Event-driven asymmetric attitude constraint control method for flexible agile spacecraft

Abstract

The application discloses a kind of based on event-driven flexible agile spacecraft asymmetric attitude constraint control method, first according to the dynamics characteristics in spacecraft attitude control task, the attitude dynamics model of flexible agile spacecraft considering multiple interference is established, including flexible vibration, unknown external disturbance torque and inertia parameter uncertainty;Then modal observer and disturbance observer are designed, the designed modal observer and disturbance observer respectively obtain vibration mode online observation value and unknown external disturbance torque and inertia uncertainty composition of the online observation value of composite disturbance;Next, asymmetric TVIBLF handling attitude angle constraint and constant IBLF handling angular velocity constraint are constructed;Finally, the flexible agile spacecraft attitude controller under the condition of asymmetric full-state constraint when trigger is designed, and event-driven condition is proposed.The application can effectively reduce update frequency and reduce communication resource consumption.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft technology, specifically relating to an event-driven asymmetric attitude constraint control method for flexible and agile spacecraft. Background Technology

[0002] As one of the most complex systems in a spacecraft, the spacecraft control system has always faced numerous severe challenges in its design process. Traditional spacecraft mostly employ time-driven control, updating signals at fixed time intervals. This method inevitably generates a large number of unnecessary update operations, leading to a waste of communication resources, and when the communication frequency is too high, it is prone to communication congestion. The problem of communication congestion should not be underestimated; it can cause control delays, which in turn seriously interfere with the normal operation of the spacecraft, threatening its stability and reliability.

[0003] With the development of aerospace technology, agile spacecraft have emerged. As a three-axis spacecraft capable of rapid attitude maneuvering, agile spacecraft communicate more frequently than ordinary spacecraft, making communication congestion a more prominent issue in their control systems.

[0004] Event-driven control has the advantage of updating signals only when an event occurs, which can effectively reduce the communication frequency and alleviate the bus bandwidth load pressure while ensuring the necessity of communication. However, existing research on event-driven spacecraft attitude control methods has not fully considered the impact of attitude angle pointing constraints and angular rate constraints on agile spacecraft attitude control.

[0005] In view of the above, it is urgent to carry out research on event-driven asymmetric attitude constraint control for agile spacecraft. It has extremely important application value for improving the performance of spacecraft control systems, ensuring the stable operation of spacecraft, and successfully executing complex missions. Summary of the Invention

[0006] To overcome the shortcomings of existing technologies, this invention provides an event-driven asymmetric attitude constraint control method for flexible agile spacecraft. First, based on the dynamic characteristics of the spacecraft's attitude control mission, an attitude dynamics model of the flexible agile spacecraft considering multiple disturbances is established, including flexible vibration, unknown external disturbance torque, and inertial parameter uncertainty. Then, a modal observer and a disturbance observer are designed. These observers obtain online observations of vibration modes and online observations of the composite disturbance consisting of unknown external disturbance torque and inertial uncertainty, respectively. Next, an asymmetric TVIBLF for handling attitude angle constraints and a constant IBLF for handling angular velocity constraints are constructed. Finally, an asymmetric full-state constraint attitude controller for the flexible agile spacecraft under conditional triggering is designed, and event-driven conditions are proposed. This invention can effectively reduce the update frequency and decrease communication resource consumption.

[0007] The technical solution adopted by this invention to solve its technical problem is as follows:

[0008] Step 1: Based on the dynamic characteristics of the spacecraft attitude control mission, establish an attitude dynamic model of the flexible agile spacecraft that considers multiple disturbances, including flexible vibration, unknown external disturbance torque, and inertial parameter uncertainty.

[0009] Step 2: Design a modal observer and a disturbance observer. The designed modal observer and disturbance observer obtain online observation values ​​of vibration modes and online observation values ​​of composite disturbances composed of unknown external disturbance torque and inertial uncertainty, respectively.

[0010] Step 3: Construct the asymmetric TVIBLF for handling attitude angle constraints and the constant IBLF for handling angular velocity constraints;

[0011] Step 4: Design a flexible agile spacecraft attitude controller under asymmetric full-state constraints when conditions are triggered, and propose event-driven conditions.

