Adaptive fault-tolerant attitude control method based on dual-channel event triggering of spacecraft

Abstract

The present invention belongs to the field of spacecraft attitude control, and specifically relates to an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering, comprising: real-time acquisition of spacecraft motion state parameter data, and real-time calculation of an adaptive fault-tolerant attitude control signal for controlling the jth actuator based on the motion state parameter data, wherein the adaptive parameter value required for calculating the attitude control signal is obtained by using a parameter estimator to estimate in real time based on the motion state parameter data and the parameter estimator sends it in an event-triggered manner; the attitude control signal is sent to the jth actuator in an event-triggered manner; wherein the event triggering condition of the adaptive parameter value sending channel is: each adaptive parameter value to be transmitted is different from the value at the last transmission by a preset threshold or more; and the event triggering conditions corresponding to each channel are asynchronous. The present invention can reduce communication and calculation burdens.

Description

Technical Field

[0001] The present invention belongs to the field of spacecraft attitude control, and more specifically, relates to an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering. Background Art

[0002] In recent years, spacecraft control technology has received extensive attention and research from aerospace units and scientific research institutions around the world. Spacecraft is a large-scale complex system composed of multiple subsystems such as attitude control, measurement and control, power supply, and payload. It can perform tasks such as space exploration, military reconnaissance, and meteorological observation, and plays an important role in both military and civilian fields. Ensuring the stability of spacecraft attitude by designing appropriate control algorithms is the basis for completing various tasks. However, the actuator failure caused by the complex and harsh operating environment and the communication limitation caused by limited bandwidth have caused difficulties and challenges to the design of attitude control algorithms.

[0003] In order to facilitate maintenance and expansion and reduce costs, spacecraft are gradually developing in a modular direction. Therefore, different spacecraft components may transmit signals through wireless networks to meet plug-and-play requirements, which places high demands on the communication and computing capabilities of spacecraft. However, the limited resources of spacecraft during on-orbit operation will affect communication and computing. For example, limited wireless network communication bandwidth causes communication congestion, and limited power causes insufficient computing power, which ultimately leads to a decrease in control performance and even affects the life of the spacecraft. Therefore, event-triggered control methods that can avoid continuous signal transmission and reduce the frequency of signal transmission have been widely studied. However, the existing event-triggered control methods for spacecraft only focus on event triggering from the controller to the actuator channel, and there is still a problem of waste of communication and computing resources to a certain extent. Summary of the invention

[0004] In view of the defects of the prior art and the need for improvement, the present invention provides an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering, which aims to realize event triggering of the controller to actuator channel and the parameter estimator to controller channel, thereby reducing the communication and computing burden.

[0005] To achieve the above object, according to one aspect of the present invention, there is provided an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering, comprising:

[0006] Collect spacecraft motion state parameter data in real time;

[0007] Based on the motion state parameter data, an adaptive fault-tolerant attitude control signal for controlling the j-th actuator is calculated in real time, wherein the adaptive parameter value required for calculating the attitude control signal is obtained by using a parameter estimator to estimate in real time based on the motion state parameter data and the parameter estimator is sent in an event-triggered manner;

[0008] Sending a posture control signal to the jth actuator in an event-triggered manner;

[0009] Among them, the event triggering condition of the adaptive parameter value sending channel is: the adaptive parameter values ​​to be transmitted differ from the values ​​at the last transmission by a preset threshold or more; the event triggering condition of the attitude control signal sending channel corresponding to the j-th actuator is: the adaptive fault-tolerant attitude control signal values ​​to be transmitted to the j-th actuator differ from the values ​​at the last transmission by a preset threshold or more, and the event triggering conditions corresponding to each channel are asynchronous.

[0010] Furthermore, the adaptive fault-tolerant attitude control signal from the controller to the j-th actuator m represents the total number of actuators; Indicates the moment when the event trigger condition from the controller to the j-th actuator channel is triggered k times; the event trigger condition is: is the preset trigger threshold constant, inf{} means taking the lower bound;

[0011] Intermediate control signal D∈R 3×m is a known steady-state control distribution matrix, and v0 represents the control signal to be distributed calculated based on the spacecraft attitude tracking model.

