Construction method and device of micro-vibration integrated model for screening disturbance vibration sources of spacecraft, computer program and storage medium

Abstract

The invention discloses a construction method and device of a micro-vibration integrated model for screening a spacecraft disturbance vibration source, a computer program and a storage medium, and relates to the technical field of spaceflight, in particular to the construction method of the micro-vibration integrated model for screening the spacecraft disturbance vibration source. The method comprises the following steps: S1, establishing an empirical model of a spacecraft disturbance vibration source; s2, establishing a dynamical model of the spacecraft structure; s3, establishing a model of the spacecraft optical system; and S4, establishing a micro-vibration model, and outputting a result. According to the spacecraft vibration disturbance source screening method based on the micro-vibration integrated model, dynamic coupling between an actual spacecraft structure and the vibration disturbance source is considered, six components of vibration disturbance generated by the vibration disturbance source can be comprehensively evaluated, and the situation that individual vibration disturbance component meets the requirement and other components do not meet the requirement is avoided. The method is suitable for screening vibration disturbance sources on spacecrafts, such as flywheels, control moment gyroscopes, SADA and the like.

Description

Technical Field

[0001] The invention relates to the field of aerospace technology, and in particular to a method for constructing a micro-vibration integrated model for screening spacecraft disturbance sources. Background Art

[0002] In order to meet the needs of high-precision observation and obtain more details of the observed targets, the resolution of optical remote sensing satellites is getting higher and higher. At the same time, the requirements for pointing stability of optical remote sensing satellites during imaging are also getting higher and higher. However, there are disturbance sources such as flywheels and CMG on the satellite. The micro-vibrations they generate can cause the visual axis of the high-resolution camera to shake, reduce its stability, and deteriorate the imaging quality. Therefore, the disturbance sources on the satellite must be suppressed. Common micro-vibration suppression methods can be divided into four aspects: reducing interference at the disturbance source, reducing the sensitivity of the payload, changing the propagation path of micro-vibrations, and installing vibration isolation devices. Among them, screening disturbance sources with smaller disturbance forces is the most direct and effective method.

[0003] At present, the screening of disturbance sources is mainly based on the screening of supplier control characteristics, while the screening based on micro-vibration, such as the method described in the prior art patent document CN109141471A - a method for implementing the screening of high-stability satellite flywheels based on micro-disturbance torque (published on January 4, 2019), uses an empirical formula, namely the screening factor, to evaluate the disturbance source, and the one with a smaller screening factor is selected as the better disturbance source. This method is based on the disturbance force measured by a fixed installation on the disturbance force test bench, ignoring the dynamic coupling between the actual satellite structure and the disturbance source, that is, the disturbance at the satellite structure mode will be amplified, and the disturbance source screened by this method may not be the best after being installed on the satellite structure. In addition, the disturbance generated by the disturbance source is a three-way disturbance force and a three-way disturbance torque. The influence of each component on the normal operation of the sensitive load is different. It is difficult to comprehensively evaluate each component through an empirical formula. The disturbance source screened by this method may not be the best under the joint action of the six disturbance components. Summary of the invention

[0004] In order to solve the problem in the prior art that the dynamic coupling between the actual spacecraft structure and the disturbance source is ignored and that some disturbance components meet the requirements while the other components do not, the present invention proposes a method for constructing a micro-vibration integrated model for screening spacecraft disturbance sources.

[0005] The construction method comprises:

[0006] S1. Steps to establish an empirical model of spacecraft disturbance sources;

[0007] S2. Steps for establishing a dynamic model of the spacecraft structure;

[0008] S3, steps to establish a spacecraft optical system model;

[0009] S4, a step of establishing a micro-vibration integrated model based on the models described in S1, S2 and S3.

