A low earth orbit satellite parameter adaptive control method with a solar sail drive device

Abstract

The application discloses a low-orbit satellite parameter adaptive control method with a solar sail driving device, belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The method solves the technical problem of overcoming the influence of satellite moment of inertia and compensating the influence of SADA disturbance torque and atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in the form of a quaternion, and on the basis, constructing a satellite attitude kinematics and dynamics equation for attitude controller design; defining the error attitude kinematics and dynamics equation of the satellite through an attitude control error and an angle control error; dividing and converting the parameters to be estimated and known state variables, and representing the negative influence of moment of inertia error on the attitude by matrix operation; designing an attitude controller control torque in a manner of combining nonlinear feedback control with adaptive control; and updating the adaptive estimation value in the attitude controller control parameter by using a parameter adaptive law.

Description

Technical Field

[0001] This invention belongs to the field of aerospace technology, specifically relating to the field of low-Earth orbit satellite parameter control technology. Background Technology

[0002] With the rapid development of aerospace technology, the applications of communication, navigation, and remote sensing satellites are becoming increasingly widespread. To improve satellite efficiency, many satellites are equipped with Solar Actuation and Control (SADA) systems, which enable the satellite to control its solar panels to always point towards the sun under different orbits and attitudes, ensuring energy balance and enhancing mission adaptability. However, because SADA drives the solar panels to rotate at different angles, the satellite's moment of inertia changes in real time. Simultaneously, SADA generates reaction forces during acceleration and deceleration, significantly impacting the satellite's attitude stability. Furthermore, due to increased solar activity, satellites in low Earth orbit also face significant atmospheric drag torque, posing a substantial challenge to achieving high-precision pointing control. Summary of the Invention

[0003] To address the technical challenges of overcoming the influence of satellite rotational inertia and compensating for the impact of SADA (Satellite Adaptive Damping and Agitation) disturbance torque and atmospheric drag torque on satellite attitude control accuracy, this invention provides a low-Earth orbit satellite parameter adaptive control method with a solar panel drive mechanism. Through dynamic parameter adaptive estimation, the satellite's rotational inertia parameters and disturbance torque can be estimated in real time, and compensation is achieved through nonlinear control. Ultimately, this allows the satellite to balance mission execution efficiency and control accuracy.

[0004] The method of the present invention includes the following steps:

[0005] S1. Perform satellite attitude kinematics and dynamics modeling: Construct satellite attitude kinematics and dynamics equations in quaternion form, and on this basis, construct satellite attitude kinematics and dynamics equations for attitude controller design;

[0006] S2. Model the error attitude kinematics and dynamics equations: Define the error attitude kinematics and dynamics equations of the satellite by using attitude control error and angle control error;

[0007] S3. Perform rotational inertia error matrix transformation: Define a transformation function to separate and transform the parameters to be estimated with the known state variables, and represent the negative impact of rotational inertia error on attitude using matrix operations;

[0008] S4. Perform attitude controller design: Design the attitude controller control torque using a combination of nonlinear feedback control and adaptive control.

[0009] S5. Design the adaptive law for attitude controller parameters: Update the adaptive estimates in the attitude controller control parameters using the adaptive law.

[0010] Furthermore, the quaternion form of the satellite attitude kinematics and dynamics equations are constructed as follows:

[0011] Using attitude quaternions To represent the satellite's body coordinate system Relative to the inertial coordinate system The posture, among which Representing the scalar part of a quaternion. The vector part of a quaternion. express The transpose of , and satisfies the constraints. ;

[0012] The quaternion form of the satellite attitude kinematics equations is: ;

[0013] The quaternion form of the satellite attitude dynamics equations is: ;

[0014] in, The matrix representing the satellite's moment of inertia; Indicates the satellite's attitude angular velocity; This represents the control angular momentum of the reaction flywheel; This represents the sum of the vibration disturbance torque and the atmospheric drag torque of the SADA. This indicates the control torque of the reaction flywheel; They represent The differential; Representing vectors The antisymmetric matrix; express identity matrix; express The differential; express antisymmetric matrix, subscript These represent the satellite's rotation axes.

