A bionic sequential control method for spacecraft based on disturbance immune inertia estimator

Abstract

The present invention provides a spacecraft bionic sequential control method based on an interference-immune inertia estimator, which belongs to the field of spacecraft control and is aimed at a spacecraft attitude control system under flexible vibration interference and inertia uncertainty. The method comprises: firstly, designing an induced angular velocity for a spacecraft composite interference dynamics model under flexible vibration interference and inertia uncertainty, and completing the design of an inertia identification control law based on rapid tracking of the induced angular velocity; secondly, designing an interference-immune inertia estimator including an interference dynamic separation estimation module and an inertia static solution module, and isolating the influence of the flexible vibration interference on the inertia estimation; finally, based on the output of the interference-immune inertia estimator, obtaining a spacecraft attitude dynamics correction equation, and completing the design of a spacecraft bionic sequential control law. The present invention has the advantages of high parameter estimation accuracy, strong anti-interference ability, etc., and can be used for high-precision attitude control of flexible spacecraft with uncertain inertia.

Description

Technical Field

[0001] The invention belongs to the field of spacecraft control, and in particular relates to a spacecraft bionic sequential control method based on a disturbance immune inertia estimator. Background Art

[0002] With the rapid development of aerospace technology, aerospace missions such as on-orbit grasping, laser communication and scientific experiments have become the direction of development for major aerospace powers. On the one hand, these complex aerospace missions put forward urgent demands on the attitude pointing and tracking accuracy of spacecraft, and high-precision, high-stability and strong robustness attitude control technology has become one of the key technologies to ensure the smooth implementation of aerospace missions. On the other hand, spacecraft often carry components such as space manipulators, large solar panels and large-scale deployable antennas, making the structure and operating environment of spacecraft increasingly complex. For example, the Japanese Engineering Test Satellite No. 8 carries two 19×2m solar panels and two 17×19m large antennas. In this case, "variable configuration" behaviors such as grasping non-cooperative targets and unfolding large space structures will cause drastic changes in the spacecraft inertia matrix. In order to meet the requirements of high-precision, high-stability and strong robustness attitude control, it is necessary to re-identify and estimate the inertia matrix of the spacecraft on-orbit.

[0003] However, spacecraft are faced with multiple sources of interference and uncertainty during their on-orbit operation. Inertia uncertainty is mixed with other interferences such as flexible vibration and environmental interference, and the identification of the inertia matrix is ​​easily "contaminated" by other interferences. Secondly, interferences such as inertia uncertainty and flexible vibration present complex characteristics. Inertia uncertainty and unknown angular acceleration present a static coupling relationship, while flexible vibration interference presents a dynamic coupling relationship with angular velocity, control input, and inertia uncertainty. In this complex interference environment, how to extract inertia uncertainty information and isolate the impact of other interferences on inertia identification is a challenging problem.

[0004] Existing estimation methods for spacecraft inertia uncertainty can be divided into two categories. One is an inertia estimation method based on adaptation or learning, which can estimate and identify the true value of the inertia matrix in real time. However, this type of method cannot separate the composite disturbance, and the estimation accuracy is easily affected by other disturbances such as flexible vibration; the other type of method is an estimation and control method based on observers (such as extended state observers, disturbance observers, etc.). This type of method can treat all disturbances such as inertia uncertainty and flexible vibration as a single "bounded rate" lumped disturbance. However, when the system has high-frequency dynamics, the performance of these observers is easily affected by the system dynamics.

