CN109612664B - Method and system for identifying on-orbit vibration state of satellite flexible accessory by utilizing gyroscope data - Google Patents
Method and system for identifying on-orbit vibration state of satellite flexible accessory by utilizing gyroscope data Download PDFInfo
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- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
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Abstract
The invention provides a method for identifying the in-orbit vibration state of a satellite flexible accessory by utilizing gyro data, which comprises the following steps: an attitude angular velocity obtaining step: obtaining attitude angular velocity measurement data corresponding to i axis of under-damped free vibration area after satellite on-orbit jet closed-loop control(ii) a A filtering step: for omegai(t) filtering to obtain filtered i-axis attitude angular velocity data omega'i(t); time series calculation step: according to ω'i(t) calculating η time series of acquired modal variablesi(t); obtaining integral displacement; calculating end displacement; and calculating the relative displacement. Correspondingly, the invention also provides a system for identifying the in-orbit vibration state of the satellite flexible accessory by utilizing the gyro data. According to the method, other sensors do not need to be installed on the flexible accessory, and only the measurement data of the existing attitude sensor of the satellite platform is utilized for analysis and processing.
Description
Technical Field
The invention relates to the field of satellite monitoring, in particular to a method and a system for identifying an in-orbit vibration state of a satellite flexible accessory by utilizing gyroscope data. In particular to an in-orbit identification method and system for vibration state information of a flexible satellite, and particularly relates to an identification method and system for integral displacement, end displacement and relative displacement of the flexible satellite during in-orbit vibration.
Background
With the continuous development of aerospace technology, a large space structure is an important development direction in the aerospace field and also is a necessary infrastructure for space development (in-orbit spacecraft dynamics parameter identification technology research-cloud-climbing). In order to obtain the dynamic characteristics of the flexible satellite structure, the vibration mode parameters are usually calculated by methods such as finite element modeling, ground test, in-orbit identification and the like. Model simplification, condition assumption and the like in the finite element modeling process influence model precision, and particularly accurate modeling of contact mechanisms such as hinges and the like is difficult. Due to the influence of factors such as gravity and atmospheric resistance, large flexible structures are difficult to assemble on the ground and perform full-scale kinetic parameter identification tests. Therefore, for spacecraft with large flexible appendages, it is difficult to obtain precise structural dynamics through finite element modeling or ground testing. Based on the factors, the on-orbit modal parameter identification research on large flexible structures such as solar arrays, space spread antennas and the like is urgent and necessary, and meanwhile, the on-orbit modal parameter identification research method has higher theoretical significance and practical application value. The flexible satellite structure dynamic parameters comprise modal frequency, modal damping, array type, coupling coefficient and the like, have important physical significance, and can provide necessary support for structural design, structural health monitoring, structural fault diagnosis, structural vibration control and other aspects of space flexible parts [ CN102982196A ].
At present, modal parameter identification methods in the field of structural dynamics mainly comprise a frequency domain method, a time domain method and a recently-developed time-frequency domain method. Most methods require on-orbit excitation and arrangement of sensors on a flexible accessory to acquire information, such as [ large solar panel mode parameters on-orbit identification research _ li xiao-master paper ] research on minimum configuration number and optimal layout scheme of sensors on a flexible accessory; CN105486474A introduces a system and method for realizing in-orbit modal identification of satellite flexible components, which needs to perform pulse excitation on a flexible accessory, receive and monitor pulse response signals of each measuring point, and acquire acceleration response signals generated by the satellite flexible components in the steady-state operation process of the satellite in orbit; two methods of satellite solar array sensor layout are introduced [ CN106557633A ] and [ CN107609296A ]. The binocular vision measurement-based on-orbit identification method for the dynamic characteristics of the unfolding structure of the solar sail is characterized in that structural vibration displacement information is directly extracted from an image, and then the dynamic characteristics of the structure are obtained in real time through a working mode analysis technology, so that the on-orbit identification of the dynamic characteristics of the structure is realized. There are also two patents that propose on-orbit identification methods for flexible satellite modal parameters by using attitude angular velocity data measured by a gyroscope: CN103970964A introduces an in-orbit identification method for modal parameters of flexible satellites, which needs to acquire the moment applied to the flexible satellite body by the actuator and the angular velocity information of the flexible satellite body relative to the inertial coordinate system, and obtain the transfer function from the modal parameters and the moment to the angular velocity by using a subspace identification algorithm. CN105157728A proposes a flexible satellite modal parameter identification method capable of suppressing gyro noise influence, which is also to identify the modal frequency and modal damping ratio parameter of the whole satellite by using the measurement data of the satellite body angular velocity when the satellite is in orbit flight, and perform differential processing on the gyro data to suppress the identification error caused by the two noise parts.
