to achieve satellite attitude control.

[0007] According to the attitude control method of the spin detection satellite of the present invention, the satellite dynamic equation is: Where J is the actual moment of inertia, T d is the disturbance torque, T c is the control torque, ω is the angular velocity, ω × is the angular velocity cross product matrix, h is the flywheel angular momentum, According to J=J0+ΔJ, where J0 is the measured value of the moment of inertia and ΔJ is the uncertainty of the moment of inertia, the dynamic equation can be rewritten as: Where d is the total interference term: Convert the satellite dynamics equations into state-space equations: Where x1 is the state variable 1, x1 = ω; x2 is the state variable 2, x2 = d * , d * is the disturbance, f is the function term of the state variable x1, b0 is the control torque coefficient;

[0008] According to the attitude control method of the spin detection satellite of the present invention, the first-layer linear state observer is: Where z 1，1 is the current angular velocity estimate, z 1，2 is the estimated value of the current disturbance, l1 is the linear coefficient 1 of the first-layer linear state observer, e1 is the difference between the angular velocity measurement value and the angular velocity estimate, and l2 is the linear coefficient 2 of the first-layer linear state observer.

[0009] According to the attitude control method of the spin detection satellite of the present invention, the second-layer linear state observer is: Where z 2,1 is the angular velocity tracking error, z 2,2 is the disturbance tracking error, l3 is the linear coefficient 1 of the second-layer linear state observer, e2 is the difference between the angular velocity measurement value and the angular velocity tracking error, l4 is the linear coefficient 2 of the second-layer linear state observer; e1 = y--z 1,1 , e2=yz 2,1 , where y is the angular velocity measurement value.

[0010] According to the attitude control method of the spin detection satellite of the present invention, the estimated value of the total interference torque inside and outside the satellite is expressed as

[0011] According to the attitude control method of the spin detection satellite of the present invention, the PD control law is: Where D is the differential term coefficient, ω e is the angular velocity error, K is the proportional coefficient, w e The error of the attitude parameter w in the direction of the spin axis is represented by the complex number w e Expressed as: w e =w e1 +1w e2 , the matrix is ​​expressed as: w e =[w e1 , w e2 ,0] T , D=k d J0, K=k p J0, where w e1 w e The real part of w e2 w e The imaginary part, k d is the differential control law coefficient, k p is the proportional control law coefficient.

[0012] According to the attitude control method of the spin detection satellite of the present invention, w e Convert to Euler angles in the order of 3-2-1: Where φ is the roll angle and θ is the pitch angle.

[0013] According to the attitude control method of the spin detection satellite of the present invention, the differential control law coefficient k d and proportional control law coefficient k p The method for determining is: k d =2ξω c , where ξ is the system damping ratio, ω c is the undamped oscillation angular frequency of the system.

[0014] According to the attitude control method of the spin detection satellite of the present invention, the calculation method of the four linear coefficients of the linear cascade state observer is: l3=l1=2ω o , Where ω o is the angular frequency at which the observer's observation results converge; tracking convergence time t o for:

[0015] Beneficial effects of the present invention: The method of the present invention uses the w parameter alone to describe and control the movement of the spin axis pointing deflection, without the need to obtain complete attitude parameters, reducing the requirements for sensors, and meeting the needs of reliable repeated spin-up and despinning and high-precision control of high-speed spinning processes.

[0016] The method of the present invention uses a two-layer linear cascade state observer to estimate internal and external disturbances. The internal and external disturbances include complex interference torque and measurement errors. It can improve the attitude control accuracy, and the parameter adjustment is simple, and the potential for engineering practice application is great.

[0017] The method of the present invention uses (w, z) parameters to describe the attitude of a spinning satellite, solving the problem that the satellite cannot obtain complete three-axis information under high-speed spinning conditions. At the same time, only the w parameter is used to describe and control the direction of the spin axis. Compared with the traditional quaternion method, the calculation process is simplified and the amount of calculation is reduced.