[0005] The above methods have certain limitations when dealing with the problem of spacecraft inertia identification under complex interference. This is largely because the control signal is designed to ensure control performance and is not constructed from the perspective of signal identification and estimation. In nature, organisms have formed a strong ability to recognize uncertainty in behaviors such as foraging, hunting, and homing. For example, animals often use the sequence of "approach-hide-raid" to deal with various uncertainties and improve the success rate of hunting. Therefore, it is urgent to empower the spacecraft attitude control system with "bionic intelligent sequential behavior". In the first sequence, the control input is constructed from the perspective of signal identification and estimation, the inertia matrix is ​​identified, and the influence of interference such as flexible vibration is isolated, that is, the "interference immune" inertia matrix estimation is realized. In the second sequence, the control input is designed from the perspective of attitude control performance, and the identification value of the inertia matrix is ​​incorporated into the design of the attitude control law to achieve accurate, robust and fast attitude control. Summary of the invention

[0006] In view of the problem that existing spacecraft attitude control methods are difficult to achieve accurate estimation of the inertia matrix and ensure high-precision and high-stability attitude control in a complex interference environment, the present invention provides a spacecraft bionic sequential control method based on an interference-immune inertia estimator. Inspired by the biological "bionic intelligent sequential behavior", in the first order, an inertia identification control law and an inertia matrix estimation method with interference immunity are designed. In the second order, according to the estimated value of the inertia matrix, an attitude maneuver control law based on a sliding mode observer is designed, thereby completing the design of the bionic sequential control law, providing theoretical and technical support for space missions such as on-orbit services, space / earth observations, and precision scientific experiments.

[0007] In order to achieve the above object, the present invention adopts the following technical scheme:

[0008] A method for bionic sequential control of a spacecraft based on a disturbance immune inertia estimator comprises the following steps:

[0009] Step 1: Design the induced angular velocity for the spacecraft composite disturbance dynamics model under flexible vibration disturbance and inertia uncertainty, and complete the design of the inertia identification control law based on the rapid tracking of the induced angular velocity;

[0010] Step 2: Design a disturbance immune inertia estimator including a disturbance dynamic separation estimation module and an inertia static solution module to isolate the influence of flexible vibration disturbance on inertia estimation;

[0011] Step 3: Based on the output of the disturbance immune inertia estimator, the spacecraft attitude dynamics correction equation is obtained to complete the design of the spacecraft bionic sequential control law.

[0012] The beneficial effects of the present invention compared with the prior art are:

[0013] The present invention makes full use of the deep coupling relationship of the spacecraft's composite interference and designs an inertia estimator with interference immunity, overcoming the difficult problem that inertia estimation is easily affected by interference such as flexible vibration. In addition, a spacecraft bionic sequential control method is invented, which separates the inertia identification control law and the attitude maneuver control law in chronological order, which can significantly improve the accuracy of the spacecraft's inertia matrix identification and improve the accuracy, stability and robustness of the attitude control system. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] Figure 1 A flow chart of a method for bionic sequential control of a spacecraft based on a disturbance immune inertia estimator according to the present invention;

[0015] Figure 2 is the time response curve of the flexible vibration disturbance and its estimated value;

[0016] Figure 3 is the time response curve of inertia uncertainty and its estimated value;

[0017] Figure 4 is the time response curve of the spacecraft angular velocity and its estimated value;

[0018] Figure 5 This is the time response curve of the spacecraft angular velocity. DETAILED DESCRIPTION

[0019] The present invention is further described in detail below with reference to the accompanying drawings and embodiments.

[0020] like Figure 1 As shown, a method for spacecraft bionic sequential control based on a disturbance immune inertia estimator of the present invention comprises the following steps:

[0021] The first step is to design the induced angular velocity for the spacecraft composite disturbance dynamics model under flexible vibration disturbance and inertia uncertainty, and complete the design of the inertia identification control law based on the rapid tracking of the induced angular velocity, including:

[0022] The spacecraft attitude dynamics equation is:

[0023] ;

[0024] in, is the spacecraft inertia matrix, is the spacecraft angular velocity, , and They represent the roll angular velocity, pitch angular velocity, and yaw angular velocity, respectively. The superscript “T” indicates the transpose operation of the matrix; for The first time derivative of Indicates about The cross product matrix of represents the rigid-flex coupling matrix, represents the bionic sequential control law of the spacecraft, represents the environmental disturbance torque; represents the modal displacement of the flexible component, and Respectively The first time derivative of The second-order time derivative of and Respectively represent the damping matrix and stiffness matrix of the flexible component; for the actual spacecraft system, there is a known positive number , , and , so that the inequality , , and Established, symbol Indicates the calculation of the matrix bi-norm;

