CN109612664A - The in-orbit vibrational state method and system of satellite flexible appendage is recognized using gyro data - Google Patents
The in-orbit vibrational state method and system of satellite flexible appendage is recognized using gyro data Download PDFInfo
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- CN109612664A CN109612664A CN201910017079.1A CN201910017079A CN109612664A CN 109612664 A CN109612664 A CN 109612664A CN 201910017079 A CN201910017079 A CN 201910017079A CN 109612664 A CN109612664 A CN 109612664A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M7/00—Vibration-testing of structures; Shock-testing of structures
- G01M7/02—Vibration-testing by means of a shake table
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M7/00—Vibration-testing of structures; Shock-testing of structures
- G01M7/02—Vibration-testing by means of a shake table
- G01M7/025—Measuring arrangements
Abstract
The present invention provides a kind of methods using the gyro data identification in-orbit vibrational state of satellite flexible appendage comprising the steps of: attitude angular velocity obtaining step: the corresponding attitude angular velocity measurement data of i axis in the underdamping free vibration area after obtaining the in-orbit jet closed-loop control of satellite；Filter step: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；Time series calculates step: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；Global displacement obtaining step；End displacement calculates step；Relative displacement calculates step.Correspondingly, the present invention also provides a kind of systems using the gyro data identification in-orbit vibrational state of satellite flexible appendage.This method does not need other installation sensors on flexible appendage, is analyzed and processed merely with the measurement data of the existing attitude sensor of satellite platform.
Description
Technical field
The present invention relates to satellite monitoring fields, and in particular, to a kind of to be existed using gyro data identification satellite flexible appendage
The method and system of rail vibrational state.More particularly to a kind of in-orbit identification method for flexible satellite vibrational state information be
System, global displacement, end displacement and the discrimination method of relative displacement during the specifically a kind of pair of in-orbit vibration of flexible satellite
With system.
Background technique
With the continuous development of space technology, large space structure has been an important development side in space industry
To also will being the infrastructure [in-orbit spacecraft dynamics parameter identification technique research-Yu Dengyun] of space development indispensability.For
Flexible satellite structural dynamic characteristics are obtained, calculates and vibrates frequently with the methods of finite element modeling, ground experiment and in-orbit identification
Modal parameter.Model simplification, condition hypothesis in finite element modeling process etc. affect model accuracy, are especially difficult to hinge
Etc. contacts mechanism Accurate Model.Due to the influence of the factors such as gravity and atmospheric drag, Large Flexible Structure be difficult ground into
Row assembles and implements full-scale dynamic parameters identification test.Therefore, for the spacecraft with large-scale flexible appendage, it is difficult to
Accurate structural dynamic parameter is obtained by finite element modeling or ground experiment.Based on above-mentioned factor, to sun battle array, space exhibition
It is very urgent and necessary while also with higher that the big flexible structure such as Kaitian's line, which carries out in-orbit Modal Parameter Identification research,
Theory significance and practical application value.Flexible satellite structural dynamic parameter includes modal frequency, modal damping, formation, coupling
Coefficient etc., these parameters have important physical significance, can for the structure design of space flexible part, monitoring structural health conditions,
The application of structure failure diagnosis, structural vibration control etc. provides necessary support [CN102982196A].
The Modal Parameters Identification in Structural Dynamics field mainly includes frequency domain method, time domain approach and in recent years at present
Three kinds of the time-frequency domain method of rise.Most methods are needed using in-orbit excitation and placement sensor obtains letter on flexible appendage
Breath, such as [large-sized solar windsurfing modal parameter in-orbit identification research _ Li Xiaoran-Master's thesis] has studied senses on flexible appendage
The minimal configuration quantity and optimal location scheme of device；[CN105486474A] describes a kind of in-orbit mode of satellite flexible part
The realization system and method for identification needs to carry out pulse excitation to flexible appendage, receives and monitor the impulse response of each measuring point
Signal, and acquire acceleration responsive signal caused by the in-orbit steady state operation Satellite flexible part of satellite；
[CN106557633A] and [CN107609296A] describes the two methods of satellite sun array sensor placement.
