Spacecraft in-orbit attitude control and structural dynamics coupling simulation method and system

Abstract

The invention provides a spacecraft in-orbit attitude control and structural dynamics coupling simulation method and system. The method comprises the following steps: establishing a spacecraft simplified three-dimensional model; defining an analysis step, a moment and a concentrated force, researching an attitude control and flexible structure vibration suppression equation and compiling the equation into a UAMP subprogram, and simulating an attitude control moment and a flexible structure vibration suppression acting force; and finally, evaluating whether the attitude angle precision of the spacecraft and the vibration suppression effect of the flexible structure reach expectations or not. According to the method, the influence of the motion and deformation of each flexible part on the attitude and mass center displacement of the central rigid body during the on-orbit attitude control of the spacecraft is considered, the structural dynamics characteristics of the large-flexibility spacecraft under the on-orbit attitude control condition are disclosed, and a new technical approach is provided for evaluating the structural mechanical property, reliability and pointing precision of the spacecraft.

Description

Technical Field

[0001] The present invention relates to the technical field of dynamics simulation, and in particular to a method and system for simulating the coupling of on-orbit attitude control and structural dynamics of a large flexible spacecraft using Abaqus software. Background Art

[0002] When a spacecraft is in orbit, it needs to perform attitude stabilization or attitude maneuvering control to complete specific tasks, change orbits, or avoid danger. Early spacecraft were small in size and had high structural rigidity. For the simulation study of their in-orbit attitude control, the spacecraft only needed to be regarded as a simple rigid body or rigid system, ignoring the structural dynamic response of the spacecraft itself. For the most common spacecraft configurations such as "large mass center rigid body + small mass flexible attachment", the coupling coefficient method is generally used to extract the constraint modes of the flexible attachment, calculate the inertia force of each modal vibration mode, and finally convert it into the reaction force or torque acting on the large mass center rigid body, so as to carry out the simulation study of the spacecraft's in-orbit attitude control.

[0003] However, with the development of modern spacecraft design concepts and application requirements, "large flexible spacecraft with highly integrated payload and platform" (hereinafter referred to as large flexible spacecraft) has become one of the current important development directions. For example, high-orbit imaging satellites equipped with ultra-large aperture optical payloads, large width, ultra-high resolution and long-term continuous observation capabilities are important tools for meteorological research, scientific search and rescue and other activities. This type of satellite configuration has a large number of lightweight and low-damping truss structures to reduce weight. Its overall dimensions are large but the central rigid body characteristics are not significant, the structural flexibility is large but the fundamental frequency is very low, and the mass characteristics are dispersed but the modes are dense. Since the coupling coefficient method does not pay attention to the influence of the movement and deformation of the flexible attachment itself on the attitude and displacement of the central rigid body, and it is difficult to clearly distinguish the central rigid body and the flexible attachment in the structural design of the large flexible spacecraft, the coupling coefficient method cannot realize the coupled simulation analysis of the on-orbit attitude control and structural dynamics of this type of large flexible spacecraft, and thus meet the evaluation requirements of the mechanical performance, reliability, pointing accuracy and other aspects of the flexible structure of the spacecraft under large-scale attitude maneuvers.

[0004] Abaqus is a powerful engineering simulation finite element software. It not only has dynamic analysis modules to solve simple linear or complex nonlinear problems of various structures, but also provides a wide range of flexible subroutine interfaces, allowing users to define complex material constitutive, unit or acoustic, optical, thermal, magnetic and other physical fields to meet specific coupling analysis needs. At present, there is still a lack of attention to the coupled simulation analysis of on-orbit attitude control and structural dynamics of large flexible spacecraft. Since in actual engineering problems, the attitude of the spacecraft is indirectly changed by applying forces or torques instead of directly giving the attitude motion parameters of the spacecraft, it is closer to the actual engineering needs to couple the on-orbit attitude motion and structural dynamic characteristics of the spacecraft. Summary of the invention

[0005] In view of the defects in the prior art, the object of the present invention is to provide a method and system for coupling simulation of spacecraft on-orbit attitude control and structural dynamics.

[0006] A spacecraft on-orbit attitude control and structural dynamics coupling simulation method provided by the present invention comprises:

[0007] Step S1: Establish a simplified three-dimensional geometric model of the spacecraft in the finite element analysis software, mesh the model and assign material parameters to each component;

[0008] Step S2: Use implicit dynamics as the analysis step, turn on geometric nonlinearity, and define the total simulation time, initial incremental step, minimum incremental step, and maximum incremental step;

[0009] Step S3: setting three-axis moments and concentrated forces at the central star and truss of the three-dimensional simplified geometric model, simulating attitude control moments and flexible structure vibration suppression forces, and setting coefficients and amplitudes of the three-axis moments and concentrated forces;

[0010] Step S4: establishing the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure, and compiling them into a subroutine script of the finite element analysis software;

[0011] Step S5: Establish an implicit dynamics analysis task, select the subroutine script and submit it for calculation, draw the attitude angle curve and three index curves for measuring the vibration suppression effect of the flexible structure, and analyze whether the stable attitude angle accuracy of the spacecraft and the vibration suppression effect of the flexible structure meet expectations.