The in-orbit kinetic parameter identification method for the flexible satellite has two constraints in the aspect of engineering application: firstly, the on-orbit spacecraft is difficult to apply known excitation required by kinetic parameter identification, and excitation sources are only generated for excitation by unfolding and folding of the on-orbit spacecraft structure, butt joint and separation of the structure, ignition of an engine and the like, and signals of the excitation sources are difficult to measure [ research on the kinetic parameter identification technology of the on-orbit spacecraft-cloud acquisition ]. Secondly, the number of the vibration sensors arranged on the flexible accessory is limited by the implementation of engineering. General flexible accessories need to be unfolded on the rail, and various speed, acceleration and displacement sensors are installed and fixed, cables are laid and the like to have adverse effects on an unfolding mechanism, so that the design difficulty is increased. In addition, the number of slip ring signal channels of the driving mechanism needs to be increased for rotating flexible parts such as a solar array.
Disclosure of Invention
In view of the shortcomings in the prior art, an object of the present invention is to provide a method and system for identifying an in-orbit vibration state of a satellite flexible accessory by using gyroscope data.
The method for identifying the in-orbit vibration state of the satellite flexible accessory by utilizing the gyro data comprises the following steps of:
an attitude angular velocity obtaining step: obtaining attitude angular velocity measurement data omega corresponding to i axis of under-damped free vibration area after satellite on-orbit air injection closed-loop controli(t), wherein the i-axis is any one of an X-axis, a Y-axis and a Z-axis in the space rectangular coordinate system;
a filtering step: for omegai(t) filtering to obtain filtered i-axis attitudeAngular velocity data ω'i(t)；
Time series calculation step: according to ω'i(t) calculating η time series of acquired modal variablesi(t)；
Overall displacement acquisition step according to ηi(t) and vibrational mode matrix data DiCalculating to obtain the integral displacement condition of each node on the flexible accessory;
end displacement calculating step according to ηi(t) and end node matrix data diAnd calculating the on-orbit vibration end displacement m of the satellite flexible accessoryi(t)；
Calculating the relative displacement according to ηi(t) calculating the array data of any two nodes on the flexible accessory to obtain the relative displacement m of any two nodes on the flexible accessory12(t)。
Preferably, in the attitude angular velocity acquisition step, ωiAnd (t) is obtained by measuring through a satellite gyroscope during the period of under-damped free vibration after the on-orbit air injection control of the satellite.
Preferably, the filtering step comprises the steps of:
a focus vibration band setting step: setting a frequency band of interest in the i-axis direction to [ f1,f2]Hz；
A stripping step: let the band-pass filter pair omegai(t) line filtering, stripping omegaiIn the (t), the periodic motion and the dither component exceeding the set value are obtained by retaining only the information on the vibration in the frequency band of interest'i(t)。
Preferably, in the overall displacement obtaining step, η is addedi(t) and vibrational mode matrix data DiAnd multiplying to obtain the integral displacement condition of each node on the flexible accessory in a set period.
Preferably, the method further comprises a main vibration mode judging step: judging whether the flexible accessory has only one main vibration mode in each vibration direction according to the overall displacement condition; if yes, continuing to execute the end displacement calculation step and the relative displacement calculation step; if not, returning to the step of executing the filtering and adjusting the concerned vibration frequency band.
