Flexible agile spacecraft asymmetric attitude constraint control method based on event driving

Abstract

The invention discloses a flexible agile spacecraft asymmetric attitude constraint control method based on event driving, and the method comprises the steps: building a flexible agile spacecraft attitude dynamic model considering multiple interferences according to the dynamic characteristics in a spacecraft attitude control task; comprising flexible vibration, unknown external disturbance torque and inertia parameter uncertainty; then designing a modal observer and a disturbance observer, wherein the designed modal observer and disturbance observer respectively obtain a vibration modal online observation value and an online observation value of composite disturbance composed of unknown external disturbance torque and inertia uncertainty; the method comprises the following steps: constructing an asymmetric TVIBLF for processing attitude angle constraint and a constant IBLF for processing angular velocity constraint; and finally, designing a flexible agile spacecraft attitude controller under the asymmetric full-state constraint when the condition is triggered, and proposing an event driving condition. The invention can effectively reduce the updating frequency and reduce the consumption of communication resources.

Description

Technical Field

[0001] The present invention belongs to the technical field of spacecraft, and in particular relates to an event-driven asymmetric attitude constraint control method for a flexible and agile spacecraft. Background Art

[0002] As one of the most complex systems in a spacecraft, the design process of the spacecraft control system has always faced many severe challenges. Traditional spacecraft mostly use time-driven control, which updates signals at fixed time intervals. This method inevitably generates a large number of unnecessary update operations, resulting in a waste of communication resources, and when the communication frequency is too high, it is very easy to cause communication congestion. The problem of communication congestion cannot be underestimated. It will cause control delays, which will seriously interfere with the normal operation of the spacecraft and threaten the stability and reliability of the spacecraft.

[0003] With the development of aerospace technology, agile spacecraft came into being. Agile spacecraft is a spacecraft with three-axis rapid attitude maneuvers. Compared with ordinary spacecraft, its communication is more frequent, which makes the communication blocking problem more prominent in the control system of agile spacecraft.

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

[0005] In view of the above situation, it is urgent to carry out research on event-driven asymmetric attitude constraint control of agile spacecraft, which has extremely important application value for improving the performance of spacecraft control systems, ensuring the stable operation of spacecraft, and the successful execution of complex tasks. Summary of the invention

[0006] In order to overcome the shortcomings of the prior art, the present invention provides an event-driven flexible agile spacecraft asymmetric attitude constraint control method. First, according to the dynamic characteristics of the spacecraft attitude control task, the attitude dynamics model of the flexible agile spacecraft considering multiple interferences is established, including flexible vibration, unknown external interference torque and inertial parameter uncertainty; then the modal observer and the interference observer are designed, and the designed modal observer and interference observer respectively obtain the online observation value of the vibration mode and the online observation value of the composite interference composed of the unknown external interference torque and the inertial uncertainty; then the asymmetric TVIBLF for processing attitude angle constraints and the constant IBLF for processing angular velocity constraints are constructed; finally, the flexible agile spacecraft attitude controller under the asymmetric full-state constraint when the condition is triggered is designed, and the event-driven condition is proposed. The present invention can effectively reduce the update frequency and reduce the consumption of communication resources.

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

[0008] Step 1: According to the dynamic characteristics of the spacecraft in the attitude control mission, the attitude dynamics model of the flexible agile spacecraft considering multiple disturbances is established, including flexible vibration, unknown external disturbance torque and inertial parameter uncertainty;

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

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

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

[0012] Preferably, the step 1 is specifically:

[0013] According to the dynamic characteristics of the spacecraft attitude control task, the agile spacecraft attitude dynamics model under multiple interferences is established as follows:

[0014]

[0015] Where:

[0016]