[0012] Further, the three-dimensional simplified geometric model includes: a primary mirror, a secondary mirror, a primary mirror segment truss, a secondary mirror segment truss and a central star;

[0013] The specific structure is ignored during modeling, and the equivalent is performed in the form of uniformly distributed mass or mass points; the primary mirror, secondary mirror, central star body and the triangle plate in the truss cell are modeled with shell units, and the truss structure is modeled with beam units; when the model is meshed, the number of meshes is controlled, taking into account the calculation accuracy and amount of the model.

[0014] Furthermore, the finite element analysis software adopts Abaqus software;

[0015] The step S3 comprises: in the Load module of the Abaqus software, setting the coefficient and amplitude of the flexible structure vibration suppression force and the attitude control torque, and selecting user as the amplitude type, indicating that the UAMP subroutine interface is allowed to be called;

[0016] The magnitude of the attitude control torque and the flexible structure vibration suppression force is the torque or the coefficient of the torque multiplied by the amplitude.

[0017] Furthermore, the step S4 includes: compiling the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure into a UAMP subroutine interface script based on the Fortran language, and selecting and calling it when submitting the calculation in the CAE interface of the Abaqus software; the output of the UAMP subroutine script is the amplitude of the attitude control torque and the flexible structure vibration suppression force.

[0018] Furthermore, in step S5:

[0019] The pointing angle of the central star axis is used as the attitude angle of the spacecraft. The pointing angle is represented by a follow-up coordinate system established with the help of some unit nodes on the central star. The attitude angles of the target pointing and the current real-time pointing are both represented by a set of quaternions.

[0020] The deviation D between the geometric center distance of the primary and secondary mirrors and the nominal axial distance of the spacecraft, the distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α between the normal vectors of the primary and secondary mirror planes are used as three indicators to measure the effect of suppressing the spacecraft structure vibration. When the three indicators are all less than the expected values, it is considered that the spacecraft flexible structure vibration is effectively suppressed.

[0021] A spacecraft on-orbit attitude control and structural dynamics coupling simulation system provided by the present invention comprises:

[0022] Module M1: Establish a simplified three-dimensional geometric model of the spacecraft in the finite element analysis software, mesh the model and assign material parameters to each component;

[0023] Module M2: Use implicit dynamics as the analysis step, turn on geometric nonlinearity, and define the total simulation time, initial incremental step, minimum incremental step, and maximum incremental step;

[0024] Module M3: Set three-axis moments and concentrated forces at the central star and truss of the three-dimensional simplified geometric model to simulate the attitude control moment and the flexible structure vibration suppression force, and set the coefficients and amplitudes of the three-axis moments and concentrated forces;

[0025] Module M4: Establish the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure, and compile them into subroutine scripts of the finite element analysis software;

[0026] Module M5: Establish an implicit dynamics analysis task, select the subroutine script and submit the calculation, draw the attitude angle curve and three index curves for measuring the vibration suppression effect of the flexible structure, and analyze whether the stable attitude angle accuracy of the spacecraft and the vibration suppression effect of the flexible structure meet expectations.

[0027] Further, the three-dimensional simplified geometric model includes: a primary mirror, a secondary mirror, a primary mirror segment truss, a secondary mirror segment truss and a central star;

[0028] The specific structure is ignored during modeling, and the equivalent is performed in the form of uniformly distributed mass or mass points; the primary mirror, secondary mirror, central star body and the triangle plate in the truss cell are modeled with shell units, and the truss structure is modeled with beam units; when the model is meshed, the number of meshes is controlled, taking into account the calculation accuracy and amount of the model.

[0029] Furthermore, the finite element analysis software adopts Abaqus software;

[0030] The module M3 includes: in the Load module of the Abaqus software, setting the coefficient and amplitude of the flexible structure vibration suppression force and the attitude control torque, and selecting user as the amplitude type, indicating that the UAMP subroutine interface is allowed to be called;

[0031] The magnitude of the attitude control torque and the flexible structure vibration suppression force is the torque or the coefficient of the torque multiplied by the amplitude.