[0010] Furthermore, the steps of establishing the empirical model of the spacecraft disturbance source are:

[0011]

[0012] Among them, F j (t) is the jth component of F, F is the disturbance force and torque generated by the disturbance source, expressed as a vector, F = [F x ,F y ,F z ,M x ,M y ] T , F x 、F y 、F z Respectively represent the disturbance forces generated by the disturbance source along the x, y, and z directions, M x 、M y Respectively represent the torques generated by the disturbance source around the x and y directions, n is the harmonic number of the disturbance source, M i is the amplitude of the ith harmonic, h i is the speed multiple of the i-th harmonic, α i is the initial phase of the i-th harmonic, and Ω is the angular velocity of the disturbance source.

[0013] Furthermore, the steps of establishing the dynamic model of the spacecraft structure are:

[0014]

[0015] Where η is the modal coordinate, Φ is the mode shape matrix,

[0016] ζ is the damping matrix, expressed as:

[0017]

[0018] Ω 2 is the modal frequency matrix, expressed as:

[0019]

[0020] Take the state variable as The state space expression corresponding to the dynamic model is:

[0021]

[0022] Among them, A sis the dynamic model system matrix, B s is the kinetic model input matrix, C s is the output matrix of the dynamics model, which can be expressed as:

[0023] C s =[Φ 0],

[0024] Furthermore, the steps of establishing the spacecraft optical system model are:

[0025] The spacecraft optical system model includes an optical camera, and the optical camera includes a plurality of optical elements;

[0026] The image shift p(t) generated by the optical camera is expressed as:

[0027]

[0028] Where m is the number of optical elements, L i is the sensitivity coefficient of the ith optical element, T i is the rigid body displacement of the i-th optical element, respectively expressed as:

[0029]

[0030] T i =C i q i ,

[0031] Among them, l 11 , l 12 , l 13 represents the image shift in the X direction caused by moving the i-th optical element unit displacement along the x, y, and z directions, respectively, and l 14 , l 15 , l 16 represents the image shift in the X direction caused by rotating the i-th optical element by a unit angle along the x, y, and z directions, respectively, and l 21 , l 22 , l 23 represents the image shift in the Y direction caused by moving the i-th optical element unit displacement along the x, y, and z directions, respectively, and l 24 , l 25 , l 26 They represent the image shift in the Y direction caused by rotating the i-th optical element by a unit angle along the x, y, and z directions, respectively.

[0032] qi is the displacement vector of each node on the i-th optical element, expressed as:

[0033] q i =[u i1 ,v i1 ,w i1,…,u ij ,v ij ,w ij ,…,u in ,v in ,w in ] T ,

[0034] Among them, u ij 、v ij 、w ij are the displacements of the jth node on the ith optical element in the x, y, and z directions, respectively;

[0035] Ci is a constant matrix expressed as:

[0036] C i =(A i T A i ) -1 A i T ,

[0037] Among them, Ai is a matrix related to the coordinates of the nodes on the optical element, expressed as:

[0038]

[0039] Among them, x ij ,y ij 、z ij are the coordinates of the jth node on the ith optical element in the x, y, and z directions, respectively;

[0040] The rigid body displacement T i Substituting the image shift p(t) generated by the optical camera into the model of the optical system, we can get:

[0041] p(t)=C o q o ,

[0042] Among them, q o is the degree of freedom of the optical element and is expressed as:

[0043] q o =[q1,…,q2,…,q m ] T ,

[0044] C o is the comprehensive coefficient matrix, expressed as:

[0045]

[0046] Among them, L 1x , L 1yare the sensitivity coefficients of the first optical element in the x and y directions,

[0047] L ix , L iy are the sensitivity coefficients of the i-th optical element in the x and y directions,

[0048] L mx , L my are the sensitivity coefficients of the mth optical element in the x and y directions respectively.

[0049] Further, the steps of establishing a micro-vibration integrated model based on the models described in S1, S2 and S3 are as follows:

[0050] The spacecraft's degrees of freedom q are divided into the degrees of freedom on the optical element q o and the remaining degrees of freedom q r , expressed as:

[0051]

[0052] Where k is the mode number;

[0053] The degree of freedom q on the optical element o , expressed as:

[0054]

[0055] The degree of freedom q on the optical element o Substituted into the optical system model p(t), we can get the final optical system model:

[0056]

[0057] According to the state space expression corresponding to the dynamic model and the final optical system model, the micro-vibration integrated model of the spacecraft can be obtained:

[0058]

[0059] Where P is the output of the optical system, C p is the comprehensive coefficient matrix.

