A method for joint simulation of a micro-satellite attitude control system based on a control moment gyro

Abstract

The application discloses a kind of microsatellite attitude control system joint simulation method based on control moment gyro, it is related to satellite attitude control system simulation field.This method specifically includes: in SolidWorks, the three-dimensional model of control moment gyro and satellite shell is built;Solar wing flexibility model is built in HyperMesh;The above-mentioned all models are imported into Adams to build Adams satellite virtual prototype model;Satellite attitude control system model is built in Simulink;Finally, through Adams and Simulink data interaction, joint simulation is realized.The application has the advantages that the integration of mechanical-control system simulation is realized, the dynamic performance of control moment gyro is accurately verified, and model parameters can be adjusted at will, shorten the product development cycle;Meanwhile, the control strategy extension, iteration and optimization of satellite attitude control system can also be realized, to provide simulation support for the development of system.

Description

Technical Field

[0001] This invention relates to the field of satellite attitude control system simulation, and specifically to a co-simulation method for a microsatellite attitude control system based on a control moment gyroscope. Background Technology

[0002] As a crucial actuator in a satellite attitude control system, the control moment gyroscope directly determines the quality of the system's operation, making its design paramount. The numerous configurations of control moment gyroscopes contribute to their structural complexity, increasing the difficulty of developing physical prototypes.

[0003] Currently, the development of physical prototype control moment gyroscopes still adopts a development cycle of development, prototyping, processing, and testing. If structural parameters or configurations need to be modified during this process, redevelopment and testing are required. While this cycle-based development model ensures product quality, its drawbacks—long design cycles, high development costs, and high testing costs—cannot be ignored. This model is gradually becoming unsuitable for the needs of today's high-quality, information-driven modern life.

[0004] To address the aforementioned issues, researchers have proposed using software simulation methods for developing attitude control systems for microsatellites. However, the development of attitude control systems for microsatellites based on control moment gyroscopes still primarily relies on single-software simulation. Furthermore, most current spacecraft are equipped with solar panels, whose structure, elastically connected to the spacecraft via a rotating shaft, constitutes a rigid-flexible coupled system. The core of its dynamics research lies in the dynamic modeling of flexible multibody systems. In single-software simulations, to reduce the modeling difficulty of flexible multibody systems, the solar panels are often simplified into rigid models, which undoubtedly reduces the realism of the simulation.

[0005] Therefore, there is an urgent need for a co-simulation method for satellite attitude control systems to achieve the integration of mechanical and control system simulation models. Summary of the Invention

[0006] The purpose of this invention is to provide a co-simulation method for a microsatellite attitude control system based on a control moment gyroscope, integrating the satellite virtual prototype model and the satellite attitude control model, thus overcoming the shortcomings of single-software simulation. This co-simulation method can accurately verify the dynamic performance of the control moment gyroscope, while reducing the difficulty of dynamic modeling of flexible multibody systems and improving simulation efficiency and realism. The specific technical solution is as follows:

[0007] A co-simulation method for a microsatellite attitude control system based on a control moment gyroscope includes the following steps:

[0008] Step S1: Create a 3D model. Specifically:

[0009] A 3D model of the control moment gyroscope and a 3D model of the satellite's outer shell were created using SolidWorks. A flexible model of the satellite's solar panels was created using HyperMesh. During the creation of the flexible solar panel model, the original plate-like structure was equated to a shell structure to reduce the computational load of the simulation. The connecting rods between the solar panels and the satellite body were ignored, and the connection point between the solar panels and the satellite body was defined as the master node. The area between the master node and the solar panels was defined as the rigid connection area. The 3D model of the control moment gyroscope is applicable to all configurations, specifically including double-parallel, triple-parallel, pyramidal, square-pyramid, and pentagonal pyramidal configurations.

[0010] Step S2: Build the Adams virtual prototype model. Specifically:

[0011] Import the 3D model established in step S1 into Adams, and add appropriate constraints and drives based on the motion form of the control torque gyroscope to establish a complete Adams satellite virtual prototype model.

[0012] Step S3: Build a Simulink satellite attitude control system model. Specifically:

[0013] A mathematical simulation model for satellite attitude control is built in Simulink. This model includes a controller module, a control torque gyro manipulation law module, and a satellite attitude kinematics module. The controller module can employ any type of controller, including PD control, sliding mode control, and MPC control, to generate control torque commands. The control torque gyro manipulation law module can use any type of manipulation law, including Moore-Ponros pseudo-inverse manipulation law, singular robust pseudo-inverse manipulation law, generalized singular robust pseudo-inverse manipulation law, and pseudo-inverse manipulation law with idle motion, to convert control torque commands into frame motor speed commands. The satellite attitude kinematics module is used to recursively calculate attitude quaternion changes and update the satellite attitude.

