A spacecraft safety maneuver control method based on a fixed-time observer

Abstract

The application discloses a spacecraft safety maneuver control method based on a fixed-time observer and belongs to the technical field of spacecraft control. The application realizes the method as follows: a motion model and an attitude forbidden area model of a rigid spacecraft attitude system are established; an observer is used to perform actuator fault information reconstruction; a robust CLF constraint and a robust CBF constraint are designed based on the reconstructed actuator fault information; a quadratic programming problem is established to balance control cost and performance on the basis; a robust CLF-CBF QP controller is designed; a fixed-time state expansion observer is constructed to realize observation on external uncertain interference; and a CLF-CBF QP robust safety attitude controller based on the fixed-time state expansion observer is designed. The CLF-CBF QP robust safety attitude controller ensures that the spacecraft can timely avoid the forbidden pointing area even in the case of actuator failure and realize attitude stabilization within a fixed time.

Description

Technical Field

[0001] This invention relates to a spacecraft safety maneuver control method based on a fixed-time observer, belonging to the field of spacecraft control technology. Background Technology

[0002] Because the space environment in which spacecraft perform space missions is relatively complex and harsh, in order to successfully complete the mission, it is necessary not only to maintain attitude stability, but also to consider various state constraints introduced by the flight mission, sensitive payloads, and actuators. Different missions have different attitude requirements; therefore, researching a feasible and reliable safe attitude control method has practical engineering significance, enabling spacecraft to complete space missions while satisfying all attitude constraints, and ensuring that onboard equipment can always function normally, thereby guaranteeing the safety of the spacecraft and the successful completion of the mission. Figure 1 The diagram shown is a block diagram of the attitude control system for which this invention is applied.

[0003] To address the safety constraint problem of forward invariance of sets, a paper titled "Control Barrier Function Based Quadratic Programs for Safety Critical Systems" was published in IEEE Transactions on Automatic Control in 2017. This paper proposed the Control Barrier Function (CBF) method, which is then combined with algorithms such as Control Lyapunov Function (CLF), backstepping, and quadratic programming (QP) to achieve optimal system safety trajectory tracking control. However, none of the methods presented in the paper consider uncertainties such as actuator failures and external disturbances, which can lead to a decrease in the overall performance of the spacecraft system.

[0004] To address the challenge of real-time handling of various attitude constraints, a paper titled "Kinematic Steering Law for Conically Constrained Torque-Limited Spacecraft Attitude Control" was published in the Journal of Guidance, Control, and Dynamics in 2018. This paper introduced a method based on Lyapunov's stability theorem and a logarithmic barrier potential function to derive the steering law for gyroscope attitude control under conical constraints and restricted regions. The algorithm also considered control torque constraints. However, the potential function method in this paper is prone to local minima, causing the system to get trapped in the minimum region and fail to achieve the control objective.

[0005] To address the safety control problem of uncertain systems with unknown parameters, a paper titled "Disturbance Observers for Robust Safety-Critical Control With Control Barrier Functions" was published in IEEE ControlSystems Letters in 2023. This paper proposes a high-gain input observer method that utilizes the first-order time derivative of the control flow factor (CBF) to estimate the time-varying unmodeled dynamics of the CBF with error bounds. However, the controller in this paper converges in finite time, and the convergence time depends on the initial state of the system.

[0006] In order to maintain the stability of the spacecraft's attitude and meet various state constraints, while minimizing uncertainties such as actuator failures and external interference, a spacecraft control method is needed to realize a safe attitude control method for spacecraft that takes into account actuator failures. Summary of the Invention

[0007] To address the challenge of spacecraft carrying numerous sensitive payloads performing attitude orientation maneuvers in complex space environments while avoiding prohibited pointing areas and considering actuator failures, the main objective of this invention is to provide a spacecraft safe maneuver control method based on a fixed-time observer. This method utilizes an extended state observer to reconstruct actuator failure information, designs fixed-time robust control Lyapunov function constraints and robust control obstacle function constraints, and establishes a quadratic programming problem to balance control cost and performance. This ensures that even with actuator failures, the spacecraft can still avoid prohibited pointing areas in a timely manner and achieve attitude stability within a fixed time.

[0008] The objective of this invention is achieved through the following technical solution.