Preferably, in the step of setting the vibration frequency band of interest, the vibration frequency band of interest in the i-axis direction is set according to a calculation result of a ground whole star structure dynamics analysis.
Preferably, the time series calculating step comprises the steps of:
a first derivative transformation step: according to the following formula to ω'i(t) carrying out treatment:
in the formula:
the first derivative of the modal variable corresponding to the i axis;
Jithe moment of inertia of the satellite in the direction of the i axis;
Bithe rotational coupling coefficient of the vibration mode of the flexible accessory corresponding to the i axis;
an integration step: to pair
Integral calculation is carried out to obtain ηi(t)。
Preferably, mi(t) is obtained by the following calculation formula:
mi(t)＝diηi(t)。
preferably, m12(t) is obtained by the following calculation formula:
m12(t)＝(d1-d2)ηi(t)
in the formula: d1、d2The two nodes on the flexible accessory corresponding to the relative displacement are array data respectively.
The invention also provides a system for identifying the in-orbit vibration state of the satellite flexible accessory by utilizing the gyro data, which comprises the following modules:
an attitude angular velocity acquisition module: obtaining the i-axis correspondence of the under-damped free vibration region after the on-orbit jet closed-loop control of the satelliteAttitude angular velocity measurement data ωi(t), wherein the i-axis is any one of an X-axis, a Y-axis and a Z-axis in the space rectangular coordinate system;
a filtering module: for omegai(t) filtering to obtain filtered i-axis attitude angular velocity data omega'i(t)；
A time series calculation module: according to ω'i(t) calculating η time series of acquired modal variablesi(t)；
Integral displacement acquisition module according to ηi(t) and vibrational mode matrix data DiCalculating to obtain the integral displacement condition of each node on the flexible accessory;
end displacement calculation Module according to ηi(t) and end node matrix data diAnd calculating the on-orbit vibration end displacement m of the satellite flexible accessoryi(t)；
Relative displacement calculation module according to ηi(t) calculating the array data of any two nodes on the flexible accessory to obtain the relative displacement m of any two nodes on the flexible accessory12(t)。
Compared with the prior art, the invention has the following beneficial effects:
1. the method meets the requirement of the on-orbit structure dynamic characteristic identification of the flexible satellite, and extracts the on-orbit vibration end displacement information of the flexible satellite.
2. Compared with other methods for calculating the dynamic parameters of the in-orbit structure of the satellite, the method does not need to install other sensors on the flexible accessory, and only utilizes the measurement data of the existing attitude sensor of the satellite platform to carry out analysis processing.
3. The increased satellite design, manufacturing difficulty and risk of on-orbit operation due to the installation of other vibration sensors are avoided.
4. The identified vibration state information of the flexible accessory can be used for dynamic analysis, fault diagnosis, performance evaluation and the like of the satellite on-orbit structure.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a schematic view of a configuration of a single-wing solar array satellite;
FIG. 2 is a change curve of attitude angular velocity of a satellite during an on-orbit jet closed-loop control period;
FIG. 3 is a time domain comparison graph before and after satellite attitude angular velocity filtering;
FIG. 4 is a frequency domain comparison before and after filtering of satellite attitude angular velocity;
FIG. 5 is a graph of the change in flexural modal variables;
FIG. 6 is a schematic illustration of an array of flexural attachment vibrational modes;
FIG. 7 is a graph of end displacement variation of vibration of a flexible attachment;
FIG. 8 is a graph of the end displacement variation of the vibration of two nodes on a flexible attachment;
FIG. 9 is a flow chart of a method for identifying an in-orbit vibration state of a satellite flexure attachment using gyroscope data.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that variations and modifications can be made by persons skilled in the art without departing from the spirit of the invention. All falling within the scope of the present invention.
In the description of the present invention, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, are not to be construed as limiting the present invention.