[0014] Step S4: Implement Adams-Simulink co-simulation and output simulation data. Specifically:

[0015] Import the Adams virtual prototype model into the Simulink satellite attitude control system model; determine the transmission signals between the Adams virtual prototype model and the Simulink satellite attitude control system model, and establish a two-way co-simulation interface; the Simulink satellite attitude control system model transmits signals to the Adams virtual prototype model, and the Adams virtual prototype model performs dynamic simulation; the Adams virtual prototype model transmits the signals generated by the dynamic simulation to the Simulink satellite attitude control system model, and the Simulink satellite attitude control system model performs kinematic simulation.

[0016] Preferably, the three-dimensional model is established in step S1, specifically by establishing a three-dimensional model of the control moment gyroscope and a three-dimensional model of the satellite shell in SolidWorks. The three-dimensional model of the control moment gyroscope adopts a pyramid configuration, including four single-frame control moment gyroscopes and a support frame. The four single-frame control moment gyroscopes are fixed on the support frame in a pyramid configuration with a tilt angle of 54.7°. Each single-frame control moment gyroscope includes a frame motor, an outer frame, an inner frame, bearings, and a rotor system (including a rotor motor and a rotor).

[0017] Preferably, in step S1, a flexible satellite solar panel model is established in HyperMesh. Specifically, the original plate-like structure of the solar panel is converted into a shell structure using the Extract Mid-Layer command. The equivalent solar panel is then divided into multiple shell elements using finite element methods. The size of each shell element can be selected according to requirements. Material parameters are set for the shell elements, such as aluminum alloy. The material properties of the shell elements are defined. A master node is defined; its function is to ensure that the flexible solar panel model can establish connections with feature points of other structures in Adams. In this invention, only the impact of solar panel vibration on the overall structure is considered; therefore, the connecting rods between the solar panel and the satellite body are ignored. The master node is defined at the connection point between the solar panel and the satellite body, and a rigid connection region is defined between the master node and the solar panel. Constraints are created at the connection point; specifically, fixed constraints can be established at the connection point to obtain the flexible satellite solar panel model. The model obtained above can be exported as an MNF file according to the following steps: setting the load collector, creating an ASET, setting constraints and load types, setting the output, and generating the MNF file.

[0018] Preferably, in step S2, the Adams virtual prototype model is built by importing the built three-dimensional model of the control moment gyroscope, the three-dimensional model of the satellite shell, and the flexible model of the satellite solar array into Adams. Constraints and drives are added between the components of the three-dimensional model according to the motion form of the control moment gyroscope. The constraints and drives specifically include fixed joints, rotary joints, and rotary drives. Finally, the complete Adams virtual prototype model is assembled.

[0019] Specifically, the constraints are as follows: the bracket and outer frame form a fixed pair constraint G1; the outer frame and bearing form a fixed pair constraint G2; the inner frame and bearing form a rotating pair constraint X1; the frame stepper motor and inner frame form a rotating pair constraint X2; the rotor motor and inner frame form a fixed pair constraint G3; and the rotor and rotor motor form a rotating pair constraint X3. The drive is achieved by adding rotational drives to the rotating pair constraints X2 and X3 respectively.

[0020] Preferably, in step S3, a Simulink satellite attitude control mathematical simulation model is built. Specifically, the control torque gyro manipulation law module uses a singular robust pseudo-inverse manipulation law. The controller module uses a PD controller to calculate the control torque, and the calculation formula is as follows:

[0021] T=K p q e +K d w e

[0022] Where T is the control torque, q e and w e These are the error quaternion and the error angular velocity, respectively, K p and K d These are the PD parameters corresponding to the error quaternion and the error angular velocity, respectively.

[0023] Preferably, in step S4, the Adams virtual prototype model is imported into the Simulink satellite attitude control system model. Specifically, the FMI function in Adams is used to export the Adams virtual prototype model into a corresponding FMU model, and the FMU model is imported into Simulink.

[0024] Specifically, the step of using the FMI function in Adams to export the Adams virtual prototype model as the corresponding FMU model involves setting parameters, including the name of the control module, the prefix of the generated file, the input and output variables of the Adams virtual prototype model, the target software, the analysis type, and the solver, and then exporting it as an FMU model.

[0025] Specifically, importing the FMU model into Simulink involves setting parameters, including whether to display it dynamically, the simulation analysis mode, and the interaction frequency between Adams and Simulink, and then importing it into Simulink.