[0012] Preferably, step 1 specifically comprises:

[0013] Based on the dynamic characteristics of spacecraft attitude control missions, the following dynamic model of agile spacecraft attitude under multiple disturbances is established:

[0014]

[0015] In the formula:

[0016]

[0017] In the formula: Let 'a' be the orbital angular velocity and 'a' be the semi-major axis of the orbit. The attitude angle vector, ψ and ω represent the roll angle, pitch angle, and yaw angle of an agile spacecraft, respectively; ω = [ω x ω y ω z ] T Let η be the angular velocity vector of the spacecraft's intrinsic system relative to the inertial frame; η∈R n Let n be the modal coordinate vector. For the defined total modal velocity, F s ∈R 3×n Let C be the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body, where C = 2ζΩ, and ζ and Ω ∈ R. n×n Let J be an n-dimensional diagonal matrix, representing the damping ratio and modal frequency of the flexible structure, respectively; m =JF s F s T, J∈R 3×3 J is the overall attitude inertia matrix of the flexible spacecraft. m Let J be the rotational inertia matrix of the rigid body part of the flexible spacecraft. m =J m0 +△J, where J m0 Let ΔJ be the nominal part of the rotational inertia of the rigid body portion of the flexible spacecraft, and ΔJ be the bounded unknown inertia matrix. u represents the control torque vector, μ = 3.986012 × 10⁻⁶. 5 km 3 / s 2 For gravitational parameters, sat u = [sat u] x sat u y sat u z ] T It is a control torque vector with saturation constraints, expressed as:

[0018]

[0019] Among them, u max It is the maximum control torque provided by the actuator, denoted as sat u = u + Δu, where u is defined as the control vector, and Δu = [Δu] x Δu y Δu z ] T This indicates the saturation degree of the control torque; d1 = F s CΨ+F s Kη-s(ω)F s Ψ represents the modal vibration correlation term, and d represents the vector of unknown external disturbance torques acting on the agile spacecraft. For the disturbance consisting of an unknown external disturbance torque and an unknown moment of inertia, matrices A and B are represented as follows:

[0020]

[0021] Where K = Ω 2 .

[0022] Preferably, step 2 specifically comprises:

[0023] Design modal observers and multivariable nonlinear disturbance observers to obtain online observations of disturbances;

[0024] The structures of the modal observer and the disturbance observer are as follows:

[0025]

[0026] in and These are the observed values ​​of modes η and Ψ. The observed value of d2 is a composite disturbance consisting of the uncertainty of rotational inertia and the unknown external disturbance torque, P∈R. 2n×2n D is the gain matrix of the modal observer, z is the auxiliary variable of the perturbation observer, and D = diag[d x ,d y ,d z ]>0 is the gain matrix of the interference observer, p(ω)=DJ m0 ω,

[0027]

[0028] Preferably, step 3 specifically comprises:

[0029] Step 3-1: Design the asymmetric time-varying integral barrier Lyapunov function TVIBLF:

[0030]

[0031] Where z i =ii d It is the angular velocity error, k cai (t) and k cbi (t) represents the upper and lower bounds of the attitude angle, δ i Let i be the integral variable. d For the desired attitude angle,

[0032]

[0033] Step 3-2: Define the attitude angle error vector;

[0034]

[0035] in For the desired attitude angle, As an auxiliary variable;

[0036] Step 3-3: Introduce a first-order filter:

[0037]

[0038] Where t ω =diag{t α t θ t ψ The filter error is

[0039] Construct the following constant IBLF to constrain angular velocity:

[0040]

[0041] in It is angular velocity error. It is a continuously differentiable function and has positive constants. satisfy k ωj It is the angular velocity boundary;

[0042] Define angular velocity error:

[0043]

[0044] in It is an asymmetric virtual control vector, where ξ2=[ξ x ξ y ξ z ] T It is an auxiliary variable.

[0045] Preferably, S4 specifically comprises:

[0046] Step 4-1: Design form of asymmetric stable control signal:

[0047]

[0048] Where c1 is a positive design constant.

[0049]

[0050] β i >0 is a constant.

[0051]

[0052] Step 4-2: Design a flexible, agile spacecraft attitude controller under asymmetric full-state constraints when the design conditions are triggered, i.e., in the following form:

[0053]

[0054] Where c2 is a positive design constant. It is a design parameter matrix.

[0055]

[0056] Step 4-3: Design the event triggering mechanism as follows:

[0057] ||(u(t)-v(t))|| 2 ≥b||Z ω || 2 +γ (15)

[0058] In the formula, u(t)=v(t) k ) represents the control input at the time of the last trigger, where t∈[tk t k+1 ) represents the time within two trigger intervals, b is the design parameter, and γ>0.