[0009] This invention discloses a spacecraft safety maneuver control method based on a fixed-time observer, whereby the controlled object is the spacecraft's safe attitude. The safe attitude is achieved by defining attitude exclusion zones, reconstructing actuator fault information using an extended state observer, designing a robust CLF-CBF QP controller, and introducing a fixed-time method to construct a robust safe attitude controller based on the observer's fixed time.

[0010] This invention discloses a spacecraft safety maneuver control method based on a fixed-time observer, comprising the following steps:

[0011] Step 1: Establish the motion model and attitude exclusion zone model of the rigid spacecraft attitude system.

[0012] Step 2: Reconstruct actuator fault information using an observer. Based on the reconstructed actuator fault information, design robust CLF constraints and robust CBF constraints. On this basis, establish a quadratic programming problem to balance control cost and performance, and design a robust CLF-CBF QP controller.

[0013] Step 3: Construct a fixed-time state extension observer to observe uncertain external disturbances.

[0014] Step 4: Based on the fixed-time state extension observer from Step 3, design a CLF-CBF QP robust safe attitude controller. The CLF-CBF QP robust safe attitude controller ensures that the spacecraft can still avoid prohibited pointing areas in a timely manner and achieve attitude stability within a fixed time, even in the event of actuator failure.

[0015] Furthermore, the implementation method for step one is as follows:

[0016] Define attitude quaternions q0 is the scalar part of the attitude quaternion, q v =[q1 q2 q3] T For the vector part. The spacecraft's attitude kinematics equations, expressed using unit quaternions, are as follows:

[0017]

[0018] ω=[ω1 ω2 ω3] T This refers to the angular velocity of the spacecraft's attitude relative to the inertial coordinate system, within its own coordinate system. For a = [a1, a2, a3] T , operator × Defined as:

[0019]

[0020] The attitude kinematic equations of the spacecraft, based on the error quaternion attitude model, are as follows:

[0021]

[0022]

[0023] q is the moment of inertia matrix of the spacecraft in its body coordinate system, and it is a positive definite matrix. e It is the body coordinate system F b Relative to a reference coordinate system F r The attitude error quaternion is defined as:

[0024]

[0025] q r It is the body coordinate system F r Relative to reference coordinate system F r The posture quaternion, i.e. the expectation quaternion, It is q r The conjugate quaternion, ω e It is the body coordinate system F b Relative to reference coordinate system F r The error angular velocity is defined as:

[0026]

[0027] C(e) is the body coordinate system relative to the reference coordinate system F. r Coordinate transformation matrix; ω r It is the reference coordinate system relative to the inertial coordinate system F i In the body coordinate system F b The component below, i.e., the desired angular velocity.

[0028] The spacecraft attitude exclusion zone model is as follows:

[0029]

[0030] It represents the unit vector in the inertial coordinate system pointing from the spacecraft's center of mass toward the strong light (heat) source that needs to be avoided; This represents the unit vector along the line of sight of the sensitive payload in the spacecraft's body coordinate system; the cone angle is θ. s The conical range is the orientation-prohibited pointing region, and R is the rotation matrix.

[0031] Furthermore, the implementation method for step two is as follows:

[0032] Define variable x = [q T ,ω T ] T The spacecraft attitude model is organized as follows:

[0033]

[0034]

[0035] These are the state variables of the spacecraft's attitude system. It is the control input torque of the spacecraft's attitude system. This is the variable part that contains actuator fault information. The total fault information of the actuators in the above attitude control model is entirely included in the φ(t) term.

[0036] For controlling the Lyapunov function, the Lyapunov candidate function is selected as follows:

[0037]

[0038] In the formula, a, b, and c are non-negative constants. Since matrix J is symmetric and positive definite, V(x) can be rearranged as follows:

[0039]

[0040] The condition for V(x) to be positive definite is that a>0 and acJ>b. 2 J 2 For the control barrier function, define the function:

[0041]

[0042] When h(x) ≥ 0, the spacecraft attitude system will not enter the attitude restricted area and will always satisfy the safety constraints. Therefore, the control obstacle function is selected as follows:

[0043]

[0044] The Lie derivative operator is used to simplify the formula expression. Combining the control Lyapunov function and the control barrier function, and taking into account the constraint of the actual input torque, a quadratic programming problem is established:

[0045]

[0046] stL f V(x)+L g V(x)(u+d)≤-λV(x)+δ

[0047] L f B(x)+L g B(x)(u+d)≥-γB(x)