[0018] The linear cascade extended state observer designed by the method of the present invention has the advantages of fewer setting parameters, easy adjustment and stable performance compared with the traditional nonlinear extended state observer; compared with the traditional linear extended state observer, the observation accuracy is greatly improved. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 is a flow chart of the linear cascade state observer in the attitude control method of the spin detection satellite described in the present invention; Figure 2 is a schematic diagram of the definition of the w parameter; in the figure, i′1i′2i′3 represent the inertial coordinate system, b3 represents the Z axis of the body coordinate system, and u represents the vector direction of the rotation of the coordinate axis b3 to the coordinate axis i′3, that is, the direction of the rotation axis; Figure 3 is a schematic diagram of the tracking error between the estimated value of the total internal and external interference torque of the satellite estimated by the linear cascade state observer and the actual value; Figure 4 is a schematic diagram of the control result of attitude control using the PD control law designed with the w parameter; Figure 5 is a schematic diagram of the tracking error between the estimated value of the total internal and external interference torque of the satellite estimated by the nonlinear state observer in the second specific implementation method and the actual value. DETAILED DESCRIPTION

[0020] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making any creative efforts shall fall within the scope of protection of the present invention.

[0021] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments may be combined with each other.

[0022] The present invention will be further described below with reference to the accompanying drawings and specific embodiments, but they are not intended to limit the present invention.

[0023] Specific embodiment 1, as shown in conjunction with Figures 1 and 2, the present invention provides an attitude control method for a spin detection satellite, including: establishing a satellite dynamics equation based on rigid body dynamics and converting it into a state-space equation; designing a two-layer linear cascade state observer based on the state-space equation, wherein the first-layer linear state observer is used to observe and obtain the satellite's current angular velocity estimate and current interference estimate; the second-layer linear state observer observes the tracking error of the first-layer linear state observer based on the observation results of the first-layer linear state observer; the tracking error is the difference between the current angular velocity estimate and the current interference estimate and the current true angular velocity and the current true interference; then, an estimate of the total internal and external interference torque of the satellite is calculated from the current interference estimate and the tracking error; using the attitude parameter w representing the direction of the spin axis to design a PD control law, and compensating the estimated total internal and external interference torque of the satellite to calculate a control torque for satellite attitude control in the current cycle; then, the calculated control torque is input into the linear cascade state observer to observe the interference in the next cycle and calculate the control torque in the next cycle to achieve satellite attitude control.

[0024] In this embodiment, angular momentum exchange and transposition are used to realize the rotation and de-spinning and attitude control of the entire satellite, which has many advantages such as saving fuel, improving control accuracy, and flexible and reliable operation.

[0025] Considering that the payload detection only requires the control accuracy of the satellite's spin axis pointing, the (w, z) parameter can be used to decouple the motion of the two directions of rotation around the spin axis and deflection of the spin axis pointing. Where w is the attitude parameter that characterizes the direction of the spin axis pointing, and z is the attitude parameter that characterizes the angle of rotation around the spin axis direction. If the rotation matrix C from the reference coordinate system to the body coordinate system is defined br Middle C br (1, 3) = a, C br (2, 3) = b, C br (3, 3) = c, then further define Where C br(1,3), C br (2, 3) and C br (3, 3) represents the three components of the third column of the rotation matrix from the reference coordinate system to the body coordinate system. w1 represents the real part of w, and w2 represents the imaginary part of w. The complex variable w represents the deflection of the spin axis during rotation. Therefore, the current satellite w parameter can be obtained using a sun sensor and used in the design of the control law to ensure the accuracy of the spin axis pointing.

[0026] This embodiment improves the overall estimation accuracy by cascading linear extended state observers, and sets parameters for each layer of linear state observers through convergence time.

[0027] The basic principle of a linear cascade state observer is that the lower-level observer observes the error in the tracking signal of the upper-level observer, improving overall observation accuracy by re-tracking the error. Its form can be designed based on the state space form of the system model. As shown in Figure 1, the number of layers in the linear cascade state observer can be designed based on actual needs. This implementation uses a two-layer cascade to achieve the required accuracy for satellite attitude control.

[0028] The satellite dynamics equation of this embodiment is: Where J is the actual moment of inertia, T d is the disturbance torque, T c is the control torque, ω is the angular velocity, ω × is the angular velocity cross product matrix, h is the flywheel angular momentum, According to J=J0+ΔJ, where J0 is the measured value of the moment of inertia and ΔJ is the uncertainty of the moment of inertia, the dynamic equation can be rewritten as: Where d is the total interference term: Convert the satellite dynamics equations into state-space equations: Where x1 is the state variable 1, x1 = ω; x2 is the state variable 2, x2 = d * , d * is the disturbance variable, f is the function term of the state variable x1, and b0 is the control torque coefficient;

[0029] According to the state space equation, the corresponding linear cascade state observer can be designed.