[0059] A computer program that causes a computer to execute the above-described asymmetric attitude constraint control method.

[0060] An electronic device includes a processor and a memory; the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to enable the electronic device to perform the above-described asymmetric attitude constraint control method.

[0061] A computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the above-described asymmetric attitude constraint control method.

[0062] A chip includes a processor for calling and running a computer program from a memory, causing a device on which the chip is installed to perform the aforementioned asymmetric attitude constraint control method.

[0063] A computer program product includes a computer storage medium storing a computer program, the computer program including instructions executable by at least one processor, which, when executed by the at least one processor, implement the aforementioned asymmetric attitude constraint control method.

[0064] The beneficial effects of this invention are as follows:

[0065] This invention addresses the attitude control problem of agile spacecraft by employing an event-driven mechanism to update the control signal. This ensures that the control signal is updated only when certain conditions are met, remaining unchanged at other times. This effectively reduces the update frequency and communication resource consumption, which is highly advantageous for agile spacecraft with significant communication requirements. Furthermore, it considers the asymmetric time-varying attitude angle and angular rate constraints of agile spacecraft under multiple disturbances. The proposed asymmetric TVIBLF and constant IBLF can directly handle state constraints without error transformation, exhibiting good constraint and control performance. The presence of an observer enhances the system's robustness. Therefore, under multiple disturbances and constraints, the problem of limited communication resources is alleviated, while attitude control performance and system robustness are guaranteed. Attached Figure Description

[0066] Figure 1 The event-driven three-axis attitude constraint tracking curve of this invention is shown in (a) the roll angle. (b) Pitch angle θ, (c) Yaw angle ψ ;

[0067] Figure 2 The event-driven triaxial angular velocity constraint curve of this invention is shown in (a) angular velocity ω. x (b) angular velocity ω y (c) angular velocity ω z ;

[0068] Figure 3 The composite interference estimation error of the present invention Time response curve;

[0069] Figure 4 The modal estimation error z of this invention η Time response curve;

[0070] Figure 5 The modal estimation error z is defined in this invention. ψ Time response curve;

[0071] Figure 6 This is the control torque variation curve of the present invention;

[0072] Figure 7 This is an event-driven time interval diagram of the present invention. Detailed Implementation

[0073] The present invention will be further described below with reference to the accompanying drawings and embodiments.

[0074] The purpose of this invention is to provide an event-driven asymmetric attitude constraint control method for agile spacecraft. This method can handle asymmetric time-varying attitude constraint problems and angular velocity limitation problems under multiple disturbances, and can effectively solve communication resource limitations. This method has attitude constraint performance, strong robustness and communication resource saving capabilities, and is suitable for application in attitude constraint control tasks of agile spacecraft with multiple disturbances.

[0075] To achieve the above objectives, the technical solution adopted by the present invention includes the following steps:

[0076] S1: Based on the dynamic characteristics of spacecraft attitude control missions, establish an attitude dynamics model for flexible agile spacecraft that considers multiple disturbances, including flexible vibration, unknown external disturbance torques, and inertial parameter uncertainties.

[0077] S2: Using the flexible and agile spacecraft attitude control model established in S1, design a modal observer and a disturbance observer. The designed modal observer and disturbance observer can obtain online observation values ​​of vibration modes and online observation values ​​of composite disturbances composed of unknown external disturbance torque and inertial uncertainty, respectively.

[0078] S3: Based on the attitude control model established in S1 and the full-state constraints established in S2, construct the asymmetric TVIBLF for handling attitude angle constraints and the constant IBLF for handling angular velocity constraints respectively.

[0079] S4: Based on the attitude control model established in S1, the observations obtained in S2, and the asymmetric TVIBLF and constant IBLF established in S3, design a flexible agile spacecraft attitude controller under asymmetric full-state constraints when conditions are triggered, and propose event-driven conditions.