[0048] u min ≤u≤u max

[0049] H(x) is an arbitrary positive definite matrix, and k(x) is a locally Lipschitz continuous fundamental control law. For a constant The slack variables for the penalty, λ and γ are manually chosen parameters, u min ,u max These represent the minimum and maximum values ​​of the input torque. To eliminate the adverse effects of time-varying actuator malfunctions, it is necessary to address unknown input disturbances. Compensation is performed, resulting in the final robust CLF-CBF QP controller:

[0050]

[0051]

[0052]

[0053] u min ≤u≤u max

[0054] because Seeking

[0055] Furthermore, the implementation method for step three is as follows:

[0056] To construct the extended state observer, we first define x1 = ω, and set the extended state variable as x2 = φ(t) and let... r(t) is bounded, that is i = 1, 2, 3, It is a positive constant. The spacecraft attitude system state variable x is expanded to:

[0057]

[0058] Let z1 and z2 represent the outputs of the state-expanded observer, respectively, and define e1 = x1 - z1 and e2 = x2 - z2 as the estimation errors of the observer. Then the state-expanded observer is constructed as follows:

[0059]

[0060] Define function sig r (a) is sig r (a)=[|a1| r sign(a1),|a2| r sign(a2),|a3| r sign(a3)] T , Where sign(·) is the sign function. Scalars β1, β2>0, k1∈(0.5,1), and k2=2k1-1.

[0061] Based on the obtained state extension observer, the fixed-time state extension observer is constructed as follows:

[0062]

[0063] Z1 and Z2 represent the outputs of the state extension observer, used to approximate the state variable ω and the unknown fault information φ(t), respectively. α1, α2, β1, β2, and ρ are the observer gains, and o1∈(1-ε,1); θ1∈(1,1+ε); o2=2o1-1; θ2=2θ1-1; and ε is a pre-set sufficiently small number. The indicator function Λ is:

[0064]

[0065] Define E1 = x1 - Z1 and E2 = x2 - Z2, then the observation error is expressed as:

[0066]

[0067]

[0068] Furthermore, step four is implemented as follows:

[0069] The solvability of the quadratic gauge is guaranteed by slack variable δ, and it satisfies... It is a finite constant. λ1, λ2, and γ are all positive constants, k1 = 1 + 1 / μ, k2 = 1 - 1 / μ, and μ > 1. The unknown is handled using the fixed-time state extension observer from step three. The terms are calculated and compensated to obtain a robust safe attitude controller based on a fixed-time observer, namely the CLF-CBF QP.

[0070]

[0071]

[0072]

[0073] u min ≤u≤u max

[0074] The definitions of τ1 and τ2 are the same as in step two. represent The estimated value, output by a fixed-time state expansion observer, can be used to calculate the fault information of the actuator.

[0075] The CLF-CBF QP robust safety attitude controller ensures that the spacecraft can avoid prohibited pointing areas in a timely manner and achieve attitude stability within a fixed time, even in the event of actuator failure.

[0076] Beneficial effects:

[0077] 1. The present invention discloses a spacecraft safety maneuver control method based on a fixed-time observer. The controller is designed based on actuator failure and input torque amplitude limitation, so that the controller can handle the adverse effects of external interference, environmental noise and uncertainty while handling time-varying faults, and has strong robustness.

[0078] 2. This invention discloses a spacecraft safety maneuver control method based on a fixed-time observer. A fixed-time state extension observer is constructed to observe uncertain external disturbances. Based on the fixed-time state extension observer, a robust safety attitude controller based on the fixed-time observer, using a CLF-CBF QP, is designed. Robust CBF constraints and fixed-time robust CLF constraints ensure that the spacecraft can avoid attitude exclusion zones in a timely manner and achieve attitude stability within a fixed time, exhibiting strong real-time performance. By establishing a quadratic programming QP, control cost and performance are balanced, thereby achieving optimal safety attitude control for the spacecraft system.

[0079] 3. The present invention discloses a spacecraft safety maneuver control method based on a fixed-time observer. It reconstructs the fault information of the actuator using an extended state observer, designs fixed-time robust control Lyapunov function constraints and robust control obstacle function constraints, and establishes a solution to the problem of poor real-time performance of the quadratic programming path planning method. The present invention can avoid the attitude prohibition pointing area in a timely manner under the interference of actuator faults, and has strong real-time performance.