[0017] Where: is the orbital angular velocity, a is the orbital semi-major axis; is the attitude angle vector, and ψ represent the roll angle, pitch angle and yaw angle of the agile spacecraft respectively; ω=[ω x ω y ω z ] T is the angular velocity vector of the spacecraft system relative to the inertial system; η∈R n is the n-dimensional modal coordinate vector, is the total modal velocity defined, F s ∈R 3×n is the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body, C = 2ζΩ, ζ and Ω∈R n×n is an n-dimensional diagonal matrix, which represents the damping ratio and modal frequency of the flexible structure respectively; J m =JF s F s T, J ∈ R 3×3 is the total attitude inertia matrix of the flexible spacecraft, J m is the inertia matrix of the rigid body part of the flexible spacecraft and J m = J m0 + ΔJ, where J m0 is the nominal part of the inertia of the rigid body part of the flexible spacecraft, and ΔJ is a bounded unknown inertia matrix. u represents the control torque vector, μ = 3.986012×10 5 km 3 / s 2 is the gravitational parameter, sat u = [sat u x sat u y sat u z T is the control torque vector with saturation constraints, and the expression is:

[0018]

[0019] where, u max is the maximum control torque provided by the actuator, sat u is expressed as sat u = u + Δu, u is defined as the control vector, Δu = [Δu x Δu y Δu z T represents the saturation of the control torque; d 1 = F s CΨ + F s Kη - s(ω)F s Ψ is the term related to modal vibration, d represents the vector of unknown external disturbance torques acting on the agile spacecraft, is the disturbance composed of the unknown external disturbance torque and the unknown inertia, and the matrices A and B are represented as the following matrices:

[0020]

[0021] where, K = Ω 2 .

[0022] Preferably, step 2 is specifically:

[0023] Design a modal observer and a multivariable nonlinear disturbance observer to obtain the online observation value of the disturbance;

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

[0025]

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

[0027]

[0028] Preferably, the step 3 is specifically:

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

[0030]

[0031] where z i =ii d is the angular velocity error, k cai (t) and k cbi (t) is the upper and lower bounds of the attitude angle, δ i is the integral variable, i d is the desired attitude angle,

[0032]

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

[0034]

[0035] in is the desired attitude angle, is 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 error of the filter is

[0039] Construct the following constant IBLF to implement the constraint on angular velocity:

[0040]

[0041] in is the angular velocity error, is a continuously differentiable function, and there exists a positive constant satisfy k ωj is the angular velocity boundary;

[0042] Define angular velocity error:

[0043]

[0044] in is the asymmetric virtual control vector, where ξ 2 =[ξ x ξ y ξ z ] T is an auxiliary variable.

[0045] Preferably, S4 is specifically:

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

[0047]

[0048] where c 1 is the positive design constant,

[0049]

[0050] β i >0 is a constant,

[0051]

[0052] Step 4-2: Design a flexible and agile spacecraft attitude controller under asymmetric full-state constraints when the condition is triggered, which is as follows:

[0053]

[0054] where c 2 is the positive design constant, is the design parameter matrix,

[0055]

[0056] Step 4-3: Design the event trigger 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 last trigger, where t∈[t k t k+1 ) represents the time between two trigger intervals, b is the design parameter, and γ>0.

[0059] A computer program enables a computer to execute the above-mentioned asymmetric posture constraint control method.

[0060] An electronic device comprises: 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, so that the electronic device executes the above-mentioned asymmetric posture constraint control method.

[0061] A computer-readable storage medium stores a computer program, which implements the above-mentioned asymmetric posture constraint control method when executed by a processor.

[0062] A chip includes: a processor, which is used to call and run a computer program from a memory, so that a device equipped with the chip executes the above-mentioned asymmetric posture constraint control method.

[0063] A computer program product, the computer program product comprising a computer storage medium, the computer storage medium storing a computer program, the computer program comprising instructions executable by at least one processor, and the above-mentioned asymmetric posture constraint control method is implemented when the instructions are executed by the at least one processor.