[0080] S1 includes:

[0081] First, based on the dynamic characteristics of spacecraft attitude control missions, the following agile spacecraft attitude dynamics model under multiple disturbances is established:

[0082]

[0083] In the formula:

[0084]

[0085] In the formula: Let be the orbital angular velocity, and 'a' be the semi-major axis of the orbit. Furthermore, The attitude angle vector, ψ and ω represent the roll angle, pitch angle, and yaw angle of an agile spacecraft, respectively; ω = [ω x ω y ω z ] T Let η be the angular velocity vector of the spacecraft's intrinsic system relative to the inertial frame. n Let n be the modal coordinate vector. For the defined total modal velocity, F s ∈R 3×n Let C be the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body, where C = 2ζΩ, and ζ and Ω ∈ R. n×n Let J be an n-dimensional diagonal matrix, representing the damping ratio and modal frequency of the flexible structure, respectively. m =JF s F s T , J∈R 3×3 J is the overall attitude inertia matrix of the flexible spacecraft. m Let J be the rotational inertia matrix of the rigid body part of the flexible spacecraft. m =J m0 +△J, where J m0 Let ΔJ be the nominal part of the rotational inertia of the rigid body portion of the flexible spacecraft, and ΔJ be the bounded unknown inertia matrix. u represents the control torque vector, μ = 3.986012 × 10⁻⁶. 5 km 3 / s2 For gravitational parameters, sat u = [sat u] x sat u y sat u z ] T It is a control torque vector with saturation constraints, expressed as:

[0086]

[0087] Among them, u max It is the maximum control torque provided by the actuator, sat u, which can be expressed as sat u = u + Δu, where u is defined as the control vector, and Δu = [Δu] x Δu y Δu z ] T This represents the saturation degree of the control torque. d1 = F s CΨ+F s Kη-s(ω)F s Ψ represents the modal vibration correlation term, and d represents the vector of unknown external disturbance torques acting on the agile spacecraft. For the disturbance consisting of an unknown external disturbance torque and an unknown moment of inertia, matrices A and B are represented as follows:

[0088]

[0089] S2 includes:

[0090] To address the impacts of flexible vibration, inertial uncertainty, and unknown external disturbance torques on flexible and agile spacecraft, a modal observer and a multivariable nonlinear disturbance observer were designed to obtain online observations of these disturbances.

[0091] The structures of the modal observer and the disturbance observer are as follows:

[0092]

[0093] in and These are the observed values ​​of modes η and Ψ. The observed value of d2 is a composite disturbance consisting of the uncertainty of rotational inertia and the unknown external disturbance torque, P∈R. 2n×2n D is the gain matrix of the modal observer, z is the auxiliary variable of the perturbation observer, and D = diag[d x ,d y ,d z ]>0 is the gain matrix of the interference observer, p(ω)=DJ m0 ω,

[0094]

[0095] S3 includes:

[0096] To ensure attitude under asymmetric constraints, an asymmetric time-varying integral-type barrier Lyapunov function (TVIBLF) of the following form was designed.

[0097]

[0098] Where z i =ii d It is the angular velocity error, where k cai (t) and k cbi (t) represents the upper and lower bounds of the attitude angle.

[0099]

[0100] Define attitude angle error vector

[0101]

[0102] in For the desired attitude angle, It is an auxiliary variable.

[0103] Introduce the following first-order filter

[0104]

[0105] Where t ω =diag{t α t θ t ψ The filter error is

[0106] Generally, due to measurement and performance requirements, the angular velocity of agile spacecraft is also limited within finite boundaries. Therefore, the following constant IBLF is constructed to achieve the constraint on angular velocity.

[0107]

[0108] in It is angular velocity error. It is a continuously differentiable function, and there exists a positive constant here. satisfy k ωj This refers to the angular velocity boundary. Angular velocity error is defined.

[0109]

[0110] in It is an asymmetric virtual control vector, where ξ2=[ξ x ξ yξ z ] T It is an auxiliary variable.

[0111] S4 includes:

[0112] Design form of asymmetric stable control signal

[0113]

[0114] Where c1 is a positive design constant.

[0115]

[0116] β i >0 is a constant, in addition

[0117]

[0118] A flexible, agile spacecraft attitude controller under asymmetric full-state constraints when design conditions are triggered can be described as follows:

[0119]

[0120] Where c2 is a positive design constant. It is a design parameter matrix. also

[0121]

[0122]

[0123] The event triggering mechanism is designed as follows:

[0124] ||(u(t)-v(t))|| 2 ≥b||Z ω || 2 +γ (30)

[0125] In the formula, u(t)=v(t) k ) represents the control input at the time of the last trigger, where t∈[t k t k+1 ) represents the time within two trigger intervals, and b is a design parameter, γ>0.