[0080] 4. The present invention discloses a spacecraft safety maneuver control method based on a fixed-time observer. It uses a quadratic programming method to balance control cost and performance, which can solve the problem that attitude control methods based on potential functions lack basic optimization capabilities, and also avoids the problem of local minima. Attached Figure Description

[0081] Figure 1 This diagram shows a block diagram of the attitude control system of the present invention;

[0082] Figure 2 The time response curve of the spacecraft attitude quaternion in step two of this invention is shown.

[0083] Figure 3 The time response curve of the attitude angular velocity in step two of the present invention is shown;

[0084] Figure 4 The time response curve of the controller torque u in step two of this invention is shown.

[0085] Figure 5 The time response curves of CLF, CBF, and δ in step two of this invention are shown.

[0086] Figure 6 The three-dimensional spatial attitude maneuvering path of the spacecraft in step two of the present invention is shown.

[0087] Figure 7 The angular velocity observation error curve in step four of this invention is shown;

[0088] Figure 8 The fault information observation error curve in step four of this invention is shown.

[0089] Figure 9 The time response curve of the spacecraft attitude quaternion in step four of this invention is shown.

[0090] Figure 10 The time response curve of the attitude angular velocity in step four of this invention is shown.

[0091] Figure 11 The time response curve of the controller torque u in step four of this invention is shown.

[0092] Figure 12 The time response curves of CLF, CBF, and δ in step four of this invention are shown.

[0093] Figure 13 The three-dimensional spatial attitude maneuvering path of the spacecraft in step four of the present invention is shown. Detailed Implementation

[0094] like Figure 1 As shown in the figure, this embodiment discloses a spacecraft safety maneuver control method based on a fixed-time observer, and the specific implementation steps are as follows:

[0095] Step 1: Define attitude quaternions q0 is the scalar part of the attitude quaternion, q v =[q1 q2 q3] T For the vector part. The spacecraft's attitude kinematics equations, expressed using unit quaternions, are as follows:

[0096]

[0097] ω=[ω1ω2ω3] T This refers to the angular velocity of the spacecraft's attitude relative to the inertial coordinate system, within its own coordinate system. For a = [a1, a2, a3] T , operator × Defined as:

[0098]

[0099] The attitude kinematic equations of a spacecraft, based on the error quaternion mathematical model, are as follows:

[0100]

[0101]

[0102] q is the moment of inertia matrix of the spacecraft in its body coordinate system, and it is a positive definite matrix. e It is the body coordinate system Fb Relative to a reference coordinate system F r The attitude error quaternion is defined as:

[0103]

[0104] q r It is the body coordinate system F r Relative to reference coordinate system F r The posture quaternion, i.e. the expectation quaternion, It is q r The conjugate quaternion, ω e It is the body coordinate system F b Relative to reference coordinate system F r The error angular velocity is defined as:

[0105]

[0106] C(e) is the body coordinate system relative to the reference coordinate system F. r Coordinate transformation matrix; ω r It is the reference coordinate system relative to the inertial coordinate system F i In the body coordinate system F b The component below, i.e., the desired angular velocity.

[0107] The mathematical model for the spacecraft attitude exclusion zone is as follows:

[0108]

[0109] It represents the unit vector in the inertial coordinate system pointing from the spacecraft's center of mass toward the strong light (heat) source that needs to be avoided; This represents the unit vector along the line of sight of the sensitive payload in the spacecraft's body coordinate system; the cone angle is θ. s The conical range is the orientation-restricted pointing area.

[0110] Step 2: Define variable x = [q] T ,ω T ] T The spacecraft attitude model can be organized as follows:

[0111]

[0112]

[0113] These are the state variables of the spacecraft's attitude system. It is the control input torque of the spacecraft's attitude system. This is the variable part that contains actuator fault information. The total fault information of the actuators in the above attitude control model is entirely included in the φ(t) term.