[0015] Furthermore, the specific kinematic equations for satellite attitude control design are as follows:

[0016] The kinematic equations for satellite attitude control design are:

[0017] ;

[0018] The satellite attitude dynamics equations used for attitude controller design are:

[0019] ;in, The matrix representing the measured moment of inertia of the satellite. .

[0020] Furthermore, the satellite is defined in the inertial coordinate system. Expected posture quaternion Attitude control error is expressed by error quaternion. express, , Represents the expected quaternion The reverse, Represents quaternion multiplication; , Representing error quaternions The scalar part, Representing error quaternions Vector part,

[0021] Define the satellite in the inertial coordinate system Expected angular velocity Angular velocity control error is achieved through error angular velocity. express, ; The rotation transformation matrix representing the error quaternion. ; Representing vectors An antisymmetric matrix.

[0022] Furthermore, the satellite's error attitude kinematics and dynamics equations are as follows:

[0023] The error attitude kinematic equations of the satellite: Satellite error attitude dynamics equations:

[0024] .

[0025] Furthermore, step S3 specifically involves defining the transformation function. Given vector The following relationship exists:

[0026] ;

[0027] The parameters to be estimated are separated and transformed from the known state variables, that is:

[0028] ,

[0029] ,

[0030] ,

[0031] ;

[0032] Finally, the negative impact of rotational inertia error on attitude was expressed using... express.

[0033] Furthermore, the control torque of the attitude controller pass:

[0034]

[0035] get;

[0036] Indicates proportional control gain; Indicates the differential control gain; Indicating nonlinear compensation control: express The adaptive estimate is used to compensate for the influence of external disturbance torque; express An adaptive estimate is used to compensate for the time-varying effects of the moment of inertia.

[0037] Furthermore, step S5 specifically includes:

[0038] right The update was successful using the parameter adaptive law: To carry out, on The update was successful using the parameter adaptive law: conduct, Represents the update gain of the adaptive law; Represents the estimated parameters The derivative; This represents the correction gain of the adaptive law. This represents the error angular velocity loop control variable. .

[0039] The beneficial effects of the method described in this invention are as follows:

[0040] (1) It can estimate the time-varying parameters of rotational inertia and external disturbance torque parameters in real time through parameter adaptive estimation, and reconstruct the controller to compensate for the influence of inertia parameter error and disturbance torque on attitude pointing accuracy and attitude stability.

[0041] (2) It ensured the attitude control accuracy of low-orbit satellites with solar panel drive devices, providing a technical foundation for the satellite to successfully perform stable communication tasks and high-quality remote sensing imaging tasks.

[0042] (3) The designed method is simple to control and has high economic value. It is applicable to the design of attitude control systems for low-orbit satellites, especially high-precision optical remote sensing satellites, in engineering practice. Attached Figure Description

[0043] Figure 1 This is a diagram illustrating the control method architecture in an embodiment of the present invention.

[0044] Figure 2 This is a diagram of attitude quaternion curves in an embodiment of the present invention;

[0045] Figure 3 This is a diagram showing the attitude angular velocity curve in an embodiment of the present invention;

[0046] Figure 4 This is a graph showing the error quaternion curve in an embodiment of the present invention;

[0047] Figure 5 This is an error angular velocity curve diagram in an embodiment of the present invention;

[0048] Figure 6 This is a graph showing the estimation curve of the disturbance torque parameter in an embodiment of the present invention;

[0049] Figure 7 This is a graph showing the estimation curve of the rotational inertia parameter in an embodiment of the present invention. Detailed Implementation

[0050] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the protection scope of the present invention.

[0051] The overall architecture of the method described in this invention is as follows: Figure 1 As shown, nonlinear feedback control is added to the basic feedback control method to enhance the tracking performance of attitude control. Simultaneously, adaptive parameter compensation control is introduced to compensate for the effects of time-varying rotational inertia parameters and external disturbance torques on attitude control accuracy.