[0126] Example:

[0127] This invention provides a method for attitude control of flexible and agile spacecraft under asymmetric time-varying constraints. The overall idea is as follows:

[0128] 1) To achieve attitude control of a flexible agile spacecraft under asymmetric full-state constraints, an attitude dynamics model of the flexible agile spacecraft is established;

[0129] 2) To address the impact of flexible vibration, inertial uncertainty, and unknown external disturbance torque on flexible agile spacecraft, a modal observer and a multivariable nonlinear disturbance observer were designed for online observation.

[0130] 3) Based on asymmetric TVIBLF and constant IBLF, a flexible agile spacecraft attitude controller under asymmetric full-state constraints is designed for condition triggering, and event-driven conditions are proposed to meet the attitude control requirements of agile spacecraft.

[0131] 1) Establish an attitude dynamics model for the flexible and agile spacecraft;

[0132] First, the attitude dynamics model of the agile spacecraft under multiple disturbances is established as follows:

[0133]

[0134] In the formula:

[0135]

[0136] In the formula: Let be the orbital angular velocity, a = 729 km be the semi-major axis of the orbit, and μ = 3.986012 × 10⁻⁶. 5 km 3 / s 2 ,also, The attitude angle vector, θ and ψ represent the roll angle, pitch angle, and yaw angle of the agile spacecraft, respectively; ω = [ω x ω y ω z ] T Let be the angular velocity vector of the spacecraft's own system relative to the inertial frame, with initial values ​​of Φ(0) = [305-20]. T (deg), ω(0)=[3.5-1-3] T (deg / s).

[0137] η∈R n Let n be the modal coordinate vector. For the defined total modal velocity, F s ∈R 3×n Here is the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body; C = 2ζΩ, K = Ω 2 ζ and Ω∈R n×n Let be an n-dimensional diagonal matrix, representing the damping ratio and modal frequency of the flexible structure, respectively; considering the initial modal variables, damping ratio, modal frequency matrix, and coupling coefficient matrix of the first four modes, we have:

[0138] [η0Ψ0]T =0.001I8

[0139] ζ=diag{[0.00560.00860.01280.0252]}

[0140] Ω=diag{[0.76811.10381.87332.4596]}(rad / s)

[0141] The coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body is:

[0142]

[0143] Among them, J m =JF s F s T , J∈R 3×3 J is the overall attitude inertia matrix of the flexible spacecraft. m Let J be the rotational inertia matrix of the rigid body portion of the flexible spacecraft, and J m =J m0 +△J is a symmetric positive definite matrix, where ΔJ=0.1J m0 J is a bounded unknown inertia matrix. m0 The nominal part of the moment of inertia of the rigid body portion of the flexible spacecraft has the following value:

[0144]

[0145] sat u = [sat u] x sat u y sat u z ] T It is a control torque vector with saturation constraints, expressed as:

[0146]

[0147] Among them, u max =10N is the maximum control torque provided by the actuator. sat u can be expressed as sat u = u + Δu, where u is defined as the control vector, and Δu = [Δu...]. x △u y △u z ] T This indicates the saturation level of the control torque.

[0148] d1=F s CΨ+F s Kη-s(ω)F s Ψ represents the modal vibration correlation term; d represents the vector of unknown external disturbance torques acting on the agile spacecraft.

[0149]

[0150] For a composite disturbance consisting of an unknown external disturbance torque and an unknown inertia, matrices A and B are represented as follows:

[0151]

[0152] 2) To address the impact of flexible vibration, inertial uncertainty, and unknown external disturbance torque on flexible agile spacecraft, a modal observer and a multivariable nonlinear disturbance observer were designed to obtain online observations.

[0153] The structures of the modal observer and the disturbance observer are as follows:

[0154]

[0155] in and These are the observed values ​​of modes η and Ψ. The observed value of the composite disturbance d2 is composed of the uncertainty of rotational inertia and the unknown external disturbance torque. P = 0.0001I8 is the gain matrix of the modal observer, z is the auxiliary variable of the disturbance observer, D = diag{[11.21]} is the gain matrix of the disturbance observer, and p(ω) = DJ m0 ω,

[0156]

[0157] 3) Design of attitude controller under asymmetric full-state constraints;

[0158] To compensate for the nonlinearity of control torque caused by actuator saturation, the following second-order auxiliary system is introduced:

[0159]

[0160] Among them, the positive design constants are c1 = 0.5 and c2 = 0.4. In addition... and ξ2=[ξ x ξ y ξ z ] T It is an auxiliary variable.