[0060] Based on the same inventive concept, the present invention also provides a construction device for a micro-vibration integrated model for screening spacecraft disturbance sources, the construction device comprising:

[0061] Modules for building empirical models based on spacecraft disturbance sources;

[0062] Modules for building dynamic models based on spacecraft structures;

[0063] Build a module based on the spacecraft optical system model;

[0064] Establishing a module of a micro-vibration integrated model based on the module of the empirical model, the module of the dynamic model and the module of the optical system model;

[0065] A module that substitutes the disturbance source data into the micro-vibration model and outputs the results.

[0066] Based on the same inventive concept, the present invention also provides a computer storage medium for storing a computer program, characterized in that when the computer program is read by a computer, the computer executes a method for constructing a micro-vibration integrated model.

[0067] Based on the same inventive concept, the present invention also provides a computer, including a processor and a storage medium, characterized in that when the processor reads the computer program stored in the storage medium, the computer executes a method for constructing a micro-vibration integrated model.

[0068] Based on the same inventive concept, the present invention also provides a computer program product as a computer program, characterized in that when the computer program is read, a method for constructing a micro-vibration integrated model is implemented.

[0069] Beneficial Effects

[0070] The present invention proposes a method for constructing a micro-vibration integrated model for screening spacecraft disturbance sources:

[0071] 1. Consider the dynamic coupling between the actual spacecraft structure and the disturbance source;

[0072] 2. Comprehensive evaluation can be performed on the six components of disturbance generated by the disturbance source;

[0073] 3. Avoid the situation where some disturbance components meet the requirements while the other components do not.

[0074] It is suitable for screening disturbance sources on spacecraft, such as flywheels, control moment gyroscopes, SADA, etc. BRIEF DESCRIPTION OF THE DRAWINGS

[0075] Figure 1 It is the X and Y direction micro-vibration image shift index of the present invention;

[0076] Figure 2 The flywheel screening process of the present invention;

[0077] Figure 3 This is the Z04 flywheel disturbance waterfall diagram of the present invention;

[0078] Figure 4 This is a frequency domain image shift waterfall diagram caused by the Z04 flywheel disturbance described in the present invention;

[0079] Figure 5This is the Z05 flywheel disturbance waterfall diagram of the present invention;

[0080] Figure 6 This is a frequency domain image shift waterfall diagram caused by the Z05 flywheel disturbance described in the present invention. Implementation

[0081] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0082] Embodiment 1: This embodiment provides a method for constructing a micro-vibration integrated model for screening spacecraft disturbance sources. The method for constructing the micro-vibration integrated model includes:

[0083] S1. Establish an empirical model of spacecraft disturbance sources:

[0084]

[0085] Among them, F j (t) is the jth component of F, F is the disturbance force and torque generated by the disturbance source, expressed as a vector, F = [F x ,F y ,F z ,M x ,M y ] T , F x 、F y 、F z Respectively represent the disturbance forces generated by the disturbance source along the x, y, and z directions, M x 、M y Respectively represent the torques generated by the disturbance source around the x and y directions, n is the harmonic number of the disturbance source, M i is the amplitude of the ith harmonic, h i is the speed multiple of the i-th harmonic, α i is the initial phase of the i-th harmonic, and Ω is the angular velocity of the disturbance source.