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

[0065] The present invention aims at the attitude control problem of agile spacecraft, and adopts an event-driven mechanism to update the control signal, so that the control signal is updated only when the conditions are met, and the control signal remains unchanged at other times, which can effectively reduce the update frequency and reduce the consumption of communication resources. This is very beneficial for agile spacecraft with large communication requirements. In addition, the asymmetric time-varying attitude angle constraint and angular rate constraint problem of agile spacecraft under multiple interferences are also considered. The proposed asymmetric TVIBLF and constant IBLF can directly process the state constraints without error transformation, and have good constraint performance and control performance. The existence of the observer improves the robustness of the system. Therefore, in the presence of multiple interferences and constraints, the problem of limited communication resources can be alleviated, and the attitude control performance and system robustness are guaranteed. BRIEF DESCRIPTION OF THE DRAWINGS

[0066] Figure 1 is the three-axis attitude constraint tracking curve driven by the event of the present invention, (a) rolling angle (b) pitch angle θ, (c) yaw angleψ ;

[0067] Figure 2 is the three-axis angular velocity constraint curve under event-driven of the present invention, (a) angular velocity ω x , (b) angular velocity ω y , (c) angular velocity ω z ;

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

[0069] Figure 4 is the modal estimation error z of the present invention η The time response curve of

[0070] Figure 5 The modal estimation error z is defined as ψ The time response curve of

[0071] Figure 6 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 DESCRIPTION

[0073] The present invention is further described below in conjunction with the accompanying drawings and embodiments.

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

[0075] In order to achieve the above object, the technical solution adopted by the present invention comprises the following steps:

[0076] S1: According to the dynamic characteristics of the spacecraft in the attitude control mission, the attitude dynamics model of the flexible agile spacecraft is established considering multiple disturbances, including flexible vibration, unknown external disturbance torque and inertial parameter uncertainty;

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

[0078] S3: According to the attitude control model established in S1 and the full-state constraints established in S2, an asymmetric TVIBLF for handling attitude angle constraints and a constant IBLF for handling angular velocity constraints are constructed 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, a flexible agile spacecraft attitude controller under asymmetric full-state constraints when conditions are triggered is designed, and event-driven conditions are proposed.

[0080] S1 includes:

[0081] Firstly, according to the dynamic characteristics of the spacecraft attitude control task, the attitude dynamics model of the agile spacecraft under multiple disturbances is established as follows:

[0082]

[0083] Where:

[0084]

[0085] Where: is the orbital angular velocity, and a is the orbital semi-major axis. In addition, is the attitude angle vector, and ψ represent the roll angle, pitch angle and yaw angle of the agile spacecraft respectively; ω=[ω x ω y ω z ] T is the angular velocity vector of the spacecraft system relative to the inertial system. η∈R n is the n-dimensional modal coordinate vector, is the total modal velocity defined, F s ∈R 3×n is the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body, C = 2ζΩ, ζ and Ω∈R n×n is an n-dimensional diagonal matrix, which represents the damping ratio and modal frequency of the flexible structure. m =JF s F s T , J∈R 3×3 is the total attitude inertia matrix of the flexible spacecraft, J m is the moment of inertia matrix of the rigid body of the flexible spacecraft and J m =J m0 +△J, where J m0 is the nominal part of the moment of inertia of the rigid body of the flexible spacecraft, △J is the bounded unknown inertia matrix. u represents the control torque vector, μ=3.986012×10 5 km 3 / s2 is the gravitational parameter, sat u=[sat u x sat u y sat u z ] T is the control torque vector with saturation constraint, expressed as:

[0086]

[0087] Among them, u max 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, Δu = [Δu x Δu y Δu z ] T Indicates the saturation of the control torque. 1 =F s CΨ+F s Kη-s(ω)F s Ψ is the modal vibration related term, d represents the unknown external disturbance torque vector acting on the agile spacecraft, The interference is composed of unknown external interference torque and unknown inertia. The matrices A and B are expressed as follows:

[0088]

[0089] S2 includes:

[0090] In order to solve the influence of flexible vibration, inertial uncertainty and unknown external disturbance torque on flexible agile spacecraft, a modal observer and a multivariable nonlinear disturbance observer are designed to obtain the online observation values ​​of the above disturbances.