[0012] Further, when the spacecraft attitude tracking model is derived by using the rigid spacecraft attitude motion model and the rigid spacecraft attitude dynamics model represented by quaternions, the

[0013] In the formula, ε is a positive constant used to avoid the denominator being 0. z2=ω e +2g1z1,z1=q ev , g1 and g2 both represent control gains with preset positive values, ω=[ω1,ω2,ω3] T ∈R 3 represents the spacecraft's attitude angular velocity, ω × represents the antisymmetric matrix of ω, ω e ∈R 3 represents the attitude angular velocity tracking error, Represents ω e The antisymmetric matrix, C = (q e0 2-q ev T q ev )I3+2q ev q ev T -2q e0 q ev × represents the rotation matrix, q ev Represents the vector components of the attitude tracking error quaternion, Indicates q ev The antisymmetric matrix, ω d =[ω d，1 ,ω d，2 ,ω d，3 ] T ∈R 3 represents the desired angular velocity that the spacecraft is to track, q e0 Represents the scalar component of the attitude tracking error quaternion, I3∈R 3×3 is the third-order identity matrix; and are the estimated values ​​of l and p respectively, which are estimated by the parameter estimator. l and p represent unknown constants related to the fault and are used by the controller to handle unknown faults. is the estimated value of G, G=[J 11 , J 22 , J 33 , J 23 , J 13 , J 12 ] T , J represents the unknown positive definite moment of inertia matrix of the spacecraft;

[0014] For any matrix a = [a1, a2, a3] T ∈R 3 , L(a) represents the following operation:

[0015]

[0016] Furthermore, the parameters l, p and G are estimated as follows:

[0017]

[0018]

[0019]

[0020] Among them, t l,k ,t p,k and tG,k They respectively represent the moment when the event triggering condition of the parameter estimator to controller channel is triggered for the kth time;

[0021] The event triggering condition of the parameter estimator to controller channel is:

[0022] t l,k+1 = inf{t>t l,k |||e l (t)||≥T l};

[0023] t p,k+1 = inf{t>t p,k |||e p (t)||≥T p};

[0024] t G,k+1 = inf{t>t G,k |||e G (t)||≥T G};

[0025] e l (t) = τ l (t)-τ l (t l,k );

[0026] e p (t) = τ p (t)-τ p (t p,k );

[0027] e G (t) = τ G (t)-τ G (t G,k );

[0028] T G >0, T l >0, T p >0 are all preset trigger threshold constants, τ l (t), τ p (t) and τ G (t) are and The adjustment function.

[0029] Furthermore, when the intermittent fault model of the actuator is: 0≤ρ j,h ≤1 is the unknown coefficient representing the degree of failure of the hth actuator of the jth actuator. represents the unknown bounded offset control torque of the hth failure of the jth actuator; u j represents the control torque of the jth actuator; and They represent the occurrence and end time of the hth fault of the jth actuator respectively;

[0030] Then ρ=diag{ρ 1,h ,…,ρ m,h}∈R m×m ; Indicated by each The offset fault vector formed, d∈R 3 is a bounded unknown external disturbance; sup{} represents the supremum, λ min (DρD T ) represents the matrix DρD T The minimum eigenvalue of , inf{} means taking the infimum.

[0031] Furthermore, the τ l (t), τ p (t) and τ G The expression of (t) is:

[0032]

[0033]

[0034]

[0035] Where, Γ and Π are the preset positive definite gain matrices; γ l >0,κ l >0,γ p >0,κ p >0 is the preset estimator gain.

[0036] The present invention also provides a computer-readable storage medium, which includes a stored computer program, wherein when the computer program is executed by a processor, the device where the storage medium is located is controlled to execute the adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering as described above.

[0037] The present invention also provides an adaptive fault-tolerant attitude control system based on spacecraft dual-channel event triggering, which is used to execute the adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering as described above, comprising:

[0038] Sensors, used to collect spacecraft motion state parameter data in real time;

[0039] A parameter estimator, used for estimating in real time the values ​​of the adaptive parameters required for calculating the attitude control signal based on the received motion state parameter data;

[0040] A controller is used to calculate an adaptive fault-tolerant attitude control signal from the controller to the jth actuator based on the received motion state parameter data, and transmit the attitude control signal to each actuator in an event-triggered manner; wherein the adaptive parameter values ​​required for calculating the attitude control signal are obtained through parameter estimation, and the adaptive parameter values ​​are transmitted to the controller by the parameter estimator in an event-triggered manner.