[0052] This invention patent is achieved through the following technical solution. This method for adaptive control of low-Earth orbit satellite parameters with a solar panel drive device includes the following steps:

[0053] Step 1, Satellite attitude kinematics and dynamics modeling:

[0054] The attitude control reference for satellites performing space missions in orbit is the geocentric inertial coordinate system. and the body coordinate system To describe the satellite's attitude changes in all dimensions without singularity, attitude quaternions are used. To represent the satellite's body coordinate system Relative to the inertial coordinate system The posture, among which Representing the scalar part of a quaternion. The vector part of a quaternion. express The transpose of , and satisfies the constraints. The quaternion form of the satellite attitude kinematics and dynamics equations is as follows:

[0055] The quaternion form of the satellite attitude kinematics equations is: ; The quaternion form of the satellite attitude dynamics equations is: ;

[0056] in: The matrix representing the satellite's moment of inertia; Indicates the satellite's attitude angular velocity; This represents the control angular momentum of the reaction flywheel; This represents the sum of the vibration disturbance torque and the atmospheric drag torque of the SADA. This indicates the control torque of the reaction flywheel; They represent The differential; Representing vectors The antisymmetric matrix; express identity matrix; express The differential; express The antisymmetric matrix. Wherein Specifically:

[0057] ;

[0058] The satellite's true moment of inertia is subscript Let represent the satellite's rotation axes, and define its measured moment of inertia as . Then the error between the measured moment of inertia and the actual moment of inertia is... Since only the measured moment of inertia can be obtained in practical engineering, the kinematic equations of satellite attitude for controller design can be described as follows:

[0059] ; ;

[0060] Simplifying the above equation, we get:

[0061] ; ;

[0062] Step 2, Modeling the kinematics and dynamics equations of error attitude:

[0063] Define the satellite in the inertial coordinate system Expected posture quaternion Then the attitude control error can be expressed by the error quaternion. The calculation formula is as follows:

[0064] ;

[0065] in: Represents the expected quaternion The reverse, This represents quaternion multiplication.

[0066] Define the satellite in the inertial coordinate system Expected angular velocity Then the angular velocity control error can be controlled by the error angular velocity. The calculation formula is as follows:

[0067] ;

[0068] in The rotation transformation matrix of a quaternion is calculated using the following formula:

[0069] ;

[0070] in: Representing vectors An antisymmetric matrix.

[0071] Based on the preceding text, the error attitude kinematics equations of the satellite are as follows:

[0072] ;

[0073] .

[0074] Step 3, Transformation of the moment of inertia error matrix:

[0075] To design a parameter adaptive controller, a new transformation function is defined. If given vector The following relationship exists:

[0076] ;

[0077] The parameters to be estimated are separated and transformed from the known state variables, that is:

[0078]

[0079] in:

[0080] ; ; ;

[0081] Ultimately, the negative impact of rotational inertia error on attitude can be attributed to matrix operations. .

[0082] Step 4, Attitude Controller Design:

[0083] To ensure the excellent performance of the attitude tracking controller, a combination of nonlinear feedback control and adaptive control is used in the controller design. The specific form of the attitude controller is as follows:

[0084] ;

[0085] in: This indicates that the control torque calculated by the controller is sent to the reaction flywheel (actuator).

[0086] Indicates proportional control gain; Indicates the differential control gain; Indicating nonlinear compensation control: express The adaptive estimate is used to compensate for the influence of external disturbance torque; express An adaptive estimate is used to compensate for the time-varying effects of the moment of inertia.

[0087] Step 5, Parameter Adaptive Law Design:

[0088] To dynamically compensate for the effects of external disturbance torques and time-varying moments of inertia, real-time calculations are required. and This enables real-time compensation for factors affecting attitude control accuracy. and The parameter adaptive law is used for updating the estimate, specifically in the form of: The update was successful using the parameter adaptive law: To carry out, on The update was successful using the parameter adaptive law: conduct, Represents the update gain of the adaptive law; Represents the estimated parameters The derivative; This represents the correction gain of the adaptive law. This represents the error angular velocity loop control variable. .in: Represents the update gain of the adaptive law; Represents the estimated parameters The derivative; This represents the correction gain of the adaptive law.