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

[0092]

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

[0094]

[0095] S3 includes:

[0096] In order to ensure the posture under asymmetric constraints, an asymmetric time-varying integral barrier Lyapunov function (TVIBLF) of the following form is designed:

[0097]

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

[0099]

[0100] Define the attitude angle error vector

[0101]

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

[0103] Introduce the following first-order filter

[0104]

[0105] where t ω =diag{t α t θ t ψ}, the error of the filter is

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

[0107]

[0108] in is the angular velocity error, is a continuously differentiable function, and there exists a positive constant satisfy k ωj is the angular velocity boundary. Define the angular velocity error

[0109]

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

[0111] S4 includes:

[0112] Design form of asymmetric stable control signal

[0113]

[0114] where c 1 is the positive design constant,

[0115]

[0116] β i >0 is a constant, and

[0117]

[0118] The flexible and agile spacecraft attitude controller under the asymmetric full-state constraint when the design condition is triggered is as follows:

[0119]

[0120] where c 2 is the positive design constant, is the design parameter matrix, also

[0121]

[0122]

[0123] Design event trigger mechanism for

[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 last trigger, where t∈[t k t k+1 ) represents the time between two trigger intervals, and b is a design parameter, γ>0.

[0126] Example:

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

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

[0129] 2) In order to solve the influence of flexible vibration, inertial uncertainty and unknown external disturbance torque on flexible agile spacecraft, a modal observer and a multivariable nonlinear disturbance observer are designed to observe them online;

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

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

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

[0133]

[0134] Where:

[0135]

[0136] Where: is the orbital angular velocity, a=729km is the orbital semi-major axis, and the gravitational constant μ=3.986012×10 5 km 3 / s 2 ,also, is 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 is the angular velocity vector of the spacecraft system relative to the inertial system, and its initial values ​​are Φ(0)=[305-20] T (deg), ω(0) = [3.5-1-3] T (deg / s).

[0137] η∈R n is the n-dimensional modal coordinate vector, is the total modal velocity defined, F s ∈R 3×n 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×nis an n-dimensional diagonal matrix, which represents 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:

[0138] [η 0 Ψ 0 ] T =0.001I 8

[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 is the total attitude inertia matrix of the flexible spacecraft, J m is the moment of inertia matrix of the rigid body of the flexible spacecraft, and J m =J m0 +△J is a symmetric positive definite matrix, where △J=0.1J m0 is the bounded unknown inertia matrix, J m0 is the nominal part of the moment of inertia of the rigid body of the flexible spacecraft, and its value is:

[0144]

[0145] sat u=[sat u x sat u y sat u z ] T is the control torque vector with saturation constraint, 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+△u, u is defined as the control vector, △u=[△u x △u y △u z ] T Indicates the saturation of the control torque.

[0148] d 1 =F s CΨ+F s Kη-s(ω)F s Ψ is the modal vibration related term; d represents the unknown external disturbance torque vector acting on the agile spacecraft,

[0149]

[0150] It is a composite interference composed of unknown external interference torque and unknown inertia. Matrices A and B are expressed as follows:

[0151]

[0152] 2) In order to solve the influence of flexible vibration, inertial uncertainty and unknown external disturbance torque on flexible agile spacecraft, a modal observer and a multivariable nonlinear disturbance observer are designed to obtain online observation values;

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

[0154]

[0155] in and are the observed values ​​of the modes η and Ψ, It is a composite disturbance composed of the uncertainty of the moment of inertia and the unknown external disturbance torque. 2 Observed value, P = 0.0001I 8 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, p(ω) = DJ m0 ω,

[0156]

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

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

[0159]

[0160] The positive design constant c 1 =0.5, c 2 =0.4, in addition and 2 =[ξ x ξ y ξ z ]T is an auxiliary variable.