[0032] Furthermore, the module M4 includes: compiling the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure into the UAMP subroutine interface script based on the Fortran language, and selecting and calling it when submitting the calculation in the CAE interface of the Abaqus software; the output of the UAMP subroutine script is the amplitude of the attitude control torque and the flexible structure vibration suppression force.

[0033] Furthermore, in the module M5:

[0034] The pointing angle of the central star axis is used as the attitude angle of the spacecraft. The pointing angle is represented by a follow-up coordinate system established with the help of some unit nodes on the central star. The attitude angles of the target pointing and the current real-time pointing are both represented by a set of quaternions.

[0035] The deviation D between the geometric center distance of the primary and secondary mirrors and the nominal axial distance of the spacecraft, the distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α between the normal vectors of the primary and secondary mirror planes are used as three indicators to measure the effect of suppressing the spacecraft structure vibration. When the three indicators are all less than the expected values, it is considered that the spacecraft flexible structure vibration is effectively suppressed.

[0036] Compared with the prior art, the present invention has the following beneficial effects:

[0037] The present invention takes into account the influence of the movement and deformation of each flexible component on the central rigid body attitude and center of mass displacement during the on-orbit attitude control of a large flexible spacecraft. It can reveal the structural dynamic characteristics of a large flexible spacecraft under the on-orbit attitude control conditions, and realize the coupled simulation of the on-orbit attitude control and structural dynamics of a large flexible spacecraft. It provides a new technical approach for the evaluation of the structural mechanical properties, reliability and pointing accuracy of a large flexible spacecraft, further reduces the technical risks in the development of spacecraft, and shortens the design cycle of the spacecraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0039] Figure 1 The figure is a flow chart of the simulation method proposed in the present invention.

[0040] Figure 2 It is a three-dimensional geometric model of a large flexible spacecraft with a high degree of integration of a load and a platform established by using Abaqus software in an embodiment of the present invention.

[0041] Figure 3 Graph showing the distribution of six observation points for vibration suppression of a highly flexible spacecraft structure in an embodiment of the present invention.

[0042] Figure 4 It is a graph of attitude angles of a highly flexible spacecraft during large-amplitude maneuvers in an embodiment of the present invention.

[0043] Figure 5 Graph D is a curve of the deviation between the geometric center distance between the primary and secondary mirrors of a large flexible spacecraft and the axial nominal distance L in an embodiment of the present invention.

[0044] Figure 6 It is a distance d curve from the geometric center of the secondary mirror of the large flexible spacecraft to the normal vector of the primary mirror plane in the embodiment of the present invention.

[0045] Figure 7 It is the angle α curve of the normal vectors of the primary and secondary mirror planes of the large flexible spacecraft in the embodiment of the present invention. DETAILED DESCRIPTION

[0046] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0047] like Figure 1As shown, the present invention provides a method for coupling simulation of on-orbit attitude control and structural dynamics of a large flexible spacecraft using Abaqus software, comprising:

[0048] Step 1: Take a large flexible spacecraft with a high degree of integration of a certain payload and platform as the research object, analyze the configuration composition of the large flexible spacecraft, obtain the structural material parameters of the large flexible spacecraft, as well as the layout and performance parameters of the flywheel or torque gyro and truss actuator.

[0049] Step 2: Use Abaqus software to establish a three-dimensional geometric model of a large flexible spacecraft, ignore the specific structure of each payload type of the spacecraft, and use the equivalent uniformly distributed mass and mass points. Focusing on the platform structure of the large flexible spacecraft, use shell elements and beam elements to establish a simplified model, mesh the model, and assign material parameters to each component.

[0050] Step 3: Select "Implicit Dynamics" as the analysis step, turn on geometric nonlinearity, and define the total simulation time, initial incremental step, minimum incremental step, and maximum incremental step.

[0051] Step 4: According to the layout and performance parameters of the flywheel or torque gyro and truss actuator in step 1, set three-axis torque and concentrated force at the central star and truss of the model to simulate the attitude control torque and flexible structure vibration suppression force, and set the coefficients and amplitudes of the above forces and torques. Select user as the amplitude type to allow calling the UAMP subroutine interface.

[0052] Step 5: Establish the attitude control equations and flexible structure vibration suppression equations of the large flexible spacecraft, and compile them into UAMP subroutine scripts based on Fortran language.

[0053] Step 6: Establish an implicit dynamics analysis task, select the UAMP subroutine script and submit the calculation, draw the attitude angle curve and three index curves to measure the vibration suppression effect of the flexible structure, and analyze whether the stable attitude angle accuracy and flexible structure vibration suppression effect of the large flexible spacecraft meet expectations based on the simulation results.