[0041] In general, the above technical solutions conceived by the present invention can achieve the following beneficial effects:

[0042] (1) The present invention proposes a dual-channel event-triggered adaptive fault-tolerant attitude control method for spacecraft. The event trigger condition is designed by comparing the difference between two adjacent transmission signals with a preset threshold, so as to achieve the effect of non-periodic signal transmission according to the event trigger mechanism. By designing independent event trigger conditions for different channels, event trigger mechanisms are designed for the attitude control signal sending channel from the controller to each actuator and the adaptive parameter value sending channel from the parameter estimator to the controller, respectively, so as to achieve asynchronous triggering of each channel. While ensuring the fault-tolerant performance of the spacecraft, the signal transmission frequency is reduced, thereby alleviating the communication burden and computing burden of the control system.

[0043] (2) According to the boundary of the intermittent fault of the actuator, the relevant constant parameters are constructed, and the processing of the intermittent fault is converted into the processing of the constant parameters related to the fault; the constructed constant parameters related to the fault are estimated by using adaptive technology, and the estimated signal is transmitted according to the event trigger mechanism from the parameter estimator to the controller channel, thereby removing the real-time integration operation when calculating the parameter estimation value, which can effectively handle the intermittent fault of the actuator and reduce the demand for system computing power, further saving system resources. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 A framework diagram of an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering provided by an embodiment of the present invention;

[0045] Figure 2 A schematic diagram of a quaternion curve of an attitude tracking error provided by an embodiment of the present invention;

[0046] Figure 3 A schematic diagram of an attitude angular velocity tracking error curve provided by an embodiment of the present invention;

[0047] Figure 4 A schematic diagram of a control input curve provided by an embodiment of the present invention;

[0048] Figure 5 A schematic diagram of event triggering from a controller to an actuator channel provided by an embodiment of the present invention;

[0049] Figure 6 A schematic diagram of event triggering of a parameter estimator to controller channel provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0050] In order to make the purpose, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0051] Embodiment 1

[0052] An adaptive fault-tolerant attitude control method based on dual-channel event triggering of spacecraft, such as Figure 1 As shown, including:

[0053] Collect spacecraft motion state parameter data in real time;

[0054] Based on the motion state parameter data, an adaptive fault-tolerant attitude control signal for controlling the j-th actuator is calculated in real time, wherein the adaptive parameter value required for calculating the attitude control signal is obtained by using a parameter estimator to estimate in real time based on the motion state parameter data and sent by the parameter estimator in an event-triggered manner;

[0055] Sending a posture control signal to the jth actuator in an event-triggered manner;

[0056] Among them, the event triggering condition of the adaptive parameter value sending channel is: the adaptive parameter values ​​to be transmitted differ from the values ​​at the last transmission by a preset threshold or more; the event triggering condition of the attitude control signal sending channel corresponding to the j-th actuator is: the adaptive fault-tolerant attitude control signal values ​​to be transmitted to the j-th actuator differ from the values ​​at the last transmission by a preset threshold or more, and the event triggering conditions corresponding to each channel are asynchronous.

[0057] For spacecraft adaptive control systems, both the controller-to-actuator channel and the parameter estimator-to-controller channel require signal transmission. Therefore, designing event trigger mechanisms for these two channels can reduce the signal transmission frequency and communication burden on the one hand, and remove the real-time integration operation when calculating parameter estimates on the other hand, reducing the computational burden.

[0058] As a preferred embodiment, the adaptive fault-tolerant attitude control signal from the controller to the j-th actuator is m represents the total number of actuators; Indicates the moment when the event trigger condition from the controller to the j-th actuator channel is triggered k times; the event trigger condition is: is the preset trigger threshold constant, inf{} means taking the lower bound, that is, when When j (t) Maintain Unchanged until When the event trigger condition is triggered again, the updated control signal is transmitted to v j (t).

[0059] Intermediate control signal D∈R 3×m is a known steady-state control distribution matrix, and v0 represents the control signal to be distributed calculated based on the spacecraft attitude tracking model.