As shown in fig. 9, the method for identifying the in-orbit vibration state of the satellite flexible accessory by using the gyro data provided by the invention comprises the following steps: angular velocity of attitudeA degree obtaining step: obtaining attitude angular velocity measurement data omega corresponding to i axis of under-damped free vibration area after satellite on-orbit air injection closed-loop controli(t), wherein the i-axis is any one of an X-axis, a Y-axis and a Z-axis in the space rectangular coordinate system; a filtering step: for omegai(t) filtering to obtain filtered i-axis attitude angular velocity data omega'i(t); time series calculation step: according to ω'i(t) calculating η time series of acquired modal variablesi(t) overall displacement acquisition step according to ηi(t) and vibrational mode matrix data DiCalculating the integral displacement of each node on the flexible accessory, and calculating the displacement of the end part according to ηi(t) and end node matrix data diAnd calculating the on-orbit vibration end displacement m of the satellite flexible accessoryi(t) calculating the relative displacement according to ηi(t) calculating the array data of any two nodes on the flexible accessory to obtain the relative displacement m of any two nodes on the flexible accessory12(t) of (d). Preferably, in the attitude angular velocity acquisition step, ωiAnd (t) is obtained by measuring through a satellite gyroscope during the period of under-damped free vibration after the on-orbit air injection control of the satellite. That is, the relative displacement of the flexible attachment during the in-orbit vibration is identified only by using the gyro measurement information loaded on the satellite platform, and displacement, speed and acceleration vibration sensors mounted on the flexible attachment are not used.
The filtering step comprises the steps of: a focus vibration band setting step: setting a frequency band of interest in the i-axis direction to [ f1,f2]Hz; a stripping step: let the band-pass filter pair omegai(t) line filtering, stripping omegaiIn the (t), the periodic motion and the dither component exceeding the set value are obtained by retaining only the information on the vibration in the frequency band of interest'i(t) of (d). That is, in the stripping step, the attitude angular velocity measurement data is subjected to filtering processing using a band-pass filter, and long-period motion and other dither error components of the satellite body are stripped. Preferably, in the step of setting the vibration frequency band of interest, the vibration frequency band of interest in the i-axis direction is set according to a calculation result of a ground whole star structure dynamics analysis.
In an embodiment, η is used in the overall displacement obtaining stepi(t) and vibrational mode matrix data DiMultiplying to obtain the integral displacement condition of each node on the flexible accessory in a set period; the method for identifying the in-orbit vibration state of the satellite flexible accessory by utilizing the gyroscope data further comprises a main vibration mode judging step: judging whether the flexible accessory has only one main vibration mode in each vibration direction according to the overall displacement condition; if yes, continuing to execute the end displacement calculation step and the relative displacement calculation step; if not, returning to the step of executing the filtering and adjusting the concerned vibration frequency band.
The time series calculation step comprises a first derivative conversion step and an integration step. Wherein, the first derivative conversion step: according to the following formula to ω'i(t) carrying out treatment:
in the formula:
the first derivative of the modal variable corresponding to the i axis; j. the design is a squareiThe moment of inertia of the satellite in the direction of the i axis; b isiThe rotational coupling coefficient of the vibration mode of the flexible attachment corresponding to the i axis. An integration step: to pair
Integral calculation is carried out to obtain ηi(t)。
Preferably, mi(t) is obtained by the following calculation formula:
mi(t)＝diηi(t)。
preferably, m12(t) is obtained by the following calculation formula:
m12(t)＝(d1-d2)ηi(t)
in the formula: d1、d2The two nodes on the flexible accessory corresponding to the relative displacement are array data respectively.