[0030] Furthermore, the first-layer linear state observer is: Where z 1，1 is the current angular velocity estimate, z 1，2 is the current interference estimate, l1 is the linear coefficient 1 of the first-layer linear state observer, e1 is the difference between the angular velocity measurement value and the angular velocity estimate, and l2 is the linear coefficient 2 of the first-layer linear state observer.

[0031] The tracking error of the first-layer observer is used as the observation target, and the second-layer linear state observer is: Where z 2，1 is the angular velocity tracking error, z 2，2 is the disturbance tracking error, l3 is the linear coefficient of the second-layer linear state observer, e2 is the difference between the angular velocity measurement value and the angular velocity tracking error, l4 is the linear coefficient of the second-layer linear state observer; e1 = yz 1，1 , e2=yz 2，1 , where y is the angular velocity measurement value.

[0032] The linearity of the linear cascade state observer is manifested in that e1 and e2 appear in the equation as linear functions.

[0033] The internal and external disturbances are then estimated via a linear cascade state observer.

[0034] The estimated value of the total interference torque inside and outside the satellite is expressed as

[0035] Use the w parameter to spin and control the direction of the satellite's spin axis. Using the w parameter as a control variable allows you to control the direction of the satellite's spin axis even without attitude information.

[0036] After considering the internal and external disturbances of the compensation estimate, the PD control law of this embodiment is designed as follows: Where D is the differential term coefficient, ω e is the angular velocity error, K is the proportional coefficient, w e The error of the attitude parameter w in the direction of the spin axis is represented by the complex number w e Expressed as: w e =w e1 +iw e2 , the matrix is ​​expressed as: w e =[w e1 , w e2 ,0] T , D=k d J0, K=k p J0, where w e1 w e The real part of w e2 w e The imaginary part, k d is the differential control law coefficient, k p is the proportional control law coefficient.

[0037] ω e =ω-C bd ω d, where C bd ω d is the expected angular velocity of this system, C bd is the attitude transformation matrix from the desired coordinate system to the body coordinate system. The subscript d represents the desired attitude, and no subscript represents the current attitude under this system. Since the satellite is only required to have an angular velocity around the Z axis, only C bd Part of the posture information, and w e Contains the same attitude information that can be obtained from sun sensor data.

[0038] Going further, w e Convert to Euler angles in the order of 3-2-1: Where φ is the roll angle and θ is the pitch angle.

[0039] In this embodiment, the differential control law coefficient k d and proportional control law coefficient k p The method for determining is: k d =2ξω c , where ξ is the system damping ratio, ω c is the undamped oscillation angular frequency of the system.

[0040] In this embodiment, the four linear coefficients of the linear cascade state observer are calculated as follows: Since the linear cascade state observer is linear, it can be simplified to a second-order system using the dynamic equations. Therefore, the parameters of the linear cascade state observer have a second-order system relationship. Leveraging the characteristics of the second-order system, the control law parameters are designed by designing the desired convergence time and damping ratio.

[0041] l3=l1=2ω o , Where ω o is the angular frequency at which the observer observation results converge; the two-layer observer parameters can be determined by designing the expected tracking convergence time.

[0042] Tracking convergence time t o for:

[0043] Next, we examine the estimated total interference torque inside and outside the satellite. To verify the tracking accuracy of the linear cascade state observer, a constant error (in the Z direction) and a sinusoidal external disturbance (in the X and Y directions) were added to verify the linear and nonlinear error tracking capabilities of the linear cascade state observer. The deviation between the tracking results and the set values ​​was observed, as shown in Figure 3. The results in Figure 3 show that the linear cascade state observer has high tracking accuracy for both types of disturbance signals (the constant signal in the Z direction and the sinusoidal signal in the X and Y directions) and can quickly track the disturbance signals within the expected convergence time.