[0026] Preferably, in step S4, the transmission signal between the Adams virtual prototype model and the Simulink satellite attitude control system model is determined as follows: the Simulink satellite attitude control system model transmits the control torque gyroscope frame motor speed signal (calculated by the control torque gyroscope control law module) to the Adams virtual prototype model; the Adams virtual prototype model transmits the satellite attitude angular velocity signal to the Simulink satellite attitude control system model; the control torque gyroscope frame motor speed signal is defined as an input variable in the Adams virtual prototype model, and the satellite attitude angular velocity signal is defined as an output variable in Adams.

[0027] Compared with the prior art, the advantages of the present invention are:

[0028] The co-simulation method for microsatellite attitude control systems based on control moment gyroscopes of this invention uses HyperMesh to construct a flexible solar array model, which makes the simulation results closer to reality. In the construction of the flexible solar array model, the mid-level extraction command is used to convert the original plate structure of the solar panel into an equivalent shell structure, preserving key mechanical properties while reducing the computational load in the simulation process; connecting rods are ignored, eliminating the need for detailed modeling of connecting rods and avoiding the computational burden caused by local mesh refinement of the rods; the master node is defined at the connection point between the solar panel and the satellite body, and a rigid connection region is defined between the master node and the solar panel, creating constraints at the connection point to highlight the dominant vibration effect of the solar panel.

[0029] Adams virtual prototype models are used to replace the satellite attitude dynamics and actuator modules in conventional Simulink satellite attitude control system models to achieve Adams-Simulink co-simulation.

[0030] This invention presents a co-simulation method for a microsatellite attitude control system based on a control moment gyroscope. By combining the strengths of SolidWorks, HyperMesh, Adams, and Simulink software, a co-simulation system for the satellite attitude control system is established, achieving integrated simulation of the mechanical and control systems. This method accurately verifies the dynamic performance of the control moment gyroscope, and model parameters such as PD parameters, frame tilt angle, and rotor motor speed can be arbitrarily adjusted. It can provide a reference for control moment gyroscope prototypes with different structures and configurations, shortening the product development and testing cycles. Simultaneously, it can also enable the expansion, iteration, and optimization of the control strategy of the satellite attitude control system, providing simulation support for system development. Attached Figure Description

[0031] Figure 1 This is a schematic diagram illustrating the working principle of the co-simulation method for the attitude control system of a microsatellite provided in a specific embodiment of the present invention.

[0032] Figure 2 This is a three-dimensional model of a single-frame control torque gyroscope provided in a specific embodiment of the present invention.

[0033] Among them, 1 is the frame motor; 2 is the outer frame; 3 is the inner frame; 4 is the bearing; and 5 is the rotor system (including the rotor and the rotor motor).

[0034] Figure 3 This is a flexible solar panel model provided in a specific embodiment of the present invention.

[0035] Figure 4 This is the overall virtual prototype model of the satellite provided in a specific embodiment of the present invention. Detailed Implementation

[0036] The embodiments of the present invention will be described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways as defined and covered by the claims.

[0037] An embodiment of the present invention discloses a co-simulation method for a microsatellite attitude control system based on a control moment gyroscope, referring to... Figure 1 Specifically, the steps include the following:

[0038] Step S1: Create a 3D model. Specifically:

[0039] A 3D model of the control torque gyroscope and a 3D model of the satellite shell were created in SolidWorks. The 3D model of the control torque gyroscope includes four single-frame control torque gyroscopes and a support frame. The four single-frame control torque gyroscopes are fixed on the support frame in a pyramidal configuration. Each single-frame control torque gyroscope includes a frame motor 1, an outer frame 2, an inner frame 3, a bearing 4, and a rotor system 5 (including a rotor motor and a rotor), as shown in the reference. Figure 2 .

[0040] A flexible satellite solar panel model is established in HyperMesh. The specific method is as follows: using the Extract Mid-Layer command, the original plate-like structure of the solar panel is equivalent to a shell structure. Finite element analysis is performed on the equivalent solar panel to obtain multiple shell elements. The specific size of the shell elements can be selected according to requirements. Material parameters are set for the shell elements; in this embodiment, aluminum alloy is used. The material properties of the shell elements are defined. Master nodes are defined. The role of the master nodes is to ensure that the flexible solar panel model can establish connection relationships with feature points of other structures in Adams. In this embodiment, only the impact of solar panel vibration on the overall structure is considered; therefore, the connecting rods between the solar panel and the satellite body are ignored. The master nodes are defined at the connection points between the solar panel and the satellite body, and a rigid connection region is defined between the master nodes and the solar panel. Constraints are created at the connection points; in this embodiment, fixed constraints are established at the connection points to obtain the flexible satellite solar panel model. (Refer to...) Figure 3 The model obtained above can be exported as an MNF file by following these steps: setting up the load collector, creating an ASET, setting constraints and load types, setting the output, and generating the MNF file.