Correspondingly, the invention also provides a system for identifying the in-orbit vibration state of the satellite flexible accessory by utilizing gyro data, which comprises the following modules: an attitude angular velocity acquisition module: obtaining attitude angular velocity measurement data omega corresponding to i axis of under-damped free vibration area after satellite on-orbit air injection closed-loop controli(t), wherein the i-axis is any one of an X-axis, a Y-axis and a Z-axis in the space rectangular coordinate system; a filtering module: for omegai(t) filtering to obtain filtered i-axis attitude angular velocity data omega'i(t); a time series calculation module: according to ω'i(t) calculating η time series of acquired modal variablesi(t) global displacement acquisition Module according to ηi(t) and vibrational mode matrix data DiCalculating the integral displacement of each node on the flexible accessory, and calculating the end displacement according to ηi(t) and end node matrix data diAnd calculating the on-orbit vibration end displacement m of the satellite flexible accessoryi(t) relative displacement calculation Module according to ηi(t) calculating the array data of any two nodes on the flexible accessory to obtain the relative displacement m of any two nodes on the flexible accessory12(t) of (d). Preferably, in the attitude angular velocity acquisition module, ωiAnd (t) is obtained by measuring through a satellite gyroscope during the period of under-damped free vibration after the on-orbit air injection control of the satellite. That is, the relative displacement of the flexible attachment during the in-orbit vibration is identified only by using the gyro measurement information loaded on the satellite platform, and displacement, speed and acceleration vibration sensors mounted on the flexible attachment are not used.
The filtering module comprises the following modules: focus vibration band setting module: setting a frequency band of interest in the i-axis direction to [ f1,f2]Hz; a stripping module: let the band-pass filter pair omegai(t) line filtering, stripping omegaiIn the (t), the periodic motion and the dither component exceeding the set value are obtained by retaining only the information on the vibration in the frequency band of interest'i(t) of (d). That is, in the stripping module, the attitude angular velocity measurement data is filtered by using a band-pass filter, and long-period motion and other high-frequency vibration error components of the satellite body are stripped. Preferably, the first and second electrodes are formed of a metal,and in the concerned vibration frequency band setting module, setting the concerned vibration frequency band in the i-axis direction according to the dynamic analysis and calculation result of the ground whole star structure.
In an embodiment, η is used in the overall displacement acquisition modulei(t) and vibrational mode matrix data DiMultiplying to obtain the integral displacement condition of each node on the flexible accessory in a set period; the system for identifying the on-orbit vibration state of the satellite flexible accessory by utilizing the gyroscope data further comprises a main vibration mode judgment module: judging whether the flexible accessory has only one main vibration mode in each vibration direction according to the overall displacement condition; if yes, the end displacement calculation module and the relative displacement calculation module are continuously executed; if not, returning to the execution of the filtering module and adjusting the concerned vibration frequency band.
The time sequence calculation module comprises a first derivative conversion module and an integration module. Wherein the first derivative conversion module: according to the following formula to ω'i(t) carrying out treatment:
in the formula:
the first derivative of the modal variable corresponding to the i axis; j. the design is a squareiThe moment of inertia of the satellite in the direction of the i axis; b isiThe rotational coupling coefficient of the vibration mode of the flexible attachment corresponding to the i axis. An integration module: to pair
Integral calculation is carried out to obtain ηi(t)。
Preferably, mi(t) is obtained by the following calculation formula:
mi(t)＝diηi(t)。
preferably, m12(t) is obtained by the following calculation formula:
m12(t)＝(d1-d2)ηi(t)
in the formula: d1、d2The two nodes on the flexible accessory corresponding to the relative displacement are array data respectively.
Preferred embodiments:
a remote sensing satellite is provided with a single-wing solar cell array, and the configuration of the single-wing solar cell array is shown in figure 1. The rolling X-axis attitude angular velocity measurement data during jet control after launch into orbit is shown in fig. 2.
Step 1: selecting rolling X-axis attitude angular velocity measurement data omega of satellite in orbit running periodX(t) analysis was performed (FIG. 2).