The solar sail deployed configuration dynamic characteristic in-orbit identification method of [CN106408570A] based on Binocular vision photogrammetry, from image directly
Structural vibration displacement information is extracted, and then obtains the dynamic characteristic of structure in real time by operational modal analysis technology, realizes structure
The in-orbit identification of dynamic characteristic.Also two patents propose the attitude angular velocity data using gyro to measure to flexible satellite mode
The in-orbit identification method of parameter: [CN103970964A] describes a kind of flexible satellite modal parameter in-orbit identification method, needs
The angular velocity information of torque and flexible satellite body relative inertness coordinate system that executing agency is applied on flexible satellite body is collected,
The transmission function of modal parameter with torque to angular speed is obtained using Subspace Identification algorithm.[CN105157728A] proposes one
It kind can inhibit the flexible satellite Modal Parameters Identification of gyro noise influence, satellite body when flight in-orbit also with satellite
The measurement data of angular speed recognizes the modal frequency and damping ratios parameter of whole star, carries out difference processing inhibition to gyro data
Identification Errors caused by this two parts noise.
There are two constraints in terms of engineer application for the above in-orbit dynamic parameters identification method of flexible satellite: first is that in-orbit boat
Its device is difficult to required for applying dynamic parameters identification known excitation, can only utilize the expansion and gathering of in-orbit space structure,
The docking and separation of structure, the igniting of engine etc. generate driving source and are motivated, and these excitation source signals are all difficult to measure
[in-orbit spacecraft dynamics parameter identification technique research-Yu Dengyun].Second is that installing the number of vibrating sensor on flexible appendage
Amount receives the limitation of engineering construction.General flexible appendage is required to in-orbit expansion, all kinds of speed, accelerations, displacement sensor
Installation fixation, cable laying etc. have an adverse effect to unfolding mechanism, increase design difficulty.In addition, for sun battle array
Equal rotational flexibilities component also needs to increase driving mechanism slip ring signal number of channels.
Summary of the invention
For the defects in the prior art, it is attached using gyro data identification satellite flexibility that the object of the present invention is to provide a kind of
The method and system of the in-orbit vibrational state of part.
There is provided according to the present invention using gyro data identification the in-orbit vibrational state of satellite flexible appendage method, comprising with
Lower step:
Attitude angular velocity obtaining step: the i axis in the underdamping free vibration area after obtaining the in-orbit jet closed-loop control of satellite
Corresponding attitude angular velocity measurement data ωi(t), wherein i axis is X-axis, Y-axis, any in Z axis in rectangular coordinate system in space
A axis；
Filter step: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；
Time series calculates step: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；
Global displacement obtaining step: according to ηi(t) with mode of oscillation formation data Di, it is calculated on flexible appendage and respectively ties
The global displacement situation of point；
End displacement calculates step: according to ηi(t) with end node formation data di, satellite flexible appendage is calculated and exists
Rail vibrates end displacement mi(t)；
Relative displacement calculates step: according to ηi(t) it with the formation data of any two node on flexible appendage, is calculated and scratches
The relative displacement m of any two node on property attachment12(t)。
Preferably, in attitude angular velocity obtaining step, ωiIt (t) is that underdamping is freely shaken after the in-orbit gas puff Z-pinch of satellite
It in dynamic period, is obtained by satellite gyroscope measurement.
Preferably, the filter step comprises the steps of:
Concern vibration frequency band setting steps: the concern vibration frequency band of setting i axis direction is [f1,f2]Hz；
Strip step: enable bandpass filter to ωi(t) ω is removed in row filtering processingiIt (t) is more than the period of setting value in
Movement and high-frequency vibration component only retain the vibration information closed in vibration injection band, obtain ω 'i(t)。
Preferably, in global displacement obtaining step, by ηi(t) with mode of oscillation formation data DiIt is multiplied, obtains flexible appendage
Global displacement situation of the upper each node in setting period.