[0025] definition is the flexible vibration interference, is the inertia uncertainty, is a given positive number, represents a unit matrix with 3 rows and 3 columns; at this time, the composite disturbance dynamics model of the spacecraft under flexible vibration disturbance and inertia uncertainty is obtained as follows:

[0026] ;

[0027] Design the induced angular velocity for the spacecraft composite disturbance dynamics model is in the form of a sine function, that is ;in, represents the time variable, represents the frequency of the induced angular velocity, Indicates the amplitude of the induced angular velocity; the amplitude of the induced angular velocity The design is in the form of a piecewise constant vector:

[0028] ;

[0029] in, , , , , , , , and are given constant values, , and All are given moments;

[0030] In the inertia identification phase (i.e., time variable satisfy ), design the inertia identification control law based on the rapid tracking of the induced angular velocity, and make the spacecraft bionic sequential control law is equal to the inertia identification control law; specifically, the inertia identification control law Designed for:

[0031] ;

[0032] in, represents the angular velocity tracking error, , , and Identify the control law for inertia The positive gain satisfies the following relationship:

[0033] ;

[0034] in, Induced angular velocity The first time derivative of is a given positive number.

[0035] The second step is to design a disturbance immune inertia estimator that includes a disturbance dynamic separation estimation module and an inertia static solution module to isolate the impact of flexible vibration disturbance on inertia estimation, including:

[0036] In the inertia identification stage, the design includes a disturbance dynamic separation estimation module Static inertia solver The disturbance immune inertia estimator is:

[0037] : ;

[0038] : ;

[0039] in, Indicates flexible vibration interference The internal state variables of Inertia uncertainty The internal state variables of and Respectively The estimated value and An estimated value of and They represent the flexible vibration disturbance The estimated value and inertia uncertainty of The estimated value of represents the estimated value of the spacecraft inertia matrix, Dynamic separation and estimation module for interference The gain matrix of is an auxiliary variable, for The first-order time derivative of , the superscript “-1” indicates the inverse operation of the matrix; , , , and They are all known coefficient matrices, and their expressions are: , , , , ; , and Respectively OK The zero matrix of columns, A zero matrix with 3 rows and 3 columns and a zero matrix with 3 rows and 3 columns, Represents the modal displacement of the flexible component The dimension of and Represents 6 lines respectively A zero matrix with 6 rows and 3 columns, express OK The identity matrix of columns; Indicates from time To a given time The integral value of Indicates from time To a given time The integral value of Indicates from time To a given time The integral value of

[0040] Selecting interference dynamic separation estimation module by pole placement method The gain matrix , so that the matrix All eigenvalues ​​of have negative real parts, thus isolating the effect of the flexible vibration disturbance on the inertia estimate.

[0041] The third step is to obtain the spacecraft attitude dynamics correction equation based on the output of the disturbance immune inertia estimator and complete the design of the spacecraft bionic sequential control law, including:

[0042] In the posture control phase (i.e., time variable satisfy ), based on the output of the disturbance immune inertia estimator, the spacecraft attitude dynamics correction equation is obtained as:

[0043] ;

[0044] in, represents the spacecraft inertia correction matrix, ( and ) represents the estimated value of the spacecraft inertia matrix No. Line Elements of a column; represents the spacecraft inertia correction error matrix;

[0045] Based on the spacecraft attitude dynamics correction equation, the sliding mode observer is designed as:

[0046] ;

[0047] in, express The estimated value of express The first time derivative of and is the positive gain of the sliding mode observer, is a given constant; in the attitude control stage, the attitude maneuver control law is designed , and make the spacecraft bionic order control law Attitude maneuver control law Equal; specifically, the attitude maneuver control law Designed for:

[0048] ;

[0049] in, is the attitude maneuver control law Positive gain of

[0050] Bionic Sequential Control Law for Spacecraft Identifying the control law by inertia and attitude maneuver control law In chronological order, the expression is:

[0051] ;

[0052] The spacecraft bionic sequential control law has different forms in the inertia identification stage and the attitude control stage, thereby achieving the goals of high-precision identification of the inertia matrix and rapid attitude maneuvering respectively.