[0086] S2. Establish the dynamic model of the spacecraft structure:

[0087]

[0088] Where η is the modal coordinate, Φ is the modal vibration matrix,

[0089] ζ is the damping matrix, expressed as:

[0090]

[0091] Ω 2 is the modal frequency matrix, expressed as:

[0092]

[0093] Take the state variable as Convert equation (2) into state space expression:

[0094]

[0095] Among them, A s is the dynamic model system matrix, B s is the kinetic model input matrix, C s is the output matrix of the dynamics model, which can be expressed as:

[0096]

[0097] S3. Establishing the spacecraft optical system model:

[0098] The spacecraft optical system model includes an optical camera, and the optical camera includes optical elements;

[0099] The image shift p(t) produced by the optical camera can be expressed as:

[0100]

[0101] Where m is the number of optical elements, L i is the sensitivity coefficient of the ith optical element, T i is the rigid body displacement of the i-th optical element, respectively expressed as:

[0102]

[0103] T i =C i q i (9)

[0104] Among them, l 11 , l 12 , l 13 represents the image shift in the X direction caused by moving the i-th optical element unit displacement along the x, y, and z directions, respectively, and l 14 , l 15 , l 16 represents the image shift in the X direction caused by rotating the i-th optical element by a unit angle along the x, y, and z directions, respectively, and l 21 , l 22 , l 23represents the image shift in the Y direction caused by moving the i-th optical element unit displacement along the x, y, and z directions, respectively, and l 24 , l 25 , l 26 They represent the image shift in the Y direction caused by rotating the i-th optical element by a unit angle along the x, y, and z directions, respectively.

[0105] qi is the displacement vector of each node on the i-th optical element, expressed as:

[0106] q i =[u i1 ,v i1 ,w i1 ,…,u ij ,v ij ,w ij ,…,u in ,v in ,w in ] T ,

[0107] Among them, u ij 、v ij 、w ij are the displacements of the jth node on the ith optical element in the x, y, and z directions, respectively;

[0108] Ci is a constant matrix expressed as:

[0109] C i =(A i T A i ) -1 A i T ,

[0110] Among them, Ai is a matrix related to the coordinates of the nodes on the optical element, expressed as:

[0111]

[0112] Among them, x ij ,y ij 、z ij are the coordinates of the jth node on the ith optical element in the x, y, and z directions, respectively;

[0113] Substituting equation (9) into equation (7), the model of the optical system can be obtained as follows:

[0114] p(t)=C o q o (11)

[0115] Among them, q o is the degree of freedom of the optical element and is expressed as:

[0116] q o =[q1,…,q2,…,q m ] T

[0117] C o is the comprehensive coefficient matrix, expressed as:

[0118]

[0119] Among them, L 1x , L 1y are the sensitivity coefficients of the first optical element in the x and y directions,

[0120] L ix , L iy are the sensitivity coefficients of the i-th optical element in the x and y directions,

[0121] L mx , L my are the sensitivity coefficients of the mth optical element in the x and y directions respectively.

[0122] The spacecraft's degrees of freedom q are divided into the degrees of freedom on the optical element q o and the remaining degrees of freedom q r , as shown in formula (13).

[0123]

[0124] Where k is the mode number.

[0125] From formula (13), we can get the degree of freedom q on the optical element: o , expressed as:

[0126]

[0127] Substituting equation (14) into equation (11), we can obtain

[0128]

[0129] From equations (5) and (15), the integrated model of micro-vibration of spacecraft can be obtained as follows:

[0130]

[0131] Where P is the output of the optical system, C p is the comprehensive coefficient matrix.

[0132] Substituting the measured disturbance into equation (16), the image shift p of the optical camera can be calculated. If p satisfies equation (17), the disturbance generated by the disturbance source meets the requirements of the spacecraft and can be installed on the spacecraft.

[0133]

[0134] In the formula, Micro-vibration index requirements for spacecraft.

[0135] Embodiment 2: This embodiment provides a device for constructing a micro-vibration integrated model for screening spacecraft disturbance sources, the device comprising:

[0136] A module for establishing empirical models based on spacecraft disturbance sources. This module is used to establish empirical models of disturbance sources, describe the disturbance forces and torques generated by disturbance sources under different working conditions, and determine the characteristic parameters of disturbance sources, such as harmonic components, angular velocity, amplitude, and initial phase, through experimental data and theoretical analysis.

[0137] A module for establishing a dynamic model based on the spacecraft structure. This module is used to establish a dynamic model of the spacecraft structure, taking into account the modal frequency and damping characteristics of the structure. Through modal analysis, the dynamic response of the spacecraft structure under different disturbance sources is determined. The model contains parameters such as modal coordinates, modal vibration matrix, damping matrix, and modal frequency matrix.