[0054] For step one, the role of the flywheel or torque gyro is to output the attitude control torque of the spacecraft, and the role of the truss actuator is to output the actuating force for suppressing the vibration of the flexible structure, thereby realizing the structural vibration suppression of the spacecraft; the attitude control torque and structural vibration suppression actuating force required for the large-amplitude maneuver simulation of the large-flexible spacecraft are calculated by the attitude control equation and the structural vibration suppression equation respectively. The input quantities of the two algorithms are the structural deformation and rigid body displacement during the maneuver of the large-flexible spacecraft, and the output quantities are the attitude control torque and actuating force.

[0055] The simplified large flexible spacecraft platform consists of a primary mirror, a secondary mirror, a primary mirror segment truss, a secondary mirror segment truss and a central body. A flywheel or a torque gyro is arranged at the central body, and a truss actuator is arranged on the primary mirror segment truss in an optimal layout. The primary and secondary mirrors are made of wood, and the central body and truss are made of aluminum alloy.

[0056] The material properties of wood are shown in the following table:

[0057]

[0058] The material properties of aluminum alloy are shown in the following table:

[0059]

[0060] For step two, the three-dimensional geometric model of the large flexible spacecraft in the embodiment of the present invention, whose simplified platform configuration consists of a primary mirror, a secondary mirror, a primary mirror segment truss, a secondary mirror segment truss and a central star body, will be ignored during modeling, and will be equivalent in the form of evenly distributed mass or mass points; the primary mirror, secondary mirror, central star body and the triangle plates in the truss cells will be modeled with shell units, and the truss structure will be modeled with beam units; when meshing the model, it is necessary to properly control the number of meshes, taking into account both the model calculation accuracy and the amount of calculation.

[0061] A large number of measurement and control load units, power supplies, etc. will be arranged at the central body. When modeling these loads in Abaqus software, the specific structure will be ignored and the loads will be equivalent in the form of uniformly distributed mass. An extremely complex optical imaging system will be arranged at the secondary mirror, which is also simplified to a mass point of equal mass. The primary mirror, secondary mirror, central body and triangle plates in the truss cell are modeled with shell units, and the truss structure is modeled with beam units. When using Mesh in Abaqus software to mesh the model, it is necessary to properly control the number of grids to take into account the calculation accuracy and amount of the model. Figure 2 After meshing, the model in the embodiment includes 42175 S4R shell elements and 3367 B31 beam elements, a total of 46837 nodes and 281022 degrees of freedom.

[0062] Regarding step three, the details are as follows: for the large-scale maneuvering and structural vibration suppression simulation of flexible spacecraft, the total simulation time ranges from tens to hundreds of seconds. For this type of long-duration dynamic simulation, an implicit algorithm should be used, and considering the large rotation angles and large displacements in the simulation, it is necessary to turn on the geometric nonlinearity option in the Abaqus software analysis step to ensure the authenticity of the results. Implicit analysis has the problem of convergence difficulties. It is necessary to reasonably set parameters such as the initial incremental step size, the minimum incremental step size, and the maximum incremental step size in the analysis step to improve the computational convergence of the simulation. Among them, the maximum incremental step size is generally set to one-fifth to one-tenth of the sampling period or control period of the control system. The control period of the control system in the embodiment of the present invention is 0.1s, so the maximum incremental step size is set to 0.2s. The initial incremental step size has a great influence on the convergence of the simulation. After debugging, the initial incremental step size is set to 1×10 -13 s, it has good convergence. The minimum incremental step size is set to 1×10 -15 s, and the total simulation time is set to 650s. The analysis step settings of the embodiment of the present invention are specifically as follows:

[0063]

[0064] For step 4, in the Load module of Abaqus software, set the coefficients and amplitudes of the structural vibration suppression force and attitude control torque. Select user as the amplitude type, which means that the UAMP subroutine interface can be called. The size of the attitude control torque and the flexible structure vibration suppression force is the torque or the coefficient of the torque multiplied by the amplitude.

[0065] In the embodiment of the present invention, the moment vector of the attitude control torque at the central star body points to the positive direction of the Y axis, and the upper limit of the moment is set to 1000 N·m; according to the layout plan of the truss actuator, 15 flexible structure vibration suppression actuators are set, and the upper limit of a single actuator is 500 N. In the load module of the Abaqus software, the coefficients and amplitudes of the above forces and moments are set, and the amplitude type is selected as user, indicating that the UAMP subroutine interface is allowed to be called.