Preferably, also include principal oscillation mode judgment step: according to the global displacement situation, judging flexible appendage every
Whether only has a principal oscillation mode on a direction of vibration；If so, continuing to execute end displacement calculates step and relative displacement
Calculate step；It executes filter step if it is not, then returning and adjusts concern vibration frequency band.
Preferably, in concern vibration frequency band setting steps, according to the whole star structural dynamical model calculated result in ground, setting
Frequency band is vibrated in the concern of i axis direction.
Preferably, time series calculates step and comprises the steps of:
First derivative step of converting: according to the following formula to ω 'i(t) it is handled:
In formula:For the first derivative of the corresponding mode variable of i axis；
JiFor satellite i axis direction rotary inertia；
BiFor the rotation coefficient of coup of the corresponding flexible appendage mode of oscillation of i axis；
Integration step: rightIntegral calculation is carried out, η is obtainedi(t)。
Preferably, mi(t) it is obtained by following calculation formula:
mi(t)=diηi(t)。
Preferably, m12(t) it is obtained by following calculation formula:
m12(t)=(d1-d2)ηi(t)
In formula: d1、d2The formation data of two nodes on the corresponding flexible appendage of respectively required relative displacement.
The present invention also provides a kind of systems using the gyro data identification in-orbit vibrational state of satellite flexible appendage, include
With lower module:
Attitude angular velocity obtains module: the i axis in the underdamping free vibration area after obtaining the in-orbit jet closed-loop control of satellite
Corresponding attitude angular velocity measurement data ωi(t), wherein i axis is X-axis, Y-axis, any in Z axis in rectangular coordinate system in space
A axis；
Filter module: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；
Time series computing module: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；
Global displacement obtains module: according to ηi(t) with mode of oscillation formation data Di, it is calculated on flexible appendage and respectively ties
The global displacement situation of point；
End displacement computing module: according to ηi(t) with end node formation data di, satellite flexible appendage is calculated and exists
Rail vibrates end displacement mi(t)；
Relative displacement computing module: according to ηi(t) it with the formation data of any two node on flexible appendage, is calculated and scratches
The relative displacement m of any two node on property attachment12(t)。
Compared with prior art, the present invention have it is following the utility model has the advantages that
1, meet the needs of in-orbit structural dynamic characteristics of flexible satellite recognize, extract the in-orbit vibration end position of flexible satellite
Move information.
2, compared with other calculate the method for the in-orbit structural dynamic parameter of satellite, this method is not needed on flexible appendage
Other installation sensors, are analyzed and processed merely with the measurement data of the existing attitude sensor of satellite platform.
3, increased design of satellites, manufacture difficulty and wind in orbit are avoided because installing other vibrating sensors
Danger.
4, the flexible appendage vibrational state information picked out can be used for the in-orbit structural dynamical model of satellite, fault diagnosis,
Performance Evaluation etc..
Detailed description of the invention
Upon reading the detailed description of non-limiting embodiments with reference to the following drawings, other feature of the invention,
Objects and advantages will become more apparent upon:
Fig. 1 is certain single-blade sun battle array satellite configuration schematic diagram；
Fig. 2 is the attitude angular velocity change curve in satellite in-orbit jet closed-loop control period；
Fig. 3 is the time domain comparison diagram of measuring satellite angular velocities filtering front and back；
Fig. 4 is the frequency domain comparison diagram of measuring satellite angular velocities filtering front and back；
Fig. 5 is the change curve of flexible mode variable；
Fig. 6 is the formation schematic diagram of flexible appendage mode of oscillation；
Fig. 7 is the end displacement change curve of flexible appendage vibration；
Fig. 8 is the end displacement change curve of the vibration of two nodes on flexible appendage；
Fig. 9 is the method flow diagram using the gyro data identification in-orbit vibrational state of satellite flexible appendage.