[0114] For controlling the Lyapunov function, the Lyapunov candidate function is selected as follows:

[0115]

[0116] In the formula, a, b, and c are non-negative constants. Since matrix J is symmetric and positive definite, V(x) can be rearranged as:

[0117]

[0118] The condition for V(x) to be positive definite is that a>0 and acJ>b. 2 J 2 For the control barrier function, define the function:

[0119]

[0120] When h(x) ≥ 0, the spacecraft attitude system will not enter the attitude restricted area and will always satisfy the safety constraints. Therefore, the control obstacle function is selected as follows:

[0121]

[0122] Combining the control Lyapunov function and control obstacle function methods, and considering the constraint of actual input torque, a quadratic programming problem is established:

[0123]

[0124] stL f V(x)+L g V(x)(u+d)≤-λV(x)+δ

[0125] L f B(x)+L g B(x)(u+d)≥-γB(x)

[0126] u min ≤u≤u max

[0127] To eliminate the adverse effects of time-varying actuator failures, input disturbances need to be compensated for. Therefore, the final robust CLF-CBF QP controller is as follows:

[0128]

[0129]

[0130]

[0131] u min ≤u≤u max

[0132] because It can be obtained

[0133] The nominal moment of inertia matrix J in the spacecraft attitude model is set as follows:

[0134]

[0135] The time-varying actuator fault is selected as -Eu=[0.2+0.1sin(0.5t) 0.3-0.1sin(t) 0.2-0.1sin(t)]u c u a = [0.4sin(t) -0.4sin(t) -0.5sin(0.5t)] Nm. The initial attitude and initial attitude angular velocity of the spacecraft are set as q = [0.8 0.4 0.2 0.4] T ω = [0.05 -0.03 0.02] T rad / s, the expected value of the attitude quaternion is chosen as q d =[1 0 0 0] T The expected value of the attitude angular velocity is chosen as ω. d =[0 0 0] T rad / s. The unit vector of the sensitive payload along the line of sight in the spacecraft's body coordinate system is... In the inertial coordinate system, the unit vector pointing from the spacecraft's center of mass to the direction of the light source to be avoided is: The initial conditions and gain of the state-extended observer are chosen as z. i =[0 0 0] T Given i = 1, 2, β1 = 10, β2 = 150, k1 = 0.8, k2 = 0.6, after a certain observation time T0, the upper bound of the observer's observation error can be selected as ε2 = 0.005. The design parameters in the robust CLF-CBF QP controller are set as H = 1, p = 1, λ = 10, γ = 1, K... b =[1 2]. The design parameters in the control Lyapunov function are selected as a=40, b=3, c=10. Assume the semi-cone angle of the safety line-of-sight field for the sensitive load is θ. s Select the design parameter θ in the control barrier function. s =15°. The maximum and minimum values ​​of the allowed input torque constraint are u and u, respectively. max =2N and u min = -2N.

[0136] Figure 2 The time response curve of the spacecraft's attitude quaternion. Figure 3 This is the time response curve of the attitude angular velocity. Figure 4 The time response curve of the controller's control torque u. Figure 5 Here are the time response curves for CLF, CBF, and δ. Figure 6 This refers to the three-dimensional spatial attitude maneuvering path of the spacecraft.

[0137] Step 3: To construct the extended state observer, first define x1 = ω, set the extended state variable as x2 = φ(t), and let... Assume r(t) is bounded, i.e. i = 1, 2, 3, It is a positive constant. The spacecraft attitude system state variable x is expanded to:

[0138]

[0139] Let z1 and z2 represent the outputs of the state-expanded observer, respectively, and define e1 = x1 - z1 and e2 = x2 - z2 as the estimation errors of the observer. Then the state-expanded observer is constructed as follows:

[0140]

[0141] Scalars β1, β2>0, k1∈(0.5,1), and k2=2k1-1.

[0142] Based on the obtained state extension observer, the fixed-time state extension observer is constructed as follows:

[0143]

[0144] Z1 and Z2 represent the outputs of the state extension observer, used to approximate the state variable ω and the unknown fault information φ(t), respectively. α1, α2, β1, β2, and ρ are the observer gains, and o1∈(1-ε,1); θ1∈(1,1+ε); o2=2o1-1; θ2=2θ1-1; and ε is a sufficiently small number. The indicator function Λ is:

[0145]

[0146] Define E1 = x1 - Z1 and E2 = x2 - Z2, then the observation error is expressed as:

[0147]

[0148]

[0149] Step 4: Based on the robust CLF-CBF QP controller obtained in Step 2, and by unifying the fixed-time stability and safety constraints while also considering the limitation on the input torque amplitude, the quadratic programming problem can be obtained:

[0150]

[0151]

[0152] L f B(x)+L g B(x)(u+d)≥-γB(x)

[0153] u min ≤u≤u max

[0154] The slack variable δ guarantees the solvability of the quadratic gauge and satisfies It is a finite constant. λ1, λ2, and γ are all positive constants, k1 = 1 + 1 / μ, k2 = 1 - 1 / μ, and μ > 1. The unknown is handled using the fixed-time state extension observer from step three. After calculating and compensating for the items, the robust safe attitude controller based on the fixed-time observer CLF-CBF QP can be obtained:

[0155]

[0156]

[0157]

[0158] u min ≤u≤u max

[0159] The definitions of τ1 and τ2 are the same as in step two. represent The estimated value, output by a fixed-time state expansion observer, can be used to calculate the fault information of the actuator.