[0066] For step five, the attitude control equations and structural vibration suppression equations of the large flexible spacecraft are compiled into the UAMP subroutine interface script based on the Fortran language, and selected for calling when submitting the calculation in the CAE interface of the Abaqus software; the output of the UAMP subroutine script is the amplitude of the attitude control torque and the flexible structure vibration suppression force.

[0067] In an embodiment of the present invention, the attitude control torque and structural vibration suppression actuating force required for the large-amplitude maneuver simulation of a large-flexible spacecraft are calculated by the attitude control equation and the structural vibration suppression equation, respectively. The input quantities of the two algorithms are the structural deformation and rigid body displacement during the maneuver of the large-flexible spacecraft, and the output quantities are the attitude control torque and the actuating force.

[0068] (1) The attitude control equation of a large flexible spacecraft is a PID algorithm with an additional saturation constraint, which is expressed as follows:

[0069]

[0070] Among them, τ=(τ1,τ2,τ3),||τ|| ∞ =max{|τ1|,|τ2|,|τ3|}; J is the moment of inertia matrix of the controlled object; k, T and c are the control parameters corresponding to the proportional link, integral link and differential link of the PID algorithm, respectively. In the embodiment of the present invention, the parameters k=9.54, T=10 and c=0.55; e=(e1, e2, e3) is the error vector between the current posture of the controlled object and the target posture; ω=(ω1, ω2, ω3) is the angular velocity vector; u=(u1, u2, u3) is the posture control torque vector; U is the maximum value of the posture control torque allowed to be output; sat(*) is the saturation state function, which is defined as follows:

[0071]

[0072] (2) The vibration suppression of the highly flexible spacecraft structure in the embodiment of the present invention requires real-time observation of the distance changes of six observation points on the primary mirror and the secondary mirror. The distribution of the observation points is as follows: Figure 3 There are 15 observations in the structural vibration suppression equation. The observation is the relative distance d between any two observation points, which is defined as follows:

[0073] d=(d 12 , d 13 , d 14 , d 15 , d 16 , d 23 , d 24 , d 25 , d 26 , d 34 , d 35 , d 36 , d 45 , d 46 , d 56 )

[0074]

[0075] in, Represents the initial coordinates of the i-th observation point.

[0076] The observed quantity needs to be processed by a filter. In the embodiment of the present invention, a fourth-order Bessel low-pass filter is used, and the cut-off frequency is set to 0.5 Hz. According to the modal characteristics of the large flexible spacecraft in the embodiment of the present invention, the filter is written as the following linear system form:

[0077]

[0078] Where y is the input signal of the low-pass filter; is the output signal of the low-pass filter.

[0079] The output structural vibration suppression force is the thrust and tension along the truss axis (thrust is defined as a positive value and tension is defined as a negative value). The force vector F c The calculation formula is:

[0080]

[0081] in, is the measurement of relative deformation after passing through the low-pass filter; The time differential of the relative deformation measurement after passing through the low-pass filter; G y and G v is the flexible structure vibration control rate matrix in the embodiment of the present invention, and the matrix is ​​related to the modal characteristics of the large flexible spacecraft in the embodiment.

[0082] (3) Based on the Fortran language, the attitude control equations of the large flexible spacecraft and the structural vibration suppression equations need to be compiled into the UAMP subroutine interface script and selected and called when submitting the calculation in the CAE interface of the Abaqus software. The UAMP subroutine allows the user to extract the deformation, displacement, coordinates and other information of the specified nodes or units of the model in real time during the calculation, and finally obtain the amplitude (AMPVALUENEW) of the force or moment through the script program calculation. The UAMP subroutine interface script template has been provided in the help document of the Abaqus software. The user only needs to edit the program in the reserved compilation section to output the amplitude parameter AMPVALUENEW.

[0083] For step six, the pointing angle of the central star body's axis is used as the attitude angle of the large flexible spacecraft. The pointing angle is characterized by a follow-up coordinate system established with the help of some unit nodes on the central star body. The attitude angles of the target pointing and the current real-time pointing are represented by a set of quaternions. The deviation D between the geometric center distance of the primary and secondary mirrors and the nominal axial distance of the large flexible spacecraft, the distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α between the normal vectors of the primary and secondary mirror planes are used as three indicators to measure the vibration suppression effect of the large flexible spacecraft structure. When the above three indicators are all less than the expected values, it can be considered that the vibration of the spacecraft flexible structure is effectively suppressed.