Specific embodiment
The present invention is described in detail combined with specific embodiments below.Following embodiment will be helpful to the technology of this field
Personnel further understand the present invention, but the invention is not limited in any way.It should be pointed out that the ordinary skill of this field
For personnel, without departing from the inventive concept of the premise, various modifications and improvements can be made.These belong to the present invention
Protection scope.
In the description of the present invention, it is to be understood that, term " on ", "lower", "front", "rear", "left", "right", " perpendicular
Directly ", the orientation or positional relationship of the instructions such as "horizontal", "top", "bottom", "inner", "outside" is orientation based on the figure or position
Relationship is set, is merely for convenience of description of the present invention and simplification of the description, rather than the device or element of indication or suggestion meaning are necessary
It with specific orientation, is constructed and operated in a specific orientation, therefore is not considered as limiting the invention.
As shown in figure 9, the method provided by the invention using the gyro data identification in-orbit vibrational state of satellite flexible appendage,
Comprise the steps of: attitude angular velocity obtaining step: the underdamping free vibration area after obtaining the in-orbit jet closed-loop control of satellite
The corresponding attitude angular velocity measurement data ω of i axisi(t), wherein i axis is X-axis in rectangular coordinate system in space, Y-axis, in Z axis
Any one axis；Filter step: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；Time
Sequence calculates step: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；Global displacement obtaining step: according to
ηi(t) with mode of oscillation formation data Di, the global displacement situation of each node on flexible appendage is calculated；End displacement calculates
Step: according to ηi(t) with end node formation data di, the in-orbit vibration end displacement m of satellite flexible appendage is calculatedi(t)；
Relative displacement calculates step: according to ηi(t) it with the formation data of any two node on flexible appendage, is calculated on flexible appendage
The relative displacement m of any two node12(t).Preferably, in attitude angular velocity obtaining step, ωiIt (t) is the in-orbit jet of satellite
It in underdamping free vibration period after control, is obtained by satellite gyroscope measurement.That is, the identification in-orbit vibration of flexible appendage
The gyro to measure information that relative displacement is loaded merely with satellite platform, without using displacement, speed, the acceleration installed on flexible appendage
Spend vibrating sensor.
The filter step comprises the steps of: concern vibration frequency band setting steps: the concern vibration frequency of setting i axis direction
Band is [f1,f2]Hz；Strip step: enable bandpass filter to ωi(t) ω is removed in row filtering processingiIt (t) is more than setting value in
Periodic motion and high-frequency vibration component, only retain close vibration injection band in vibration information, obtain ω 'i(t).That is,
In strip step, attitude angular velocity measurement data is filtered using bandpass filter, removes the length of satellite body
Periodic motion and other high-frequency vibration error components.Preferably, in concern vibration frequency band setting steps, according to the whole star structure in ground
Frequency band is vibrated in the concern of dynamic analysis calculated result, setting i axis direction.
In embodiment, in global displacement obtaining step, by ηi(t) with mode of oscillation formation data DiIt is multiplied, obtains flexible attached
Global displacement situation of each node in setting period on part；Utilize the gyro data identification in-orbit vibrational state of satellite flexible appendage
Method also includes principal oscillation mode judgment step: according to the global displacement situation, judging flexible appendage in each direction of vibration
On whether only have a principal oscillation mode；If so, continuing to execute, end displacement calculates step and relative displacement calculates step；If
It is no, then it returns and executes filter step and adjust concern vibration frequency band.
It includes first derivative step of converting and integration step that time series, which calculates step,.Wherein, first derivative step of converting:
According to the following formula to ω 'i(t) it is handled:
In formula:For the first derivative of the corresponding mode variable of i axis；JiFor satellite i axis direction rotary inertia；Bi
For the rotation coefficient of coup of the corresponding flexible appendage mode of oscillation of i axis.Integration step: rightIntegral calculation is carried out, η is obtainedi
(t)。
Preferably, mi(t) it is obtained by following calculation formula:
mi(t)=diηi(t)。
Preferably, m12(t) it is obtained by following calculation formula:
m12(t)=(d1-d2)ηi(t)
In formula: d1、d2The formation data of two nodes on the corresponding flexible appendage of respectively required relative displacement.