[0138] Establish a module based on the spacecraft optical system model. This module is used to establish a spacecraft optical system model, including an optical camera, which includes optical elements, describe the image shift of the optical system under the action of a disturbance source, determine the sensitivity coefficient and rigid body displacement of the optical element, and combine the displacement and sensitivity coefficient of the optical element through a comprehensive coefficient matrix to describe the dynamic behavior of the optical system.

[0139] A module for establishing a micro-vibration integrated model is based on the modules of the empirical model, the dynamic model and the optical system model. This module integrates the above three models to establish a complete micro-vibration integrated model, which accurately simulates the micro-vibration of the spacecraft under different disturbance sources by comprehensively considering the interaction between the disturbance source, the spacecraft structure and the optical system, and provides a unified framework for analyzing and optimizing the micro-vibration performance of the spacecraft.

[0140] Substitute the disturbance source data into the micro-vibration model and output the result module. This module is used to input the actual disturbance source data into the micro-vibration integrated model for simulation and analysis, output the impact of the disturbance source on the spacecraft structure and optical system, provide comprehensive evaluation indicators, help screen and optimize the disturbance source, and ensure that the micro-vibration performance of the spacecraft meets the design requirements.

[0141] These modules work together to build a complete integrated micro-vibration model for screening and optimizing the disturbance sources of the spacecraft to ensure the stability and performance of the spacecraft.

[0142] Embodiment 3: This embodiment provides a computer storage medium for storing a computer program. When the computer program is read by a computer, the computer executes the method provided in embodiment 1.

[0143] Embodiment 4: This embodiment provides a computer, including a processor and a storage medium. When the processor reads a computer program stored in the storage medium, the computer executes the method provided in embodiment 1.

[0144] Embodiment 5: This embodiment provides a computer program product, which is a computer program. When the computer program is read, the method provided in embodiment 1 is implemented.

[0145] Embodiment 6: This embodiment further describes the technical solution provided above in detail through specific embodiments:

[0146] A satellite is equipped with three flywheels in the X, Y and Z directions. The micro-vibration image shift indicators of the satellite in the X and Y directions are as follows: Figure 1 As shown in Figure 2, the flywheel screening process is as follows: Figure 2 shown.

[0147] For a flywheel numbered Z04, the disturbance force is tested as follows Figure 3 As shown, (a) represents the disturbance force generated by the Z04 flywheel along the x direction, (b) represents the disturbance force generated by the Z04 flywheel along the y direction, (c) represents the disturbance force generated by the Z04 flywheel along the z direction; (d) represents the torque generated by the Z04 flywheel around the x direction, (e) represents the torque generated by the Z04 flywheel around the y direction, and (f) represents the torque generated by the Z04 flywheel around the z direction. It can be seen that among the disturbance forces and torques generated by the Z04 flywheel, the maximum value of Fx is 0.011N, the frequency is 3.19Hz, and the speed is 2000rpm; F y The maximum value is 0.018N, the frequency is 471.2Hz, and the speed is 2000rpm; F z The maximum value is 0.03N, the frequency is 470.9Hz, and the speed is 1900rpm; M x The maximum value is 1.608N·mm, the frequency is 471.2Hz, and the speed is 2000rpm; M y The maximum value is 0.981N·mm, the frequency is 471.2Hz, and the speed is 2000rpm; M z The maximum value is 0.488N·mm, the frequency is 425Hz, and the rotation speed is 800rpm.

[0148] The frequency domain image shift waterfall diagram caused by the Z04 flywheel is calculated using formula (16) as follows: Figure 4As shown, (a) represents the frequency domain image shift caused by the Z04 flywheel in the x direction, and (b) represents the frequency domain image shift caused by the Z04 flywheel in the y direction. It can be seen that the maximum image shift in the x direction is 0.135 pixel, which occurs at 425 Hz at 800 rpm. The maximum image shift in the y direction is 0.647 pixel, which occurs at 19.94 Hz at 1200 rpm. The image shifts in the x and y directions are both beyond the tolerance value, so the flywheel is not suitable for installation on the satellite.