[0084] (1) The characterization method of the attitude angle of a large flexible spacecraft is as follows:

[0085] In the embodiment of the present invention, the pointing angle of the central star's axis is used as the attitude angle of the large flexible spacecraft. The pointing angle needs to be characterized by defining a set of follow-up coordinate systems bound to the central star. However, in the Abaqus software, it is impossible to establish a local coordinate system that moves with the central star, so the follow-up coordinate system can only be established with the help of some unit nodes on the central star. In the embodiment of the present invention, an end face of one of the star cabins of the central star is selected to establish a follow-up coordinate system, and the axial direction of the large flexible spacecraft is used as the Z axis of the follow-up coordinate system, and the X axis and Y axis point to the same global coordinate system of the model.

[0086] The target pointing and current real-time pointing attitude angle of the large flexible spacecraft model are both represented by a set of quaternions. The specific method is as follows:

[0087] Assume a rotation matrix R, we have:

[0088]

[0089] Among them, n x =(R 11 ,R 21 ,R 31 ) T ,n Y =(R 12 ,R 22 ,R 32 ) T and n Z =(R 13 ,R 23 ,R 33 ) T are the unit column vectors of the X-axis, Y-axis and Z-axis of the moving coordinate system respectively.

[0090] The quaternion Q of the current real-time pointing attitude angle c It is expressed as:

[0091] Q c =(a,b,c,w)

[0092]

[0093] In the embodiment of the present invention, the quaternion of the target pointing attitude angle is set to:

[0094]

[0095] This means that the large flexible spacecraft will rotate 5° in the XZ plane.

[0096] (2) In the embodiment of the present invention, there are three types of indicators for measuring the vibration suppression effect of the large flexible spacecraft structure, namely, the deviation D between the geometric center distance of the primary and secondary mirrors and the axial nominal distance L of the large flexible spacecraft, the distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α between the normal vectors of the primary and secondary mirror planes. When the above three indicators are all less than the expected values, it can be considered that the vibration of the spacecraft flexible structure is effectively suppressed. The calculation methods of the three indicators are as follows:

[0097] like Figure 3 As shown, the coordinates of the three observation points on the primary and secondary mirrors are and Where i = 1, 2, 3. Three observation points on the primary or secondary mirror can determine a plane equation:

[0098] Ax+By+Cz+D=0

[0099] Assume D = 1, then the coordinates of the three observation points on the primary mirror satisfy the equation:

[0100]

[0101] definition and are the normal vectors of the primary mirror plane and the secondary mirror plane respectively, then:

[0102]

[0103] Among them, O and O′ are the centers of the circumscribed circles of the triangles formed by the three observation points on the primary and secondary mirrors, respectively. Therefore, the deviation D between the geometric center distance of the primary and secondary mirrors and the nominal axial distance L of the large flexible spacecraft can be expressed as:

[0104]

[0105] The distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane can be expressed as:

[0106]

[0107] The angle α between the normal vectors of the primary and secondary mirror planes can be expressed as:

[0108]

[0109] (3) Analysis of simulation results

[0110] The total analysis time of the simulation results in the embodiment of the present invention is 650s, which is divided into two stages. In the first stage (0s-250s), attitude control is turned on to make the large flexible spacecraft turn to the target direction (rotate 5° in the XZ plane). After the spacecraft reaches the target attitude angle, the actuator control is turned on in the second stage (250s-650s). During this process, attitude control is kept on. According to the simulation results, the attitude angle curve, the deviation D curve of the geometric center distance of the primary and secondary mirrors and the axial nominal distance L of the large flexible spacecraft, the distance d curve from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α curve between the normal vectors of the primary and secondary mirror planes are respectively drawn, as shown in FIG. Figures 4 to 7 shown.

[0111] The simulation results show that the attitude control torque does not exceed the preset upper limit of 1000 N·m during the whole simulation process, indicating that the saturation state constraint in the PID algorithm is effective. Figure 4 The attitude control accuracy in the simulation is high. After about 400s, the attitude control torque approaches 0N·m, indicating that the attitude control torque converges after the attitude of the large flexible spacecraft is effectively controlled, but there is still a certain overshoot in the attitude angle curve, and the k, T and c parameters of PID need to be further adjusted. The flexible structure vibration suppression force begins to be applied after about 250s. The force does not exceed the upper limit of 500N in the whole process simulation, indicating that the saturation state constraint condition in the PID algorithm is effective. After about 350s, the force is basically stable within the range of ±100N. Figure 5 to Figure 7 They are the deviation D curve of the geometric center distance between the primary and secondary mirrors and the axial nominal distance L of the large flexible spacecraft, the distance d curve from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α curve between the normal vectors of the primary and secondary mirror planes. The curve results show that all three indicators can basically approach 0 after 350s (i.e., 100s after the flexible structure vibration suppression is turned on), close to the initial state of the model, which shows that the structural vibration of the large flexible spacecraft in the embodiment of the present invention is effectively suppressed.