[0089] This embodiment illustrates the attitude control performance of the proposed method using a scenario involving two attitude maneuvering control processes plus an imaging process. The analysis process lasts for 200 seconds, with 0s-50s representing the initial attitude stabilization process, 50s representing the start of attitude maneuvering, and 80s-120s representing the first ground imaging process. After 120s, a second attitude maneuver is performed, and 160s-200s represent the second ground imaging process. The attitude quaternion control curve is shown below. Figure 2 As shown in the figure, the actual attitude curve almost perfectly matches the desired attitude curve, demonstrating excellent attitude tracking control performance. The attitude angular velocity control curve is shown below. Figure 3 As shown, the attitude maneuver angular velocity curve approximates a triangle, indicating rapid convergence during the attitude maneuver. The attitude quaternion control error and angular velocity control error are as follows: Figures 4-5 As shown, the satellite's attitude control error was very small throughout the entire process, and the control error converged quickly. The estimated parameters for the disturbance torque and moment of inertia are as follows: Figures 6-7 As shown, it rapidly and adaptively adjusts the parameter estimates as the satellite's state changes, ensuring accurate control precision.

[0090] The relevant parameters for the embodiments are shown in Table 1.

[0091] Table 1:

[0092]

[0093] The memory in this application embodiment can be volatile memory or non-volatile memory, or it can include both volatile and non-volatile memory. The non-volatile memory can be read-only memory (ROM), programmable read-only memory (PROM), erasable programmable read-only memory (EPROM), electrically erasable programmable read-only memory (EEPROM), or flash memory. The volatile memory can be random access memory (RAM), which is used as an external cache. By way of example, but not limitation, many forms of RAM are available, such as static random access memory (SRAM), dynamic random access memory (DRAM), synchronous dynamic random access memory (SDRAM), double data rate synchronous dynamic random access memory (DDRSDRAM), enhanced synchronous dynamic random access memory (ESDRAM), synchronous linked dynamic random access memory (SLDRAM), and direct rambus RAM (DR RAM). It should be noted that the memory used in the methods described in this invention is intended to include, but is not limited to, these and any other suitable types of memory.

[0094] In the above embodiments, implementation can be achieved, in whole or in part, through software, hardware, firmware, or any combination thereof. When implemented in software, it can be implemented, in whole or in part, as a computer program product. The computer program product includes one or more computer instructions. When the computer instructions are loaded and executed on a computer, all or part of the processes or functions described in the embodiments of this application are generated. The computer can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions can be stored in a computer-readable storage medium or transmitted from one computer-readable storage medium to another. For example, the computer instructions can be transmitted from one website, computer, server, or data center to another via wired (e.g., coaxial cable, fiber optic, digital subscriber line (DSL)) or wireless (e.g., infrared, wireless, microwave, etc.) means. The computer-readable storage medium can be any available medium accessible to a computer or a data storage device such as a server or data center that integrates one or more available media. The available media may be magnetic media (e.g., floppy disks, hard disks, magnetic tapes), optical media (e.g., high-density digital video discs (DVDs)), or semiconductor media (e.g., solid-state disks (SSDs)).

[0095] In implementation, each step of the above method can be completed by integrated logic circuits in the processor's hardware or by instructions in software. The steps of the method disclosed in the embodiments of this application can be directly implemented by a hardware processor, or by a combination of hardware and software modules in the processor. The software modules can reside in random access memory, flash memory, read-only memory, programmable read-only memory, electrically erasable programmable memory, registers, or other mature storage media in the art. This storage medium is located in memory, and the processor reads information from the memory and, in conjunction with its hardware, completes the steps of the above method. To avoid repetition, detailed descriptions are omitted here.

[0096] It should be noted that the processor in the embodiments of this application can be an integrated circuit chip with signal processing capabilities. During implementation, each step of the above method embodiments can be completed by the integrated logic circuitry in the processor's hardware or by instructions in software form. The processor can be a general-purpose processor, a digital signal processor (DSP), an application-specific integrated circuit (ASIC), a field-programmable gate array (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, or discrete hardware components. It can implement or execute the methods, steps, and logic block diagrams disclosed in the embodiments of this application. The general-purpose processor can be a microprocessor or any conventional processor. The steps of the methods disclosed in the embodiments of this application can be directly embodied as execution by a hardware decoding processor, or as a combination of hardware and software modules in the decoding processor. The software modules can be located in random access memory, flash memory, read-only memory, programmable read-only memory, electrically erasable programmable memory, registers, or other mature storage media in the art. This storage medium is located in memory, and the processor reads the information in the memory and, in conjunction with its hardware, completes the steps of the above methods.