[0161] To ensure attitude under asymmetric constraints, the following form of asymmetric TVIBLF was designed:

[0162]

[0163] Where k cai (t) and k cbi (t) represents the upper and lower bounds of the attitude angle, and the attitude boundary function satisfies:

[0164]

[0165] Where z i =ii d -ξ i This refers to the attitude angle error, defined as the attitude angle error vector:

[0166]

[0167] in Assuming the flexible agile spacecraft needs to perform three maneuvers in a short period of time to achieve the desired attitude, the desired attitudes for the three maneuvers are Φ. 1d =[-102815] T (deg), Φ 2d =[24-10-15] T (deg) and Φ 3d =[-152022] T (deg).

[0168]

[0169] Introduce the following first-order filter:

[0170]

[0171] Where t ω =diag{0.05 0.05 0.05}, the filter error is

[0172] Generally, due to measurement and performance requirements, the angular velocity of agile spacecraft is also limited within finite boundaries. Therefore, the following constant IBLF is constructed to constrain the angular velocity.

[0173]

[0174] in It is the angular velocity boundary. It is angular velocity error. It is a continuously differentiable function, and there exists a positive constant here. satisfy Define angular velocity error:

[0175]

[0176] in It is an asymmetric virtual control vector.

[0177] The design form of the asymmetric stable control signal is as follows:

[0178]

[0179] in:

[0180]

[0181] k Φ =diag{[0.80.60.5]}

[0182]

[0183] β i =0.001, in addition

[0184]

[0185] The design form of the attitude controller for a flexible and agile spacecraft under asymmetric full-state constraints with conditional triggering is as follows:

[0186]

[0187] Wherein, the design parameter matrix k ω =diag{[121015]},

[0188]

[0189] The event triggering mechanism is as follows:

[0190] ||(u(t)-v(t))‖ 2 ≥a||Z ω || 2 +γ (47)

[0191] In the formula, u(t)=v(t) k ) represents the control input at the time of the last trigger, where t∈[t k t k+1 ) represents the time within two trigger intervals, where b = 0.05 and γ = 0.005;

[0192] Numerical simulations have verified that the attached... Figures 1-7 , Figures 1 to 2 The figures show the attitude and angular velocity curves of the agile satellite under an event-driven mechanism. It can be seen that under the event-driven mechanism, the designed full-state constraint controller still exhibits good constraint performance and attitude tracking performance, and the angular velocity variation curve remains within the constant constraint range. From... Figures 3 to 5 It can be seen that, under the event-driven mechanism, the designed interference observer and modal observer still have good observation results. Furthermore, in Figure 3 In the middle, the interference observer's observation accuracy for composite interference is 0.015N. Figures 4 to 5In the modal error observation plot, the observation precision of the modal error and the defined modal error are 2×10⁻⁶ respectively. -3 With 3×10 -3 . Figure 6 As shown in the curve of the control torque change, it can be seen that when the triggering condition is not met, the control torque change curve is a short horizontal line, that is, the control torque remains the same as when the triggering condition was met last time. Figure 7 The figure shows the time distribution of event updates and the sampling interval of event-driven events. The minimum sampling interval is 0.01s, and no Zeno phenomenon occurred. It can be seen that after the attitude tracking is stable, the event-driven time is mostly 0.15s to 0.2s. In this simulation, the time was 85s, the sampling interval was 0.01s, and a total of 498 events were triggered. Compared with the time-driven control that updates every 0.01s, communication resources were saved by about 94.13%, and the attitude control performance was better.

[0193] The simulation results above show that the present invention can achieve agile spacecraft attitude control under full-state constraints when the system has flexible vibration, inertia uncertainty and external disturbances.

Claims

1. An event-driven asymmetric attitude constraint control method for flexible and agile spacecraft, characterized in that, Includes the following steps: Step 1: Based on the dynamic characteristics of the spacecraft attitude control mission, establish an attitude dynamic model of the flexible agile spacecraft that considers multiple disturbances, including flexible vibration, unknown external disturbance torque, and inertial parameter uncertainty. Step 2: Design a modal observer and a disturbance observer. The designed modal observer and disturbance observer obtain online observation values ​​of vibration modes and online observation values ​​of composite disturbances composed of unknown external disturbance torque and inertial uncertainty, respectively. Step 3: Construct the asymmetric TVIBLF for handling attitude angle constraints and the constant IBLF for handling angular velocity constraints; Step 4: Design a flexible agile spacecraft attitude controller under asymmetric full-state constraints when conditions are triggered, and propose event-driven conditions.