[0160] The nominal moment of inertia matrix J in the spacecraft attitude model is set as follows:

[0161]

[0162] Actuator failure is still selected as -Eu c =[0.2+0.1sin(t) 0.2+0.1sin(t) 0.2+0.1sin(t)]u c u a= [0.5sin(t) -0.5sin(t) 0.5sin(t)] Nm. The initial attitude and initial attitude angular velocity for the unit quaternion are chosen as q = [0.8 0.4 0.2 0.4] T ω = [0.1 0.1 0.1] T rad / s, the expected value of the quaternion is set to q d =[1 0 0 0] T The desired angular velocity is ω d =[0 0 0] T rad / s. The unit vector of the sensor's line of sight in the spacecraft's body coordinate system is... In the inertial coordinate system, the unit vector pointing from the spacecraft's center of mass to the direction of the light source to be avoided is: The initial conditions and gain of the fixed-time state-extended observer are set to z. i =[0 0 0] T Given i = 1, 2, α1 = 3, α2 = 1.5, β1 = 15, β2 = 7.5, o1 = 0.9, θ1 = 1.01, after a certain observation time T0, the upper bound of the observation error ε2 = 0.005. The parameters of the CLF-CBF QP robust safety attitude controller based on a fixed-time observer are designed as H = 1, p = 1, λ1 = λ2 = 3, 1 / μ = 0.9, k1 = 1.9, k2 = 0.1, γ = 1. The parameters in the control Lyapunov function are selected as a = 30, b = 3, c = 8. The semi-cone angle of the safety line-of-sight field of the sensitive element is θ. s Design the parameter θ in the control barrier function s =15°. The maximum and minimum values ​​of the allowed input torque constraint are u and u, respectively. max =2N and u min = -2N.

[0163] Figure 7 The curve showing the observation error of angular velocity ω. Figure 8 The error curve for observing fault information φ. Figure 9 The time response curve of the spacecraft's attitude quaternion. Figure 10 This is the time response curve of the attitude angular velocity. Figure 11 The time response curve of the controller's control torque u. Figure 12 Here are the time response curves for CLF, CBF, and δ. Figure 13 This refers to the three-dimensional spatial attitude maneuvering path of the spacecraft.

[0164] The above detailed description further illustrates the purpose, technical solution, and beneficial effects of the invention. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for spacecraft safety maneuver control based on fixed-time observer, the control object is the safety attitude of spacecraft; characterized in that: Comprising the following steps, Step one: establish the motion model of rigid spacecraft attitude system and the attitude forbidden zone model; Step two: use the observer to reconstruct the actuator fault information, design the robust CLF constraint and the robust CBF constraint based on the reconstructed actuator fault information, and establish a quadratic programming problem to balance the control cost and performance on this basis, and design a robust CLF-CBF QP controller; Step three: construct a fixed-time state expanding observer to observe the external uncertain disturbance; Step four: design a CLF-CBF QP robust safety attitude controller based on the fixed-time observer according to the fixed-time state expanding observer in step three, which ensures that the spacecraft can avoid the forbidden pointing area in time and realize attitude stabilization within a fixed time through the CLF-CBF QP robust safety attitude controller even if the actuator fails; The implementation method of step four is: The solvability of the quadratic form is guaranteed by the slack variable δ, and satisfies is a finite constant; λ1, λ2, γ are all normal numbers, k1 = 1 + 1 / μ, k2 = 1 - 1 / μ, μ > 1; the unknown term is processed and compensated by the fixed-time state-extended observer in step three, and the CLF-CBF QP robust safe attitude controller based on the fixed-time observer is obtained: τ1 and τ2 are defined in step two, represent The estimated value of the output of the fixed-time state expansion observer, by operation, can obtain the fault information of the actuator The implementation method of step one is, 2. The fixed-time observer based spacecraft safety maneuver control method of claim 1, wherein: The attitude kinematics equation of the spacecraft is based on the error quaternion attitude model: Definition of attitude quaternion q0is the scalar part of the attitude quaternion, q v = [q1q2 q3] T is the vector part; the attitude kinematics equation of the spacecraft is based on the expression of the unit quaternion as: ω = [ω1 ω2 ω3] T is the rotation angular velocity of the spacecraft attitude relative to the inertial coordinate system under the body coordinate system; for a = [a1, a2, a3] T , the operator x is defined as: The spacecraft attitude forbidden zone model is: is the inertia matrix of the spacecraft in the body coordinate system, which is a positive definite matrix; q e is the attitude error quaternion of the spacecraft relative to a reference coordinate system F b is the inertia matrix of the spacecraft in the body coordinate system, which is a positive definite matrix; q r is the attitude error quaternion of the spacecraft relative to a reference coordinate system F q r It is the body coordinate system F r Relative to reference coordinate system F r The posture quaternion, i.e. the expectation quaternion, It is q r The conjugate quaternion, ω e It is the body coordinate system F b Relative to reference coordinate system F r The error angular velocity is defined as: C(e) is the body coordinate system relative to the reference coordinate system F. r Coordinate transformation matrix; ω r It is the reference coordinate system relative to the inertial coordinate system F i In the body coordinate system F b The component below, i.e., the desired angular velocity; The implementation method of step two is, denotes the unit vector pointing from the spacecraft body center of mass to the direction of the strong light (heat) source to be avoided in the inertial coordinate system; denotes the unit vector of the sensitive payload boresight direction in the spacecraft body coordinate system; the coning angle is θ s The conical range of is the attitude forbidden pointing area, and R is the rotation matrix.