[0060] It should be noted that the adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering proposed in this embodiment can be applied to rigid spacecraft and flexible spacecraft, and can also be applied to various spacecraft attitude motion models, attitude dynamics models, etc.

[0061] As a preferred implementation, when the rigid spacecraft attitude motion model and the rigid spacecraft attitude dynamics model represented by quaternions are used to calculate the control signal, the above

[0062] In the formula, ε is a positive constant used to avoid the denominator being 0. z2=ω e +2g1z1,z1=q ev , g1 and g2 both represent control gains with preset positive values, ω=[ω1,ω2,ω3] T ∈R 3 represents the spacecraft's attitude angular velocity, ω × represents the antisymmetric matrix of ω, ω e ∈R 3 represents the attitude angular velocity tracking error, Represents ω e The antisymmetric matrix, C = (q e0 2 -q ev T q ev )I3+2q ev qev T -2q e0 q ev × represents the rotation matrix, q ev Represents the vector components of the attitude tracking error quaternion, Indicates q ev The antisymmetric matrix, ω d =[ω d,1 ,ω d,2 ,ω d,3 ] T ∈R 3 represents the desired angular velocity that the spacecraft is to track, q e0 Represents the scalar component of the attitude tracking error quaternion, I3∈R 3×3 is the third-order identity matrix; and are the estimated values ​​of l and p respectively, which are estimated by the parameter estimator. l and p represent unknown constants related to the fault and are used by the controller to handle unknown faults. is the estimated value of G, G=[J 11 , J 22 ,J 33 ,J 23 , 13 ,J 12 ] T J represents the unknown positive definite moment of inertia matrix of the spacecraft;

[0063] For any matrix a = [a1, a2, a3] T ∈R 3 , L(a) represents the following operation:

[0064]

[0065] This embodiment adopts the backstepping method to design the controller. First, two intermediate variables z1=q ev , z2=ω e -α1, α1 is the virtual control law, designed as follows: α1 = -2g1z1; then z2 = ω e +2g1z1. Among them, g1>0 is the control gain to be designed, which is an empirical value and is related to the type of spacecraft.

[0066] Exemplarily, the above-mentioned rigid spacecraft attitude motion model represented by quaternion is specifically:

[0067]

[0068]

[0069] In the formula, q0∈R and qv ∈R 3 They are attitude quaternions (which is obtained by a transformation of the attitude angle) and satisfies ω=[ω1,ω2,ω3] T ∈R 3 is the spacecraft’s attitude angular velocity, I3∈R 3×3 is a third-order unit matrix. For any matrix a=[a1,a2,a3] T ∈R 3 ,a × Represents the antisymmetric matrix of a, and its operation method is:

[0070]

[0071] Exemplarily, the attitude dynamics model of the rigid spacecraft is specifically as follows:

[0072]

[0073] Where J∈R 3×3 is the unknown positive definite moment of inertia matrix of the spacecraft, ω × is the antisymmetric matrix of ω, D∈R 3×m is the control distribution matrix, u=[u1,…,u m ]∈R m is the control torque, d∈R 3 is a bounded unknown external disturbance.

[0074] The corresponding spacecraft attitude tracking model is as follows:

[0075] Set the attitude quaternion and angular velocity of the spacecraft to be tracked as and ω d =[ω d,1 ,ω d,2 ,ω d,3 ] T ∈R 3 , then according to the constructed spacecraft attitude kinematics model, attitude dynamics model and actuator intermittent fault model, the spacecraft attitude tracking model can be obtained as:

[0076]

[0077]

[0078]

[0079] In the formula, and ω e ∈R 3They represent the attitude tracking error quaternion and attitude angular velocity tracking error respectively, v=[v1,…,v m ] T ∈R m represents the control signal vector to be designed, m represents the total number of actuators in the spacecraft, ρ = diag{ρ 1,h ,…,ρ m,h}∈R m×m and are the failure matrix and the offset fault vector respectively. To ensure the controllability of the spacecraft under actuator failure, the matrix DρD T The requirement is positive.