[0044] As shown in Figure 1, the linear cascade state observer needs to obtain the control torque at the previous moment and feed back the estimated total internal and external disturbance results to the controller for compensation to further improve the control accuracy.

[0045] Under the action of two types of disturbance signals (constant signal in Z direction and sinusoidal signal in XY direction), the PD control law designed with w parameter can achieve the control accuracy as shown in Figure 4. e The degree of spin axis pointing deviation is indicated. Under the action of external disturbance, the linear cascade state observer can converge in a shorter time and have higher control accuracy.

[0046] The method of the present invention uses a linear cascade method, which simplifies parameter design and improves the accuracy of disturbance estimation.

[0047] Specific implementation method 2: As an example, to reduce the impact of the linearization process on accuracy, based on the same working principle, a nonlinear state observer can be used to control the satellite's attitude. The details are as follows: Step 1: Design a nonlinear state observer (NLESO); Establish the satellite dynamics equation based on rigid body dynamics and convert it into a state space equation; The satellite dynamics equation is: Where J is the actual moment of inertia, T d is the disturbance torque, T c is the control torque, and ω is the angular velocity, ω × is the angular velocity cross product matrix, h is the flywheel angular momentum; according to J=J0+ΔJ, where J0 is the measured value of the moment of inertia and ΔJ is the uncertainty of the moment of inertia, the dynamic equation is rewritten as: Where d is the total interference term: Convert the satellite dynamics equations into state-space equations: Where x1 is the state variable 1, x1 = ω; x2 is the state variable 2, x2 = d * , d * is the disturbance variable, f is the function term of the state variable x1, and b0 is the control torque coefficient;

[0048] Then design a nonlinear state observer based on the state space equation: Where z1 is the current angular velocity estimate, z2 is the current disturbance estimate, l1 is the nonlinear state observer coefficient 1, e is the deviation between the current angular velocity estimate and the current angular velocity measurement, I2 is the nonlinear state observer coefficient 2, and fal(e) is the nonlinear function of the deviation e, which can be designed according to the characteristics of the observer: e = y - z1 = [e1, e2, e3]T , where y is the angular velocity measurement value, e1 is the X-axis component of the deviation e, e2 is the Y-axis component of the deviation e, and e3 is the Z-axis component of the deviation e; Where i = 1, 2, 3; t is the observation time, t s is the saturation time of the nonlinear function fal(e); Step 2: Estimate the total interference torque inside and outside the satellite through the nonlinear state observer Since disturbance signals are often nonlinear, if the parameters are set appropriately, the nonlinear state observer can observe the disturbance estimate z2 with high accuracy. Under the action of two types of disturbance signals (constant signal in the Z direction and sinusoidal signal in the XY direction), the tracking results are shown in Figure 5.

[0049] Estimation of the total interference torque inside and outside the satellite for:

[0050] The estimated total interference torque inside and outside the satellite Feedback to the controller.

[0051] Step 3: The controller compensates the estimated total disturbance torque inside and outside the satellite Use the w parameter to spin and control the direction of the satellite's spin axis.

[0052] The PD control law adopted by the controller is: Where D is the differential term coefficient, ω e is the angular velocity error, K is the proportional coefficient, w e To characterize the error of the spin axis direction parameter w, the complex number w e Expressed as: w e =w e1 +1w e2 , the matrix is ​​expressed as: w e =[w e1 , w e2 ,0] T , D=k d J0, K=k p J0, where w e1 w e The real part of w e2 w e The imaginary part, k d is the differential control law coefficient, k p is the proportional control law coefficient.

[0053] Will w e Convert to Euler angles in the order of 3-2-1: Where φ is the roll angle and θ is the pitch angle.

[0054] Differential control law coefficient k d and proportional control law coefficient k p The method for determining is: k d =2ξω c , where ζ is the system damping ratio, ω c is the undamped oscillation angular frequency of the system

[0055] Use the calculated control torque T c Perform satellite attitude control in the current cycle; then calculate the control torque T c Input the nonlinear state observer to observe the disturbance of the next cycle and calculate the control torque of the next cycle to realize satellite attitude control.

[0056] Although the present invention is described herein with reference to specific embodiments, it should be understood that these embodiments are merely illustrative of the principles and applications of the invention. It should be understood that many modifications may be made to the illustrative embodiments, and that other arrangements may be devised, without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that the various dependent claims and features described herein may be combined in ways other than those described in the original claims. It should also be understood that features described in conjunction with individual embodiments may be used in conjunction with other described embodiments.