3. The fixed-time observer based spacecraft safety maneuver control method of claim 1, wherein: For the control Lyapunov function, the Lyapunov candidate function is selected as: Define the variable x = [q T , ω T ] T , and rearrange the spacecraft attitude model as: is a state variable of the spacecraft attitude system, is a control input torque of the spacecraft attitude system, is a variable part containing actuator fault information; the total actuator fault information of the above attitude control model is contained in the φ(t) term; In the formula, a, b, c are non-negative constants, since the matrix J is symmetric and positive definite, V(x) is arranged as: When h(x) ≥ 0, the spacecraft attitude system will not enter the attitude forbidden zone, and always satisfies the safety constraint condition; Therefore, the control barrier function is selected as: The condition for V(x) to be positive definite is a > 0, acJ > b 2 J 2 For the control barrier function, define the function: The implementation method of step three is: For the Lie derivative operator used to simplify the formula expression, combined with the control Lyapunov function and the control barrier function, and combined with the restriction of the actual input torque, a quadratic programming problem is established: H(x) is an arbitrary positive definite matrix, k(x) is a local Lipschitz continuous basic control law, is a constant relaxation variable, λ, γ are artificially selected parameters, u min , max is the minimum and maximum value of the input torque, in order to eliminate the adverse effects of time-varying actuator failure, it is necessary to compensate for the unknown input disturbance , so the final robust CLF-CBF QP controller is: Because Obtained 4. The fixed-time observer based spacecraft safety maneuver control method of claim 1, wherein: Let z1 and z2 represent the output of the state expanding observer respectively, and define e1 = x1-z1, e2 = x2-z2 respectively represent the estimation error of the observer; The state expanding observer is constructed as: First define x1= ω, the extended state variable is set as x2= φ(t) and let r(t) is bounded, i.e. i = 1, 2, 3, is a positive constant; the spacecraft attitude system state variable x is extended as: According to the obtained state expanding observer, the fixed-time state expanding observer is constructed as: Definition of function sig r (a) = sig(a1, a2, a3) r (a) = [ |a1| r sign(a1), |a2| r sign(a2), |a3| r sign(a3)] T , where sign(·) is the sign function, scalars β1, β2>0, k1∈(0.5, 1), and k2=2k1-1; Define E1 = x1-Z1, E2 = x2-Z2, then the observation error is represented as: Z1 and Z2 represent the output of the state extended observer, which are used to approximate the state variable ω and the unknown fault information φ(t), respectively; α1, α2, β1, β2 and ρ are the observer gains, and oi∈(1-ε, 1); θ1∈(1, 1+ε); o2=2oi-1; θ2=2θ1-1; and ε is a predetermined sufficiently small number; the indicator function Λ is: ​

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