[0112] In summary, the present invention proposes a coupling simulation method for on-orbit attitude control and structural dynamics of a large flexible spacecraft using Abaqus software, which simulates the structural dynamics state of a large flexible spacecraft under on-orbit attitude control, realizes attitude pointing control and structural vibration suppression of a large flexible spacecraft under large-amplitude attitude maneuvers, and provides a new technical approach for evaluating the structural mechanical properties, reliability and pointing accuracy of a large flexible spacecraft.

[0113] The present invention also provides a spacecraft on-orbit attitude control and structural dynamics coupling simulation system. The spacecraft on-orbit attitude control and structural dynamics coupling simulation system can be realized by executing the process steps of the spacecraft on-orbit attitude control and structural dynamics coupling simulation method, that is, those skilled in the art can understand the spacecraft on-orbit attitude control and structural dynamics coupling simulation method as a preferred implementation of the spacecraft on-orbit attitude control and structural dynamics coupling simulation system.

[0114] A spacecraft on-orbit attitude control and structural dynamics coupling simulation system, comprising:

[0115] Module M1: Establish a simplified three-dimensional geometric model of the spacecraft in the finite element analysis software, mesh the model and assign material parameters to each component.

[0116] Module M2: Use implicit dynamics as the analysis step, turn on geometric nonlinearity, and define the total simulation time, initial incremental step, minimum incremental step, and maximum incremental step.

[0117] Module M3: Set three-axis moments and concentrated forces at the central star and truss of the three-dimensional simplified geometric model to simulate the attitude control moment and the flexible structure vibration suppression force, and set the coefficients and amplitudes of the three-axis moments and concentrated forces.

[0118] Module M4: Establish the attitude control equations of the spacecraft and the vibration suppression equations of the flexible structure, and compile them into subroutine scripts of the finite element analysis software.

[0119] Module M5: Establish an implicit dynamics analysis task, select the subroutine script and submit the calculation, draw the attitude angle curve and three index curves for measuring the vibration suppression effect of the flexible structure, and analyze whether the stable attitude angle accuracy of the spacecraft and the vibration suppression effect of the flexible structure meet expectations.

[0120] Those skilled in the art know that, in addition to realizing the system and its various devices, modules, and units provided by the present invention in a purely computer-readable program code, it is entirely possible to realize the same functions in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered as a hardware component, and the devices, modules, and units included therein for realizing various functions can also be regarded as structures within the hardware component; the devices, modules, and units for realizing various functions can also be regarded as both software modules for realizing the method and structures within the hardware component.

[0121] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A method for coupling simulation of on-orbit attitude control and structural dynamics of a spacecraft, characterized in that: include: Step S1: Establish a simplified three-dimensional geometric model of the spacecraft in the finite element analysis software, mesh the model and assign material parameters to each component; Step S2: Use implicit dynamics as the analysis step, turn on geometric nonlinearity, and define the total simulation time, initial incremental step, minimum incremental step, and maximum incremental step; Step S3: setting three-axis moments and concentrated forces at the central star and truss of the three-dimensional simplified geometric model, simulating attitude control moments and flexible structure vibration suppression forces, and setting coefficients and amplitudes of the three-axis moments and concentrated forces; Step S4: establishing the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure, and compiling them into a subroutine script of the finite element analysis software; Step S5: Establish an implicit dynamics analysis task, select the subroutine script and submit it for calculation, draw the attitude angle curve and three index curves for measuring the vibration suppression effect of the flexible structure, and analyze whether the stable attitude angle accuracy of the spacecraft and the vibration suppression effect of the flexible structure meet expectations.

2. The spacecraft on-orbit attitude control and structural dynamics coupling simulation method according to claim 1 is characterized in that: The three-dimensional simplified geometric model includes: a primary mirror, a secondary mirror, a primary mirror segment truss, a secondary mirror segment truss and a central star; The specific structure is ignored during modeling, and the equivalent is performed in the form of uniformly distributed mass or mass points; the primary mirror, secondary mirror, central star body and the triangle plate in the truss cell are modeled with shell units, and the truss structure is modeled with beam units; when the model is meshed, the number of meshes is controlled, taking into account the calculation accuracy and amount of the model.

3. The spacecraft on-orbit attitude control and structural dynamics coupling simulation method according to claim 1 is characterized in that: The finite element analysis software adopts Abaqus software; The step S3 comprises: in the Load module of the Abaqus software, setting the coefficient and amplitude of the flexible structure vibration suppression force and the attitude control torque, and selecting user as the amplitude type, indicating that the UAMP subroutine interface is allowed to be called; The magnitude of the attitude control torque and the flexible structure vibration suppression force is the torque or the coefficient of the torque multiplied by the amplitude.