Claims

1. A method for adaptive control of low-Earth orbit satellite parameters with a solar panel drive device, characterized in that, The method includes the following steps: S1. Perform satellite attitude kinematics and dynamics modeling: Construct satellite attitude kinematics and dynamics equations in quaternion form, and on this basis, construct satellite attitude kinematics and dynamics equations for attitude controller design; The quaternion form of the satellite attitude kinematics and dynamics equations are constructed as follows: Using attitude quaternions To represent the satellite's body coordinate system Relative to the inertial coordinate system The posture, among which Representing the scalar part of a quaternion. The vector part of a quaternion. express The transpose of , and satisfies the constraints. ; The quaternion form of the satellite attitude kinematics equations is: ; The quaternion form of the satellite attitude dynamics equations is: ; in, The matrix representing the satellite's moment of inertia; Indicates the satellite's attitude angular velocity; This represents the control angular momentum of the reaction flywheel; This represents the sum of the vibration disturbance torque and the atmospheric drag torque of the SADA. This indicates the control torque of the reaction flywheel; They represent The differential; Representing vectors The antisymmetric matrix; express identity matrix; express The differential; express antisymmetric matrix, subscript These represent the satellite's rotation axes; S2. Model the error attitude kinematics and dynamics equations: Define the error attitude kinematics and dynamics equations of the satellite by using attitude control error and angle control error; S3. Perform rotational inertia error matrix transformation: Define a transformation function to separate and transform the parameters to be estimated from the known state variables, and represent the negative impact of rotational inertia error on attitude using matrix operations, specifically: Define transformation function Given vector The following relationship exists: ； The parameters to be estimated are separated and transformed from the known state variables, that is: ,in, Indicates the error angular velocity. The rotation transformation matrix representing the error quaternion. Indicates the satellite in the inertial coordinate system The expected angular velocity, ， ， ； Finally, the negative impact of rotational inertia error on attitude was expressed using... express; , The matrix representing the measured moment of inertia of the satellite; S4. Perform attitude controller design: Design the attitude controller control torque using a combination of nonlinear feedback control and adaptive control. S5. Design the adaptive law for attitude controller parameters: Update the adaptive estimates in the attitude controller control parameters using the adaptive law.

2. The low-Earth orbit satellite parameter adaptive control method with solar panel drive device according to claim 1, characterized in that, The specific kinematic equations for satellite attitude kinematics used in attitude controller design are as follows: The kinematic equations for satellite attitude control design are: ； The satellite attitude dynamics equations used for attitude controller design are: 。 3. The low-orbit satellite parameter adaptive control method with solar panel drive device according to claim 2, characterized in that, Define the satellite in the inertial coordinate system Expected posture quaternion Attitude control error is expressed by error quaternion. express, , Represents the expected quaternion The reverse, Represents quaternion multiplication; , Representing error quaternions The scalar part, Representing error quaternions Vector part, Define the satellite in the inertial coordinate system Expected angular velocity Angular velocity control error is achieved through error angular velocity. express, ; ; Representing vectors An antisymmetric matrix.

4. The low-Earth orbit satellite parameter adaptive control method with solar panel drive device according to claim 3, characterized in that, The specific equations of motion for the satellite's attitude error are as follows: The error attitude kinematic equations of the satellite: ; Satellite error attitude dynamics equations: 。 5. The low-Earth orbit satellite parameter adaptive control method with solar panel drive device according to claim 4, characterized in that, Control torque of attitude controller pass: get; Indicates proportional control gain; Indicates the differential control gain; Indicating nonlinear compensation control: express The adaptive estimate is used to compensate for the influence of external disturbance torque; express An adaptive estimate is used to compensate for the time-varying effects of the moment of inertia.

6. The low-Earth orbit satellite parameter adaptive control method with solar panel drive device according to claim 5, characterized in that, Step S5 is as follows: right The update was successful using the parameter adaptive law: To carry out, on The update was successful using the parameter adaptive law: conduct, Represents the update gain of the adaptive law; Represents the estimated parameters The derivative; This represents the correction gain of the adaptive law. This represents the error angular velocity loop control variable. .

7. An electronic device comprising a memory and a processor, wherein the memory stores a computer program, characterized in that, When the processor executes the computer program, it implements the steps of the method according to any one of claims 1-6.

8. A computer-readable storage medium for storing computer instructions, characterized in that, When the computer instructions are executed by the processor, they implement the steps of the method according to any one of claims 1-6.

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