[0149] Consider another flywheel numbered Z05. The disturbance force is tested as follows: Figure 5 As shown, (a) represents the disturbance force generated by the Z05 flywheel along the x direction, (b) represents the disturbance force generated by the Z05 flywheel along the y direction, (c) represents the disturbance force generated by the Z05 flywheel along the z direction; (d) represents the torque generated by the Z05 flywheel around the x direction, (e) represents the torque generated by the Z05 flywheel around the y direction, and (f) represents the torque generated by the Z05 flywheel around the z direction. It can be seen that F x The maximum value is 0.0012N, the frequency is 33.13Hz, and the speed is 2000rpm; F y The maximum value is 0.012N, the frequency is 33.13Hz, and the speed is 2000rpm; F z The maximum value is 0.003N, the frequency is 492.8Hz, and the speed is 2000rpm; M x The maximum value is 0.08N·mm, the frequency is 33.13Hz, and the speed is 2000rpm; M y The maximum value is 0.08N·mm, the frequency is 33.13Hz, and the speed is 2000rpm; M z The maximum value is 0.05N·mm, the frequency is 106Hz, and the rotation speed is 400rpm.

[0150] The frequency domain image shift waterfall diagram caused by the Z05 flywheel is calculated using formula (16) as follows: Figure 6 As shown, (a) represents the frequency domain image shift caused by the Z05 flywheel in the x direction, and (b) represents the frequency domain image shift caused by the Z05 flywheel in the y direction. It can be seen that the maximum image shift in the x direction is 0.013pixel, which occurs at 423.9Hz at 1600rpm. The maximum image shift in the y direction is 0.044pixel, which occurs at 19.88Hz at 1200rpm. The image shifts in the x and y directions are both within the tolerance range, so the flywheel is suitable for installation on this satellite.

[0151] When there are many flywheels or other disturbance sources, the above method can also be used to screen them, and finally the disturbance source with smaller disturbance is selected to be installed on the satellite, so as to reduce the risk of the disturbance generated by the disturbance source affecting the imaging quality of the optical camera.

[0152] The technical solution provided by the present invention is further described in detail above through several specific implementation modes in order to highlight the advantages and benefits of the technical solution provided by the present invention. However, the several specific implementation modes described above are not intended to be used as limitations on the present invention. Any reasonable modification and improvement of the present invention, combination of implementation modes and equivalent substitution within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for constructing a micro-vibration integrated model for screening spacecraft disturbance sources, characterized in that: The construction method comprises: S1. Steps to establish an empirical model of spacecraft disturbance sources; S2. Steps for establishing a dynamic model of the spacecraft structure; S3, steps to establish a spacecraft optical system model; S4, a step of establishing a micro-vibration integrated model based on the models described in S1, S2 and S3.

2. The method for constructing a micro-vibration integrated model according to claim 1, characterized in that: The steps of establishing the empirical model of the spacecraft disturbance source are: Among them, F j (t) is the jth component of F, F is the disturbance force and torque generated by the disturbance source, expressed as a vector, F = [f x , f y , f z , m x , m y ] T , f x 、f y 、f z Respectively represent the disturbance forces generated by the disturbance source along the x, y, and z directions, m x 、m y Respectively represent the torques generated by the disturbance source around the x and y directions, n is the harmonic number of the disturbance source, M i is the amplitude of the ith harmonic, h i is the speed multiple of the i-th harmonic, α i is the initial phase of the i-th harmonic, and Ω is the angular velocity of the disturbance source.