[0080] As a preferred implementation, the estimation of parameters l, p and G is expressed as follows:

[0081]

[0082]

[0083]

[0084] Among them, t l,k ,t p,k and t G,k They respectively represent the moment when the event triggering condition of the parameter estimator to controller channel is triggered for the kth time;

[0085] The event triggering conditions for the above parameter estimator to controller channel are:

[0086] t l,k+1 = inf{t>t l,k |||e l (t)||≥T l};

[0087] t p,k+1 = inf{t>t p,k |||e p (t)||≥T p};

[0088] t G,k+1 = inf{t>t G,k |||e G (t)||≥T G};

[0089] e l (t) = τ l (t)-τ l (t l,k );

[0090] e p (t) = τp (t)-τ p (t p,k );

[0091] e G (t) = τ G (t)-τ G (t G,k );

[0092] T G >0, T l >0, T p >0 are all preset trigger threshold constants, τ l (t), τ p (t) and τ G (t) are and The adjustment function.

[0093] The adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering proposed in this embodiment can be applied to various fault models, including various single fault models and various intermittent multiple fault models, and can be used as a preferred implementation method. When the intermittent fault model of the actuator is: 0≤ρ j,h ≤1 is the unknown coefficient representing the degree of failure of the hth actuator of the jth actuator. represents the unknown bounded offset control torque of the hth failure of the jth actuator; u j represents the control torque of the jth actuator; and Respectively represent the occurrence and end time of the hth fault of the jth actuator; and satisfy

[0094] The above parameters The above parameters Indicated by each The offset fault vector formed, d∈R 3 is a bounded unknown external disturbance; sup{} represents the supremum, λ min (DρD T ) represents the matrix DρD T The minimum eigenvalue of , inf{} means taking the infimum.

[0095] When performing complex tasks, spacecraft may be in harsh environments, such as high temperature, radiation, vibration, etc. At the same time, the actuators themselves also have problems such as aging and wear. The combination of these factors causes frequent actuator failures. After the actuator fails, the designed control instructions cannot be executed correctly, and the safety and stability of the control system are affected. To solve this problem, control methods such as adaptive fault-tolerant control, sliding mode fault-tolerant control, and active fault-tolerant control based on fault diagnosis have been proposed. It is worth noting that most existing methods focus on the single fault problem of spacecraft actuators, that is, when designing fault-tolerant control, it is assumed that the fault only occurs once. However, during the operation of the spacecraft, the actuator may have multiple faults intermittently due to the influence of various factors, which makes the existing solutions unable to meet the fault-tolerant control requirements of the spacecraft.

[0096] The method of this embodiment is to handle unknown intermittent faults of actuators. The unknown constants l and p related to the faults are defined as above. It can be seen from the parameter formula that l is related to the lower bound of the eigenvalue of the failure fault correlation matrix, and p is related to the upper bound of the offset fault. l and p are defined according to the intermittent fault, but do not change with the intermittent changes of the fault.

[0097] As a preferred implementation, when the rigid spacecraft attitude motion model, rigid spacecraft attitude dynamics model and actuator intermittent fault model represented by the above quaternion are adopted, the above Where, Γ and Π are the preset positive definite gain matrices; γ l >0,κ l >0,γ p >0,κ p >0 is the preset estimator gain.

[0098] In order to verify the feasibility and effectiveness of the method of this embodiment, the designed fault-tolerant control method is applied to a spacecraft model for numerical simulation.

[0099] In the simulation, the moment of inertia of the spacecraft is:

[0100]

[0101] The control distribution matrix is:

[0102] The external disturbance is d = 0.15[sin(πt), sin(1.2πt), sin(1.5πt)] T N·m, the initial attitude quaternion is The initial attitude angular velocity is ω(0) = [0,0,0] T rad / m. For the desired tracking posture, the initial value of the quaternion is The desired angular velocity is

[0103] The intermittent fault of the actuator is: u2=0.5v2, t∈[20h-15,20h-10), h=1,2,…; u6=v6+0.5, t∈[20h-5,20h), h=1,2,….