[0041] Step S2: Build the Adams virtual prototype model. Specifically:

[0042] Import the 3D model of the control moment gyroscope, the 3D model of the satellite shell, and the flexible model of the satellite solar array built in step S1 into Adams. Based on the motion of the control moment gyroscope, add constraints and drives (including fixed joints, revolute joints, and rotational drives) to the components in the 3D models. Finally, assemble them into a complete virtual prototype model, referring to... Figure 4 .

[0043] Specifically, the constraints are as follows: the bracket and the outer frame are a fixed pair constraint G1; the outer frame and the bearing are a fixed pair constraint G2; the inner frame and the bearing are a rotating pair constraint X1; the frame stepper motor and the inner frame are a rotating pair constraint X2; the rotor motor and the inner frame are a fixed pair constraint G3; the rotor and the rotor motor are a rotating pair constraint X3; and the drive is to add a rotation drive to the rotating pair constraints X2 and X3 respectively.

[0044] Step S3: Build a Simulink satellite attitude control system model. Specifically, build a satellite attitude control mathematical simulation model in Simulink. This model includes a controller module, a control torque gyro manipulation law module, and a satellite attitude kinematics module. Specifically, the control torque gyro manipulation law module employs a singular robust pseudo-inverse manipulation law. The controller module uses a PD controller to calculate the control torque, and the calculation formula is as follows:

[0045] T=K p q e +K d w e

[0046] Where T is the control torque, q e and w e These are the error quaternion and the error angular velocity, respectively, K p and K d These are the PD parameters corresponding to the error quaternion and the error angular velocity, respectively.

[0047] Step S4: Implement Adams-Simulink co-simulation and output simulation data. Specifically:

[0048] Import the Adams virtual prototype model into the Simulink satellite attitude control system model; determine the transmission signals between the Adams virtual prototype model and the Simulink satellite attitude control system model, and establish a two-way co-simulation interface; the Simulink satellite attitude control system model transmits signals to the Adams virtual prototype model, and the Adams virtual prototype model performs dynamic simulation; the Adams virtual prototype model transmits the signals generated by the dynamic simulation to the Simulink satellite attitude control system model, and the Simulink satellite attitude control system model performs kinematic simulation.

[0049] Specifically, the Adams virtual prototype model is imported into the Simulink satellite attitude control system model. In particular, the FMI function in Adams is used to export the Adams virtual prototype model into the corresponding FMU model, and then the FMU model is imported into Simulink.

[0050] Specifically, the step of using the FMI function in Adams to export the Adams virtual prototype model as the corresponding FMU model involves setting parameters, including the name of the control module, the prefix of the generated file, the input and output variables of the Adams virtual prototype model, the target software, the analysis type, and the solver, and then exporting it as an FMU model.

[0051] Specifically, importing the FMU model into Simulink involves setting parameters, including whether to display it dynamically, the simulation analysis mode, and the interaction frequency between Adams and Simulink, and then importing it into Simulink.

[0052] The transmission signals between the Adams virtual prototype model and the Simulink satellite attitude control system model are determined as follows: the Simulink satellite attitude control system model transmits the control torque gyroscope frame motor speed signal (calculated by the control torque gyroscope control law module) to the Adams virtual prototype model; the Adams virtual prototype model transmits the satellite attitude angular velocity signal to the Simulink satellite attitude control system model; the control torque gyroscope frame motor speed signal is defined as an input variable in the Adams virtual prototype model, and the satellite attitude angular velocity signal is defined as an output variable in Adams.