[0053] Example:

[0054] This embodiment takes the spacecraft attitude control system as an example, and the parameters of the spacecraft are shown in Table 1:

[0055] Table 1 Spacecraft parameters

[0056]

[0057] The steps of this embodiment include:

[0058] The first step is to design the induced angular velocity for the spacecraft composite disturbance dynamics model under flexible vibration disturbance and inertia uncertainty, and complete the design of the inertia identification control law based on the rapid tracking of the induced angular velocity.

[0059] In this embodiment, the known positive number , , and The values ​​are 50, 0.5, 0.5 and 0.01 respectively, and the given positive numbers The value is 3, the given constant value , , , , , , , and The values ​​are 0.005, 0.01, 0.0075, 0.006, 0.0075, 0.0065, 0.004, 0.0035 and 0.005 respectively, and the frequency of the induced angular velocity The value is 2rad / s, given the time , and The values ​​are 25.13s, 50.26s and 75.40s respectively; the given positive numbers The value is 2, inertia identification control law Positive gain , , and The values ​​are 10, 6.17, 0.80 and 50 respectively.

[0060] The second step is to design a disturbance-immune inertia estimator that includes a disturbance dynamic separation estimation module and an inertia static solution module to isolate the impact of flexible vibration disturbance on inertia estimation.

[0061] In this embodiment, the interference dynamic separation estimation module The gain matrix Select as:

[0062] ;

[0063] In order to verify the effectiveness of the designed disturbance immune inertia estimator, Figure 2 The time response curve of the flexible vibration disturbance and its estimated value are given. Figure 3 The time response curve of inertia uncertainty and its estimated value is given; Figure 2 and Figure 3 It can be seen that the disturbance dynamic separation estimation module in the disturbance immune inertia estimator It can realize the accurate separation and estimation of flexible vibration disturbance and inertia uncertainty; further, the inertia static solution module in the disturbance immune inertia estimator Get the estimated value of the spacecraft inertia matrix , and then the spacecraft inertia correction matrix is ​​obtained as ; By comparing the spacecraft inertia matrix And the spacecraft inertia correction matrix is It can be seen that the designed interference-immune inertia estimator achieves accurate estimation of the spacecraft inertia matrix.

[0064] The third step is to obtain the spacecraft attitude dynamics correction equation based on the output of the interference immune inertia estimator, and complete the design of the spacecraft bionic sequential control law.

[0065] In this embodiment, the positive gain of the sliding mode observer is and The values ​​are 1 and 10 respectively, and the given constant The value is 0.001, attitude maneuver control law Positive gain The value is 1;

[0066] To verify the attitude control phase (i.e. the time variable satisfies s) the control performance of the designed method, Figure 4 The time response curve of the spacecraft angular velocity and its estimated value is given; Figure 4 It can be seen that under the designed sliding mode observer, the estimated value of the spacecraft angular velocity can accurately track its true value, and the estimation error is less than rad / s; Figure 5 The time response curve of the spacecraft angular velocity is given by Figure 5 It can be seen that the rolling angular velocity of the spacecraft is , pitch angular velocity and yaw rate They can converge quickly to near zero, and the steady-state error is less than rad / s, which fully verifies the effectiveness of the designed bionic sequential control law for the spacecraft.

[0067] The contents not described in detail in this specification belong to the prior art known to professional and technical personnel in this field.