Step 2: according to the dynamic analysis and calculation result of the ground whole star structure, the frequency of the main vibration mode in the X direction is about 0.37Hz, and the concerned vibration frequency band is set to be [0.259,0.481]]Hz, using 5 th order Butterworth band-pass filter to measure data omega of satellite attitude angular velocity in orbit running periodX(t) carrying out filtering processing, stripping long-period motion and high-frequency vibration components in the rolling X-axis attitude angular velocity signal, only retaining vibration information in a frequency band of interest, and obtaining the filtered rolling X-axis attitude angular velocity omegaX′(t)。
The time domain contrast and the power spectral density contrast before and after the filtering of the attitude angular velocity measurement data are respectively shown in fig. 3 and fig. 4. As is clear from both time domain contrast and frequency domain contrast, signals within the [0.259,0.481] Hz band are preserved, and signals of the remaining frequency bands are significantly attenuated.
And step 3: according to the formula
The filtered single-axis attitude angular velocity omegaX' (t) conversion to the first derivative of the modal variable
Wherein, JX＝6872.13kg·m2The moment of inertia of the satellite in the direction of the rolling X axis, BX＝33.48762m·kg1/2The coefficient of rotational coupling for the vibrational modes of the flexure attachment. Second derivative of modal variables
Integration is performed to obtain a time series η of modal variablesX(t), see FIG. 5.
And 4, step 4: array data D of vibration modeX(FIG. 6) and modal variables ηXAnd (t) multiplying, and obtaining the vibration displacement condition of each node on the flexible accessory in the period. The relative displacement of any two nodes on the flexible accessory can be obtained by utilizing the difference of the vibration displacement of any two nodes on the flexible accessory.
And 5: matrix d with end nodesX(fig. 6) the end displacement of the flexible attachment during the in-orbit operation of the satellite can be obtained: m isX(t)＝dXηX(t) of (d). In this example, the pattern of end nodes is dX＝0.277kg-1/2The end displacement results obtained are shown in figure 7.
Step 6: two optional nodes are arranged on the flexible accessory, and the array type is d1And d2If the result of the relative displacement calculation is m12(t)＝(d1-d2)ηX(t), the results are shown in FIG. 8.
Those skilled in the art will appreciate that, in addition to implementing the systems, apparatus, and various modules thereof provided by the present invention in purely computer readable program code, the same procedures can be implemented entirely by logically programming method steps such that the systems, apparatus, and various modules thereof are provided in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system, the device and the modules thereof provided by the present invention can be considered as a hardware component, and the modules included in the system, the device and the modules thereof for implementing various programs can also be considered as structures in the hardware component; modules for performing various functions may also be considered to be both software programs for performing the methods and structures within hardware components.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes and modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (8)
1. A method for identifying the on-orbit vibration state of a satellite flexible accessory by utilizing gyro data is characterized by comprising the following steps:
an attitude angular velocity obtaining step: obtaining attitude angular velocity measurement data omega corresponding to i axis of under-damped free vibration area after satellite on-orbit air injection closed-loop controli(t), wherein the i-axis is any one of an X-axis, a Y-axis and a Z-axis in the space rectangular coordinate system;
a filtering step: for omegai(t) filtering to obtain filtered i-axis attitude angular velocity data omega'i(t)；
Time series calculation step: according to ω'i(t) calculating η time series of acquired modal variablesi(t)；
Overall displacement acquisition step according to ηi(t) and vibrational mode matrix data DiCalculating to obtain the integral displacement condition of each node on the flexible accessory;
end displacement calculating step according to ηi(t) and end node matrix data diAnd calculating the on-orbit vibration end displacement m of the satellite flexible accessoryi(t)；
Calculating the relative displacement according to ηi(t) calculating the array data of any two nodes on the flexible accessory to obtain the relative displacement m of any two nodes on the flexible accessory12(t)；
The time series calculating step includes the steps of:
a first derivative transformation step: according to the following formula to ω'i(t) carrying out treatment:
in the formula:
is an i-axis pairThe first derivative of the corresponding modal variable;
Jithe moment of inertia of the satellite in the direction of the i axis;
Bithe rotational coupling coefficient of the vibration mode of the flexible accessory corresponding to the i axis;
an integration step: to pair
Integral calculation is carried out to obtain ηi(t)；
Relative displacement m12(t) is obtained by the following calculation formula:
m12(t)＝(d1-d2)ηi(t)；
in the formula: d1、d2The two nodes on the flexible accessory corresponding to the relative displacement are array data respectively.