2. The event-driven asymmetric attitude constraint control method for flexible and agile spacecraft according to claim 1, characterized in that, Step 1 specifically involves: Based on the dynamic characteristics of spacecraft attitude control missions, the following dynamic model of agile spacecraft attitude under multiple disturbances is established: (1) In the formula: In the formula: The orbital angular velocity, For the semi-major axis of the track; The attitude angle vector, and These represent the roll angle, pitch angle, and yaw angle of an agile spacecraft, respectively. Let be the angular velocity vector of the spacecraft's own system relative to the inertial frame; for 3D modal coordinate vector, For the defined total modal velocity, This represents the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body. , and for A diagonal matrix, representing the damping ratio and modal frequency of the flexible structure, respectively; , Let be the total attitude inertia matrix of the flexible spacecraft. Let the rotational inertia matrix of the rigid body part of the flexible spacecraft be... In the formula This refers to the nominal part of the rotational inertia of the rigid body portion of a flexible spacecraft. The unknown inertia matrix is ​​bounded; Represents the control torque vector. For gravitational parameters, It is a control torque vector with saturation constraints, expressed as: in, It is the maximum control torque provided by the actuator. Represented as , Defined as a control vector, Indicates the degree of saturation of the control torque; For modal vibration related terms, This represents the vector of unknown external disturbance torques acting on an agile spacecraft. For disturbances consisting of unknown external disturbance torque and unknown inertia, the matrix... , The matrix is ​​represented as follows: in, .

3. The event-driven asymmetric attitude constraint control method for flexible and agile spacecraft according to claim 2, characterized in that, Step 2 specifically involves: Design modal observers and multivariable nonlinear disturbance observers to obtain online observations of disturbances; The structures of the modal observer and the disturbance observer are as follows: (2) (3) in and It is modal and The observed values, The disturbance is a complex disturbance consisting of uncertainty in rotational inertia and unknown external disturbance torque. The observed values, It is the gain matrix of the modal observer. It is an auxiliary variable for the perturbation observer. It is the gain matrix of the interference observer. , 。 4. The event-driven asymmetric attitude constraint control method for flexible and agile spacecraft according to claim 3, characterized in that, Step 3 specifically involves: Step 3-1: Design the asymmetric time-varying integral barrier Lyapunov function TVIBLF: (4) in It is angular velocity error. and These are the upper and lower bounds of the attitude angle. For integration variables, For the desired attitude angle, Step 3-2: Define the attitude angle error vector; (5) in For the desired attitude angle, As an auxiliary variable; Step 3-3: Introduce a first-order filter: (6) in The filter error is ; Construct the following constant IBLF to constrain angular velocity: (7) in It is angular velocity error. It is a continuously differentiable function and has positive constants. , satisfy , It is the angular velocity boundary; Define angular velocity error: (8) in It is an asymmetric virtual control vector, where It is an auxiliary variable.

5. The event-driven asymmetric attitude constraint control method for flexible and agile spacecraft according to claim 4, characterized in that, Step 4 specifically involves: Step 4-1: Design form of asymmetric stable control signal: (9) in It is a positive design constant. ， It is a constant. (10) (11) Step 4-2: Design a flexible, agile spacecraft attitude controller under asymmetric full-state constraints when the design conditions are triggered, i.e., in the following form: (12) in It is a positive design constant. It is a design parameter matrix. , , , (13) (14) Step 4-3: Design the event triggering mechanism as follows: (15) In the formula, This represents the control input at the time of the last trigger, where, This represents the time between two trigger intervals. For design parameters, .

6. An electronic device, characterized in that, include: Processor and memory; The memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to cause the electronic device to perform the method as described in any one of claims 1 to 5.

7. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the computer program is executed by a processor, it implements the method as described in any one of claims 1 to 5.

8. A chip, characterized in that, include: A processor for retrieving and running a computer program from memory, causing a device on which the chip is mounted to perform the method as described in any one of claims 1 to 5.

9. A computer program product, characterized in that, The computer program product includes a computer storage medium storing a computer program, the computer program including instructions executable by at least one processor, which, when executed by the at least one processor, implement the method as described in any one of claims 1 to 5.

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