3. The method for constructing a micro-vibration integrated model according to claim 1, characterized in that: The steps of establishing the dynamic model of the spacecraft structure are: Where η is the modal coordinate, Φ is the mode shape matrix, is the damping matrix, expressed as: Ω 2 is the modal frequency matrix, expressed as: Take the state variable as The state space expression corresponding to the dynamic model is: Among them, A s is the dynamic model system matrix, B s is the kinetic model input matrix, C s is the output matrix of the dynamics model, which can be expressed as:

4. The method for constructing a micro-vibration integrated model according to claim 1, characterized in that: The steps of establishing the spacecraft optical system model are: The spacecraft optical system model includes an optical camera, and the optical camera includes a plurality of optical elements; The image shift p(t) generated by the optical camera is expressed as: Where m is the number of optical elements, L i is the sensitivity coefficient of the ith optical element, T i is the rigid body displacement of the i-th optical element, respectively expressed as: T i =C i q i , Among them, l 11 , l 12 , l 13 represents the image shift in the X direction caused by moving the i-th optical element unit displacement along the x, y, and z directions, respectively, and l 14 , l 15 , l 16 represents the image shift in the X direction caused by rotating the i-th optical element by a unit angle along the x, y, and z directions, respectively, and l 21 , l 22 , l 23 represents the image shift in the Y direction caused by moving the i-th optical element unit displacement along the x, y, and z directions, respectively, and l 24 , l 25 , l 26 They represent the image shift in the Y direction caused by rotating the i-th optical element by a unit angle along the x, y, and z directions, respectively. qi is the displacement vector of each node on the i-th optical element, expressed as: q i =[u i1 ,v i1 ,w i1 ,…,u ij ,v ij ,w ij ,…,u in ,v in ,w in ] T , Among them, u ij 、v ij 、w ij are the displacements of the jth node on the ith optical element in the x, y, and z directions, respectively; Ci is a constant matrix, expressed as: C i =(A i T A i ) -1 A i T , Among them, Ai is a matrix related to the coordinates of the nodes on the optical element, expressed as: Among them, x ij ,y ij 、z ij are the coordinates of the jth node on the ith optical element in the x, y, and z directions, respectively; The rigid body displacement T i Substituting the image shift p(t) generated by the optical camera into the model of the optical system, we can get: p(t)=C o q o , Among them, q o is the degree of freedom of the optical element and is expressed as: q o =[q1,…,q2,…,q m ] T , C o is the comprehensive coefficient matrix, expressed as: Among them, L 1x , L 1y are the sensitivity coefficients of the first optical element in the x and y directions, L ix , L iy are the sensitivity coefficients of the i-th optical element in the x and y directions, L mx , L my are the sensitivity coefficients of the mth optical element in the x and y directions respectively.

5. The method for constructing a micro-vibration integrated model according to claim 1, characterized in that: Steps to establish a micro-vibration integrated model based on the models described in S1, S2 and S3: The spacecraft's degrees of freedom q are divided into the degrees of freedom on the optical element q o and the remaining degrees of freedom q r , expressed as: Where k is the mode number; The degree of freedom q on the optical element o , expressed as: The degree of freedom q on the optical element o Substituted into the optical system model p(t), we can get the final optical system model: According to the state space expression corresponding to the dynamic model and the final optical system model, the micro-vibration integrated model of the spacecraft can be obtained: Where P is the output of the optical system, C p is the comprehensive coefficient matrix.

6. A spacecraft disturbance source screening device based on a micro-vibration integrated model, characterized in that: The device comprises: Modules for building empirical models based on spacecraft disturbance sources; Modules for building dynamic models based on spacecraft structures; Build a module based on the spacecraft optical system model; Establishing a module of a micro-vibration integrated model based on the module of the empirical model, the module of the dynamic model and the module of the optical system model; A module that substitutes the disturbance source data into the micro-vibration model and outputs the results.

7. A computer storage medium for storing a computer program, characterized in that: When the computer program is read by a computer, the computer executes the method for constructing a micro-vibration integrated model according to any one of claims 1 to 5.

8. A computer, comprising a processor and a storage medium, characterized in that: When the processor reads the computer program stored in the storage medium, the computer executes the method for constructing a micro-vibration integrated model according to any one of claims 1 to 5.

9. A computer program product, being a computer program, characterized in that When the computer program is read, the method for constructing the micro-vibration integrated model according to any one of claims 1 to 5 is implemented.

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