Claims

1. A method for bionic sequential control of spacecraft based on a disturbance immune inertia estimator, characterized in that: The steps include: Step 1: Design the induced angular velocity for the spacecraft composite disturbance dynamics model under flexible vibration disturbance and inertia uncertainty, and complete the design of the inertia identification control law based on the rapid tracking of the induced angular velocity; Step 2: Design a disturbance immune inertia estimator including a disturbance dynamic separation estimation module and an inertia static solution module to isolate the influence of flexible vibration disturbance on inertia estimation; Step 3: Based on the output of the disturbance immune inertia estimator, the spacecraft attitude dynamics correction equation is obtained to complete the design of the spacecraft bionic sequential control law; The step 2 comprises: In the inertia identification stage, the disturbance dynamic separation estimation module is denoted as , inertia static solution module is , the design includes interference dynamic separation estimation module Inertia Static Solver The disturbance immune inertia estimator is: ： ； ： ； in, Indicates flexible vibration interference The internal state variables of Inertia uncertainty The internal state variables of and They represent the flexible vibration disturbance The internal state variables The estimated value and inertia uncertainty of The internal state variables An estimated value of and They represent the flexible vibration disturbance The estimated value and inertia uncertainty of The estimated value of represents the estimated value of the spacecraft inertia matrix, Dynamic separation and estimation module for interference The gain matrix of is an auxiliary variable, for The first time derivative of ; , , , and All are known coefficient matrices; and Represents 6 lines respectively A zero matrix with 6 rows and 3 columns, Represents the modal displacement of the flexible component The dimension of , , , , , , , and are given constant values, , and For a given moment, represents the frequency of the induced angular velocity; Indicates from time To a given time The integral value of Indicates from time To a given time The integral value of Indicates from time To a given time The integral value of; the superscript "-1" indicates the inverse operation of the matrix; then, the interference dynamic separation estimation module is selected by the pole placement method The gain matrix , isolating the effect of flexible vibration disturbance on inertia estimation.

2. The method for spacecraft bionic sequential control based on a disturbance immune inertia estimator according to claim 1 is characterized in that: The step 1 comprises: Based on the spacecraft attitude dynamics equation, the spacecraft composite disturbance dynamics model under flexible vibration disturbance and inertia uncertainty is obtained as follows: ； in, is the spacecraft inertia matrix, is the spacecraft angular velocity, for The first time derivative of Indicates about The cross product matrix of represents the rigid-flex coupling matrix, and the superscript "T" represents the transpose operation of the matrix; represents the bionic sequential control law of the spacecraft, represents the environmental disturbance torque; represents the modal displacement of the flexible component, express The first time derivative of ; is the flexible vibration interference, is the inertia uncertainty, is a given positive number, Represents the identity matrix with 3 rows and 3 columns; According to the spacecraft composite disturbance dynamics model, the induced angular velocity is designed as a sinusoidal function, and the amplitude of the induced angular velocity is designed as a piecewise constant vector; in the inertia identification stage, an inertia identification control law based on the rapid tracking of the induced angular velocity is designed. , and make the spacecraft bionic order control law Equivalent to the inertia identification control law; Inertia identification control law Designed for: ； in, represents the angular velocity tracking error, represents the induced angular velocity, , , and Identify the control law for inertia The positive gain, symbol It means to find the matrix bi-norm.

3. The method for spacecraft bionic sequential control based on a disturbance immune inertia estimator according to claim 1 is characterized in that: The step 3 comprises: In the attitude control stage, the spacecraft attitude dynamics correction equation is obtained based on the output of the disturbance immune inertia estimator; based on the spacecraft attitude dynamics correction equation, the sliding mode observer is designed as follows: ； in, represents the spacecraft inertia correction matrix, represents the angular velocity of the spacecraft The estimated value of express The first time derivative of and is the positive gain of the sliding mode observer, is a given constant; In the attitude control stage, design the attitude maneuver control law , and make the spacecraft bionic order control law and attitude maneuver control law Equal; attitude maneuver control law Designed for: ； in, is the attitude maneuver control law Positive gain of Bionic Sequential Control Law for Spacecraft Identify the control law by inertia and attitude maneuver control law The spacecraft bionic order control law is constructed in time sequence; It has different forms in the inertia identification stage and the attitude control stage, thereby achieving the goals of high-precision identification of the inertia matrix and rapid attitude maneuvering respectively.

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