4. The spacecraft on-orbit attitude control and structural dynamics coupling simulation method according to claim 3 is characterized in that: The step S4 comprises: compiling the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure into the UAMP subroutine interface script based on the Fortran language, and selecting and calling it when submitting the calculation in the CAE interface of the Abaqus software; the output of the UAMP subroutine script is the amplitude of the attitude control torque and the flexible structure vibration suppression force.

5. The spacecraft on-orbit attitude control and structural dynamics coupling simulation method according to claim 1, characterized in that: In step S5: The pointing angle of the central star axis is used as the attitude angle of the spacecraft. The pointing angle is represented by a follow-up coordinate system established with the help of some unit nodes on the central star. The attitude angles of the target pointing and the current real-time pointing are both represented by a set of quaternions. The deviation D between the geometric center distance of the primary and secondary mirrors and the nominal axial distance of the spacecraft, the distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α between the normal vectors of the primary and secondary mirror planes are used as three indicators to measure the effect of suppressing the spacecraft structure vibration. When the three indicators are all less than the expected values, it is considered that the spacecraft flexible structure vibration is effectively suppressed.

6. A spacecraft on-orbit attitude control and structural dynamics coupling simulation system, characterized in that: include: Module M1: Establish a simplified three-dimensional geometric model of the spacecraft in the finite element analysis software, mesh the model and assign material parameters to each component; Module M2: Use implicit dynamics as the analysis step, turn on geometric nonlinearity, and define the total simulation time, initial incremental step, minimum incremental step, and maximum incremental step; Module M3: Set three-axis moments and concentrated forces at the central star and truss of the three-dimensional simplified geometric model to simulate the attitude control moment and the flexible structure vibration suppression force, and set the coefficients and amplitudes of the three-axis moments and concentrated forces; Module M4: Establish the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure, and compile them into subroutine scripts of the finite element analysis software; Module M5: Establish an implicit dynamics analysis task, select the subroutine script and submit the calculation, draw the attitude angle curve and three index curves for measuring the vibration suppression effect of the flexible structure, and analyze whether the stable attitude angle accuracy of the spacecraft and the vibration suppression effect of the flexible structure meet expectations.

7. The spacecraft on-orbit attitude control and structural dynamics coupling simulation system according to claim 6, characterized in that: The three-dimensional simplified geometric model includes: a primary mirror, a secondary mirror, a primary mirror segment truss, a secondary mirror segment truss and a central star; The specific structure is ignored during modeling, and the equivalent is performed in the form of uniformly distributed mass or mass points; the primary mirror, secondary mirror, central star body and the triangle plate in the truss cell are modeled with shell units, and the truss structure is modeled with beam units; when the model is meshed, the number of meshes is controlled, taking into account the calculation accuracy and amount of the model.

8. The spacecraft on-orbit attitude control and structural dynamics coupling simulation system according to claim 6, characterized in that: The finite element analysis software adopts Abaqus software; The module M3 includes: in the Load module of the Abaqus software, setting the coefficient and amplitude of the flexible structure vibration suppression force and the attitude control torque, and selecting user as the amplitude type, indicating that the UAMP subroutine interface is allowed to be called; The magnitude of the attitude control torque and the flexible structure vibration suppression force is the torque or the coefficient of the torque multiplied by the amplitude.

9. The spacecraft on-orbit attitude control and structural dynamics coupling simulation system according to claim 8, characterized in that: The module M4 includes: compiling the attitude control equation of the spacecraft and the vibration suppression equation of the flexible structure into the UAMP subroutine interface script based on the Fortran language, and selecting and calling it when submitting the calculation in the CAE interface of the Abaqus software; the output of the UAMP subroutine script is the amplitude of the attitude control torque and the flexible structure vibration suppression force.

10. The spacecraft on-orbit attitude control and structural dynamics coupling simulation system according to claim 6, characterized in that: In the module M5: The pointing angle of the central star axis is used as the attitude angle of the spacecraft. The pointing angle is represented by a follow-up coordinate system established with the help of some unit nodes on the central star. The attitude angles of the target pointing and the current real-time pointing are both represented by a set of quaternions. The deviation D between the geometric center distance of the primary and secondary mirrors and the nominal axial distance of the spacecraft, the distance d from the geometric center of the secondary mirror to the normal vector of the primary mirror plane, and the angle α between the normal vectors of the primary and secondary mirror planes are used as three indicators to measure the effect of suppressing the spacecraft structure vibration. When the three indicators are all less than the expected values, it is considered that the spacecraft flexible structure vibration is effectively suppressed.

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