[0104] The results of the embodiment are as follows Figure 2-6 shown. Figure 2 is the quaternion curve of attitude tracking error, Figure 2 It can be seen that the spacecraft attitude tracking effect is good under the action of the proposed controller. Figure 3 is the attitude angular velocity tracking error curve, Figure 3 It can be seen that under the action of the proposed controller, the spacecraft attitude angular velocity can track the expected angular velocity well. Figure 4 From left to right and from top to bottom in the figure are the control signals designed for the first to sixth actuators and the actual output signals of the actuators. Figure 4 It can be seen that when a fault occurs, the control signal is quickly adjusted to achieve fault compensation. Figure 5 From top to bottom, they are the event triggering moments of the first to the sixth actuators, respectively. Figure 5 It can be seen that the signal transmission from the controller to the actuator channel is event-triggered. Figure 6 From top to bottom, they are parameter estimators The event triggering time is determined by Figure 6 It can be seen that the signal transmission from the parameter estimator to the controller channel is event-triggered.

[0105] In general, this embodiment can realize event triggering of the controller-to-actuator channel and the parameter estimator-to-controller channel, reduce the communication and computing burden, ensure the stability of the spacecraft system under intermittent failures of the actuator, and make the attitude angle and angular velocity track the expected signal.

[0106] Embodiment 2

[0107] A computer-readable storage medium includes a stored computer program, wherein when the computer program is executed by a processor, the device where the storage medium is located is controlled to execute an adaptive fault-tolerant attitude control method based on dual-channel event triggering of a spacecraft as described above.

[0108] The relevant technical solution is the same as that in Embodiment 1 and will not be described in detail here.

[0109] Embodiment 3

[0110] An adaptive fault-tolerant attitude control system based on spacecraft dual-channel event triggering, used to execute the adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering as described above, comprising:

[0111] Sensors, used to collect spacecraft motion state parameter data in real time;

[0112] A parameter estimator, used for estimating in real time the values ​​of the adaptive parameters required for calculating the attitude control signal based on the received motion state parameter data;

[0113] A controller is used to calculate an adaptive fault-tolerant attitude control signal from the controller to the jth actuator based on the received motion state parameter data, and transmit the attitude control signal to each actuator in an event-triggered manner; wherein the adaptive parameter values ​​required for calculating the attitude control signal are obtained through parameter estimation, and the adaptive parameter values ​​are transmitted to the controller by the parameter estimator in an event-triggered manner.

[0114] The relevant technical solution is the same as that in Embodiment 1 and will not be described in detail here.

[0115] It will be easily understood by those skilled in the art that the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions and improvements made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. An adaptive fault-tolerant attitude control method based on dual-channel event triggering of spacecraft, characterized in that: include: Collect spacecraft motion state parameter data in real time; Based on the motion state parameter data, an adaptive fault-tolerant attitude control signal for controlling the j-th actuator is calculated in real time, wherein the adaptive parameter value required for calculating the attitude control signal is obtained by using a parameter estimator to estimate in real time based on the motion state parameter data and the parameter estimator is sent in an event-triggered manner; Sending a posture control signal to the jth actuator in an event-triggered manner; Among them, the event triggering condition of the adaptive parameter value sending channel is: the adaptive parameter values ​​to be transmitted are all different from the values ​​at the last transmission by a preset threshold or more; the event triggering condition of the attitude control signal sending channel corresponding to the j-th actuator is: the adaptive fault-tolerant attitude control signal values ​​to be transmitted to the j-th actuator are all different from the values ​​at the last transmission by a preset threshold or more, and the event triggering conditions corresponding to each channel are asynchronous; The adaptive fault-tolerant attitude control signal for controlling the j-th actuator m represents the total number of actuators; Indicates the moment when the event trigger condition of the attitude control signal sending channel corresponding to the j-th actuator is triggered k times; the event trigger condition is: is the preset trigger threshold constant, inf{} means taking the lower bound; Intermediate control signal D∈R 3×m is a known constant control distribution matrix, v0 represents the control signal to be distributed calculated based on the spacecraft attitude tracking model; When the spacecraft attitude tracking model is derived by using the rigid spacecraft attitude motion model and the rigid spacecraft attitude dynamics model represented by quaternions, the In the formula, ε is a positive constant used to avoid the denominator being 0. z2=ω e +2g1z1,z1=q ev , g1 and g2 both represent control gains with a preset positive value, ω=[ω1,ω2,ω3] T ∈R 3 represents the spacecraft's attitude angular velocity, ω × represents the antisymmetric matrix of ω, ω e ∈R 3 represents the attitude angular velocity tracking error, Represents ω e The antisymmetric matrix, C = (q e0 2 -q ev T q ev )I3+2q ev q ev T -2q e0 q ev × represents the rotation matrix, q ev Represents the vector components of the attitude tracking error quaternion, Indicates q ev The antisymmetric matrix, ω d =[ω d,1 ,ω d,2 ,ω d,3 ] T ∈R 3 represents the desired angular velocity to be tracked by the spacecraft, qe0 represents the scalar component of the attitude tracking error quaternion, and I3∈R 3×3 is the third-order identity matrix; and are the estimated values ​​of l and p respectively, which are estimated by the parameter estimator. l and p represent unknown constants related to the fault and are used by the controller to handle unknown faults. is the estimated value of G, G=[J 11 , J 22 , J 33 , J 23 , J 13 , J 12 ] T , J represents the unknown positive definite moment of inertia matrix of the spacecraft; For any matrix a = [a1, a2, a3] T ∈R 3 , L(a) represents the following operation:

2. The adaptive fault-tolerant attitude control method according to claim 1, characterized in that: The parameters l, p and G are estimated as follows: Among them, t l,k ,t p,k and t G,k They respectively represent the moment when the event triggering condition of the parameter estimator to controller channel is triggered for the kth time; The event triggering condition of the adaptive parameter value sending channel is: t l,k+1 =inf{t>t l,k |||e l (t)||≥T l }; t p,k+1 =inf{t>t p,k |||e p (t)||≥T p }; t G,k+1 =inf{t>t G,k |||e G (t)||≥T G }; e l (t)=τ l (t)-τ l (t l,k ); e p (t)=τ p (t)-τ p (t p,k ); e G (t)=τ G (t)-τ G (t G,k ); T G >0, T l >0, T p >0 are all preset trigger threshold constants, τ l (t), τ p (t) and τ G (t) are and The adjustment function.

3. The adaptive fault-tolerant attitude control method according to claim 1, characterized in that: When the intermittent fault model of the actuator is: 0≤ρ j,h ≤1 is the unknown coefficient representing the degree of failure of the hth actuator of the jth actuator. represents the unknown bounded offset control torque of the hth failure of the jth actuator; u j represents the control torque of the jth actuator; and They represent the occurrence and end time of the hth fault of the jth actuator respectively; Then ρ=diag{ρ 1,h ,…,ρ m,h }∈R m×m ; Indicated by each The offset fault vector formed, d∈R 3 is a bounded unknown external disturbance; sup{} represents the supremum, λ min (DρD T ) represents the matrix DρD T The minimum eigenvalue of , inf{} means taking the infimum.

4. The adaptive fault-tolerant attitude control method according to claim 3, characterized in that: The τ l (t), τ p (t) and τ G The expression of (t) is as follows: Where, Γ and Π are the preset positive definite gain matrices; γ l >0,κ l >0,γ p >0,κ p >0 is the preset estimator gain.

5. A computer-readable storage medium, characterized in that: The computer-readable storage medium includes a stored computer program, wherein, when the computer program is executed by the processor, the device where the storage medium is located is controlled to execute an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering as described in any one of claims 1 to 4.

6. An adaptive fault-tolerant attitude control system based on dual-channel event triggering of spacecraft, characterized in that: The method for executing an adaptive fault-tolerant attitude control method based on spacecraft dual-channel event triggering as claimed in any one of claims 1 to 4 comprises: Sensors, used to collect spacecraft motion state parameter data in real time; A parameter estimator, for estimating in real time the adaptive parameter values ​​used to calculate the attitude control signal based on the motion state parameter data, and sending the adaptive parameter values ​​to the controller in an event-triggered manner; A controller, configured to calculate an adaptive fault-tolerant attitude control signal from the controller to the j-th actuator based on the motion state parameter data, and transmit the attitude control signal to the j-th actuator in an event-triggered manner; Among them, the event triggering condition of the adaptive parameter value sending channel is: the adaptive parameter values ​​to be transmitted differ from the values ​​at the last transmission by a preset threshold or more; the event triggering condition of the attitude control signal sending channel corresponding to the j-th actuator is: the adaptive fault-tolerant attitude control signal values ​​to be transmitted to the j-th actuator differ from the values ​​at the last transmission by a preset threshold or more, and the event triggering conditions corresponding to each channel are asynchronous.

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