[0161] In order to ensure the posture under asymmetric constraints, an asymmetric TVIBLF of the following form is designed:

[0162]

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

[0164]

[0165] where z i =ii d -ξ i is the attitude angle error, and the attitude angle error vector is defined as:

[0166]

[0167] in is the expected attitude. Assuming that the flexible agile spacecraft needs to maneuver three times in a short period of time, the expected attitudes of 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 error of the filter is

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

[0173]

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

[0175]

[0176] in 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 of the flexible and agile spacecraft under the asymmetric full-state constraint when the condition is triggered is as follows:

[0186]

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

[0188]

[0189] The event triggering mechanism is:

[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 last trigger, where t∈[t k t k+1 ) represents the time between two trigger intervals, and b = 0.05, γ = 0.005;

[0192] After numerical simulation verification, the attached Figure 1-Figure 7 , Figure 1 to Figure 2The agile satellite attitude and angular velocity curves under the event-driven mechanism. It can be seen that under the event-driven mechanism, the designed full-state constraint controller still has good constraint performance and attitude tracking performance, and the angular velocity change curve is always within the constant constraint range. Figures 3 to 5 It can be seen that under the action of the event-driven mechanism, the designed disturbance observer and modal observer still have good observation effects. Figure 3 In the above example, the observation accuracy of the disturbance observer for the composite disturbance is 0.015N. Figures 4 to 5 In the modal error observation diagram, the observation accuracy of modal error and defined modal error is 2×10 -3 With 3×10 -3 . Figure 6 It is the change curve of the control torque. It can be seen that when the trigger condition is not met, the change curve of the control torque is a short horizontal line, that is, the control torque is maintained at the last time the trigger condition is met. Figure 7 The time distribution of event updates and the sampling interval of event drive. The smallest sampling interval in the figure is 0.01s, and no Zeno phenomenon occurs. It can be seen that after the posture tracking is stable, the event drive time is mostly 0.15s~0.2s. The simulation time is 85s, the sampling interval is 0.01s, and a total of 498 triggers are triggered. Compared with the time-driven control updated every 0.01s, the communication resources are saved by about 94.13%, and it has better posture control performance.

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

Claims

1. An event-driven asymmetric attitude constraint control method for flexible and agile spacecraft, characterized in that: The steps include: Step 1: According to the dynamic characteristics of the spacecraft in the attitude control mission, the attitude dynamics model of the flexible agile spacecraft considering multiple disturbances is established, including flexible vibration, unknown external disturbance torque and inertial parameter uncertainty; Step 2: Design a modal observer and a disturbance observer, wherein the designed modal observer and disturbance observer respectively obtain online observation values ​​of vibration modes and online observation values ​​of composite disturbances composed of unknown external disturbance torque and inertia uncertainty; Step 3: Construct an asymmetric TVIBLF for handling attitude angle constraints and a constant IBLF for handling angular velocity constraints; Step 4: Design a flexible and 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 a flexible and agile spacecraft according to claim 1 is characterized in that: The step 1 is specifically as follows: According to the dynamic characteristics of the spacecraft attitude control task, the agile spacecraft attitude dynamics model under multiple interferences is established as follows: Where: Where: is the orbital angular velocity, a is the orbital semi-major axis; is 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 is the angular velocity vector of the spacecraft system relative to the inertial system; η∈R n is the n-dimensional modal coordinate vector, is the total modal velocity defined, F s ∈R 3×n is the coupling coefficient matrix between the vibration of the flexible structure and the rotation of the rigid body, C = 2ζΩ, ζ and Ω∈R n×n is an n-dimensional diagonal matrix, which represents the damping ratio and modal frequency of the flexible structure respectively; J m =JF s F s T , J∈R 3×3 is the total attitude inertia matrix of the flexible spacecraft, J m is the moment of inertia matrix of the rigid body of the flexible spacecraft and J m =J m0 +△J, where J m0 is the nominal part of the moment of inertia of the rigid body of the flexible spacecraft, △J is the bounded unknown inertia matrix; u represents the control torque vector, μ=3.986012×10 5 km 3 / s 2 is the gravitational parameter, sat u=[sat u x sat u y sat u z ] T is the control torque vector with saturation constraint, expressed as: Among them, u max is the maximum control torque provided by the actuator, satu is expressed as sat u=u+△u, u is defined as the control vector, △u=[△u x △u y △u z ] T Indicates the saturation of the control torque; d1 = F s CΨ+F s Kη-s(ω)F s Ψ is the modal vibration related term, d represents the unknown external disturbance torque vector acting on the agile spacecraft, The interference is composed of unknown external interference torque and unknown inertia. The matrices A and B are expressed as follows: Where K = Ω 2 .