2. The method for identifying the in-orbit vibration state of the satellite flexible accessory according to claim 1, wherein in the attitude angular velocity obtaining step, ω isiAnd (t) is obtained by measuring through a satellite gyroscope during the period of under-damped free vibration after the on-orbit air injection control of the satellite.
3. A method for identifying an in-orbit vibration state of a satellite flexible accessory according to claim 1, wherein the filtering step comprises the steps of:
a focus vibration band setting step: setting a frequency band of interest in the i-axis direction to [ f1,f2]Hz；
A stripping step: let the band-pass filter pair omegai(t) line filtering, stripping omegaiIn the (t), the periodic motion and the dither component exceeding the set value are obtained by retaining only the information on the vibration in the frequency band of interest'i(t)。
4. The method for identifying the in-orbit vibration state of the satellite flexible accessory according to claim 3, wherein the overall displacement isIn the acquisition step, η is addedi(t) and vibrational mode matrix data DiAnd multiplying to obtain the integral displacement condition of each node on the flexible accessory in a set period.
5. The method for identifying the in-orbit vibration state of the satellite flexible accessory according to claim 4, further comprising a main vibration mode determining step of: judging whether the flexible accessory has only one main vibration mode in each vibration direction according to the overall displacement condition; if yes, continuing to execute the end displacement calculation step and the relative displacement calculation step; if not, returning to the step of executing the filtering and adjusting the concerned vibration frequency band.
6. The method for identifying the in-orbit vibration state of the satellite flexible accessory according to the gyro data of claim 3, wherein in the interested vibration frequency band setting step, the interested vibration frequency band in the i-axis direction is set according to the calculation result of the ground whole satellite structure dynamics analysis.
7. The method for identifying the in-orbit vibration state of the satellite flexible accessory according to claim 1, wherein m isi(t) is obtained by the following calculation formula:
mi(t)＝diηi(t)。
8. a system for identifying the on-orbit vibration state of a satellite flexible accessory by utilizing gyro data is characterized by comprising the following modules:
an attitude angular velocity acquisition module: obtaining attitude angular velocity measurement data omega corresponding to i axis of under-damped free vibration area after satellite on-orbit air injection closed-loop controli(t), wherein the i-axis is any one of an X-axis, a Y-axis and a Z-axis in the space rectangular coordinate system;
a filtering module: for omegai(t) filtering to obtain filtered i-axis attitude angular velocity data omega'i(t)；
A time series calculation module: according to ω'i(t) calculation ofObtaining a time series η of modal variablesi(t)；
Integral displacement acquisition module according to ηi(t) and vibrational mode matrix data DiCalculating to obtain the integral displacement condition of each node on the flexible accessory;
end displacement calculation Module according to ηi(t) and end node matrix data diAnd calculating the on-orbit vibration end displacement m of the satellite flexible accessoryi(t)；
Relative displacement calculation module according to ηi(t) calculating the array data of any two nodes on the flexible accessory to obtain the relative displacement m of any two nodes on the flexible accessory12(t)；
The time series calculation module comprises:
first derivative transformation: according to the following formula to ω'i(t) carrying out treatment:
in the formula:
the first derivative of the modal variable corresponding to the i axis;
Jithe moment of inertia of the satellite in the direction of the i axis;
Bithe rotational coupling coefficient of the vibration mode of the flexible accessory corresponding to the i axis;
integration: to pair
Integral calculation is carried out to obtain ηi(t)；
Relative displacement m12(t) is obtained by the following calculation formula:
m12(t)＝(d1-d2)ηi(t)；
in the formula: d1、d2The two nodes on the flexible accessory corresponding to the relative displacement are array data respectively.
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