Claims

1. A method for joint simulation of a microsatellite attitude control system based on a control moment gyroscope, characterized in that, include: Step S1: Create a 3D model; Specifically: A three-dimensional model of the control moment gyroscope, a three-dimensional model of the satellite shell, and a flexible model of the satellite solar array are established. In the process of establishing the flexible model of the satellite solar array, the plate structure is equivalent to the shell structure, and the connecting rods between the solar panel and the satellite body are ignored. The connection point between the solar panel and the satellite body is defined as the main node, and the area between the main node and the solar panel is defined as the rigid connection area. Step S2: Build the Adams virtual prototype model; specifically: Import the 3D model established in step S1 into Adams, and add constraints and drives to the components in the 3D model based on the motion form of the control torque gyroscope to establish an Adams virtual prototype model. Step S3: Build a Simulink satellite attitude control system model; specifically: A mathematical simulation model for satellite attitude control was built in Simulink. The mathematical simulation model for satellite attitude control includes a controller module, a control torque gyroscope manipulation law module, and a satellite attitude kinematics module. Step S4: Implement Adams-Simulink co-simulation and output simulation data; specifically: Import the Adams virtual prototype model into the Simulink satellite attitude control system model; determine the transmission signals between the Adams virtual prototype model and the Simulink satellite attitude control system model, and establish a two-way co-simulation interface; the Simulink satellite attitude control system model transmits signals to the Adams virtual prototype model, and the Adams virtual prototype model performs dynamic simulation; the Adams virtual prototype model transmits the signals generated by the dynamic simulation to the Simulink satellite attitude control system model, and the Simulink satellite attitude control system model performs kinematic simulation; In step S1: A flexible model of the satellite solar array is established in HyperMesh. The specific method is as follows: using the extract mid-layer command, the original plate structure of the solar panel is equivalent to a shell structure; the equivalent solar panel is divided into multiple shell elements by finite element analysis, and the size of the shell elements is selected according to the requirements; material parameters are set and material properties are defined for the shell elements; master nodes are defined, ignoring the connecting rods between the solar panel and the satellite body, and only considering the impact of solar panel vibration on the whole, the master nodes are defined as the connection points between the solar panel and the satellite body, and the connection area between the master nodes and the solar panel is defined as a rigid connection area. Constraints are created at the connection points to obtain the flexible model of the satellite solar array. In step S2, the constraint and drive specifically include a fixed joint, a revolute joint, and a rotation drive; The constraints are specifically as follows: the bracket and the outer frame are fixed pair constraints G1; the outer frame and the bearing are fixed pair constraints G2; the inner frame and the bearing are rotating pair constraints X1; the frame stepper motor and the inner frame are rotating pair constraints X2; the rotor motor and the inner frame are fixed pair constraints G3; the rotor and the rotor motor are rotating pair constraints X3; the drive is to add rotation drive to the rotating pair constraints X2 and X3 respectively.

2. The method according to claim 1, wherein, In step S1, a three-dimensional model of the control torque gyroscope and a three-dimensional model of the satellite shell are created in SolidWorks. The three-dimensional model of the control torque gyroscope adopts a pyramid configuration, including four single-frame control torque gyroscopes and a support. The four single-frame control torque gyroscopes are fixed on the support, forming a pyramid configuration. The single-frame control torque gyroscope includes a frame motor, an outer frame, an inner frame, bearings, and a rotor system.

3. The co-simulation method for a microsatellite attitude control system based on a control moment gyroscope according to claim 1, characterized in that, In step S3, the controller module uses the PD controller to calculate the control torque, and the calculation formula is as follows: where T is the control torque, q e and w e are the error quaternion and error angular velocity, respectively, and K p and K d are the PD parameters corresponding to the error quaternion and error angular velocity, respectively. The control torque gyroscope manipulation law module specifically uses a singular robust pseudo-inverse manipulation law.

4. The co-simulation method for a microsatellite attitude control system based on a control moment gyroscope according to claim 1, characterized in that, In step S4, the Adams virtual prototype model is imported into the Simulink satellite attitude control system model. Specifically, the Adams virtual prototype model is exported as a corresponding FMU model using the FMI function in Adams, and then the FMU model is imported into Simulink. Exporting the Adams virtual prototype model as a corresponding FMU model using the FMI function in Adams involves setting parameters, including the name of the control module, the prefix of the generated file, the input and output variables of the Adams virtual prototype model, the target software, the analysis type, and the solver, and then exporting it as an FMU model. Importing the FMU model into Simulink involves setting parameters, including whether to display dynamically, the simulation analysis mode, and the interaction frequency between Adams and Simulink, and then importing it into Simulink.

5. The co-simulation method for a microsatellite attitude control system based on a control moment gyroscope according to claim 4, characterized in that, In step S4, the transmission signals between the Adams virtual prototype model and the Simulink satellite attitude control system model are determined. Specifically, the Simulink satellite attitude control system model transmits the control torque gyroscope frame motor speed signal to the Adams virtual prototype model; the Adams virtual prototype model transmits the satellite attitude angular velocity signal to the Simulink satellite attitude control system model; the control torque gyroscope frame motor speed signal is defined as an input variable in the Adams virtual prototype model, and the satellite attitude angular velocity signal is defined as an output variable in Adams.

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