Claims

1. A method for attitude control of a spin detection satellite, characterized in that include, Establish satellite dynamics equations based on rigid body dynamics and convert them into state space equations; A two-layer linear cascade state observer is designed based on the state-space equation, where the first layer of the linear state observer is used to observe and obtain the current angular velocity estimate and the current interference estimate of the satellite; The second-layer linear state observer observes the tracking error of the first-layer linear state observer based on the observation results of the first-layer linear state observer; the tracking error is the difference between the current angular velocity estimate and the current interference value estimate and the current true angular velocity and the current true interference value; Then, the estimated value of the total interference torque inside and outside the satellite is calculated based on the current interference amount estimate and the tracking error; The PD control law is designed using the attitude parameter w that represents the direction of the spin axis, and the control torque is calculated by compensating the estimated value of the total interference torque inside and outside the satellite, which is used for satellite attitude control in the current cycle. The calculated control torque is then input into the linear cascade state observer to observe the interference amount of the next cycle and calculate the control torque of the next cycle to realize satellite attitude control.

2. The attitude control method of a spin detection satellite according to claim 1, characterized in that: The satellite dynamics equation is: Where J is the actual moment of inertia, T d is the disturbance torque, T c is the control torque, ω is the angular velocity, ω × is the angular velocity cross product matrix, h is the flywheel angular momentum, According to J=J0+ΔJ, Where J0 is the measured value of the moment of inertia, ΔJ is the uncertainty of the moment of inertia, Rewrite the kinetic equation as: Where d is the total interference term: Convert the satellite dynamics equations into state-space equations: Where x1 is the state variable 1, x1 = ω; x2 is the state variable 2, x2 = d * , d * is the disturbance variable, f is the function term of the state variable x1, and b0 is the control torque coefficient; 3. The attitude control method of a spin detection satellite according to claim 2, characterized in that: The first-layer linear state observer is: Where z 1,1 is the current angular velocity estimate, z 1，2 is the current interference estimate, l1 is the linear coefficient 1 of the first-layer linear state observer, e1 is the difference between the angular velocity measurement value and the angular velocity estimate, and l2 is the linear coefficient 2 of the first-layer linear state observer.

4. The attitude control method of a spin detection satellite according to claim 3, characterized in that: The second-layer linear state observer is: Where z 2，1 is the angular velocity tracking error, z 2,2 is the disturbance tracking error, l3 is the linear coefficient 1 of the second-layer linear state observer, e2 is the difference between the angular velocity measurement value and the angular velocity tracking error, and l4 is the linear coefficient 2 of the second-layer linear state observer; e1=yz 1,1 ,e2=yz 2，1 , Where y is the angular velocity measurement value.

5. The attitude control method of a spin detection satellite according to claim 4, characterized in that: The estimated value of the total interference torque inside and outside the satellite is expressed as 6. The attitude control method of a spin detection satellite according to claim 5, characterized in that: The PD control law is: Where D is the differential term coefficient, ω e is the angular velocity error, K is the proportional coefficient, w e The error of the attitude parameter w in the direction of the spin axis is represented by the complex number w e Expressed as: In e =in e1 +iw e2 , The matrix is ​​represented as: In e =[in e1 ,In e2 ,0] T , D=k d J0, K=k p J0, Where w e1 w e The real part of w e2 w e The imaginary part, k d is the differential control law coefficient, k p is the proportional control law coefficient.

7. The attitude control method of a spin detection satellite according to claim 6, characterized in that , Will w e Convert to Euler angles in the order of 3-2-1: Where φ is the roll angle and θ is the pitch angle.

8. The attitude control method of a spin detection satellite according to claim 7, characterized in that: Differential control law coefficient k d and proportional control law coefficient k p The method for determining is: Where ζ is the system damping ratio, ω c is the undamped oscillation angular frequency of the system.

9. The attitude control method of a spin detection satellite according to claim 8, characterized in that: The calculation method of the four linear coefficients of the linear cascade state observer is: Where ω o is the angular frequency at which the observer's observation results converge; Tracking convergence time t o for:

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