3. The event-driven asymmetric attitude constraint control method for a flexible and agile spacecraft according to claim 2 is characterized in that: The step 2 is specifically as follows: Design modal observers and multivariable nonlinear disturbance observers to obtain online observations of disturbances; The structures of the modal observer and disturbance observer are as follows: in and are the observed values ​​of the modes η and Ψ, is the observed value of the composite disturbance d2 composed of the uncertainty of the moment of inertia and the unknown external disturbance torque, P∈R 2n×2n is the gain matrix of the modal observer, z is the auxiliary variable of the disturbance observer, D = diag[d x ,d y ,d z ]>0 is the gain matrix of the disturbance observer, p(ω)=DJ m0 ω, 4. The event-driven asymmetric attitude constraint control method for a flexible and agile spacecraft according to claim 3 is characterized in that: The step 3 is specifically as follows: Step 3-1: Design the asymmetric time-varying integral barrier Lyapunov function TVIBLF: where z i =ii d is the angular velocity error, k cai (t) and k cbi (t) is the upper and lower bounds of the attitude angle, δ i is the integral variable, i d is the desired attitude angle, Step 3-2: Define the attitude angle error vector; in is the desired attitude angle, is an auxiliary variable; Step 3-3: Introduce a first-order filter: where t ω =diag{t α t θ t ψ }, the error of the filter is Construct the following constant IBLF to implement the constraint on angular velocity: in is the angular velocity error, is a continuously differentiable function, and there exists a positive constant j=x,y,z satisfies j=x,y,z,k ωj is the angular velocity boundary; Define angular velocity error: in is an asymmetric virtual control vector, where ξ2=[ξ x ξ y ξ z ] T is an auxiliary variable.

5. The event-driven asymmetric attitude constraint control method for a flexible and agile spacecraft according to claim 4 is characterized in that: The S4 is specifically: Step 4-1: Design form of asymmetric stable control signal: where c1 is the positive design constant, β i >0 is a constant, Step 4-2: Design a flexible and agile spacecraft attitude controller under asymmetric full-state constraints when the condition is triggered, which is as follows: where c2 is the positive design constant, is the design parameter matrix, Step 4-3: Design the event trigger mechanism as follows: ||(u(t)-v(t))|| 2 ≥b||Z ω || 2 +γ (15) In the formula, u(t)=v(t k ) represents the control input at the last trigger, where t∈[t k t k+1 ) represents the time between two trigger intervals, b is the design parameter, and γ>0.

6. A computer program, characterized in that The computer program enables a computer to execute the method according to any one of claims 1 to 5.

7. 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, so that the electronic device executes the method as claimed in any one of claims 1 to 5.

8. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method according to any one of claims 1 to 5 is implemented.

9. A chip, characterized in that: include: A processor, configured to call and run a computer program from a memory, so that a device equipped with the chip executes a method as claimed in any one of claims 1 to 5.

10. A computer program product, characterized in that The computer program product comprises a computer storage medium storing a computer program, wherein the computer program comprises instructions executable by at least one processor, and when the instructions are executed by the at least one processor, the method as claimed in any one of claims 1 to 5 is implemented.

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