Correspondingly, it is using the gyro data identification in-orbit vibrational state of satellite flexible appendage the present invention also provides a kind of
System, comprising with lower module: attitude angular velocity obtains module: the underdamping free vibration after obtaining the in-orbit jet closed-loop control of satellite
The corresponding attitude angular velocity measurement data ω of the i axis in areai(t), wherein i axis is X-axis, the Y-axis, Z axis in rectangular coordinate system in space
In any one axis；Filter module: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；When
Between sequence computing module: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；Global displacement obtains module: root
According to ηi(t) with mode of oscillation formation data Di, the global displacement situation of each node on flexible appendage is calculated；End displacement meter
Calculate module: according to ηi(t) with end node formation data di, the in-orbit vibration end displacement m of satellite flexible appendage is calculatedi
(t)；Relative displacement computing module: according to ηi(t) it with the formation data of any two node on flexible appendage, is calculated flexible attached
The relative displacement m of any two node on part12(t).Preferably, it is obtained in module in attitude angular velocity, ωiIt (t) is that satellite is in-orbit
It in underdamping free vibration period after gas puff Z-pinch, is obtained by satellite gyroscope measurement.That is, identification flexible appendage is in-orbit
The gyro to measure information that is loaded merely with satellite platform of vibration relative displacement, without using installed on flexible appendage displacement, speed,
Acceleration vibrating sensor.
The filter module includes with lower module: concern vibration frequency band setup module: frequency is vibrated in the concern of setting i axis direction
Band is [f1,f2]Hz；Strip module: enable bandpass filter to ωi(t) ω is removed in row filtering processingiIt (t) is more than setting value in
Periodic motion and high-frequency vibration component, only retain close vibration injection band in vibration information, obtain ω 'i(t).That is,
In strip module, attitude angular velocity measurement data is filtered using bandpass filter, removes the length of satellite body
Periodic motion and other high-frequency vibration error components.Preferably, in concern vibration frequency band setup module, according to the whole star structure in ground
Frequency band is vibrated in the concern of dynamic analysis calculated result, setting i axis direction.
In embodiment, global displacement is obtained in module, by ηi(t) with mode of oscillation formation data DiIt is multiplied, obtains flexible attached
Global displacement situation of each node in setting period on part；Utilize the gyro data identification in-orbit vibrational state of satellite flexible appendage
System also includes principal oscillation mode judgment module: according to the global displacement situation, judging flexible appendage in each direction of vibration
On whether only have a principal oscillation mode；If so, continuing to execute end displacement computing module and relative displacement computing module；If
It is no, then it returns and executes filter module and adjust concern vibration frequency band.
Time series computing module includes first derivative conversion module and integration module.Wherein, first derivative conversion module:
According to the following formula to ω 'i(t) it is handled:
In formula:For the first derivative of the corresponding mode variable of i axis；JiFor satellite i axis direction rotary inertia；Bi
For the rotation coefficient of coup of the corresponding flexible appendage mode of oscillation of i axis.Integration module: rightIntegral calculation is carried out, η is obtainedi
(t)。
Preferably, mi(t) it is obtained by following calculation formula:
mi(t)=diηi(t)。
Preferably, m12(t) it is obtained by following calculation formula:
m12(t)=(d1-d2)ηi(t)
In formula: d1、d2The formation data of two nodes on the corresponding flexible appendage of respectively required relative displacement.
Preferred embodiment:
Certain remote sensing satellite is equipped with single-blade solar battery array, and configuration is shown in Fig. 1.After transmitting is entered the orbit, during gas puff Z-pinch
It is as shown in Figure 2 to roll X-axis attitude angular velocity measurement data.
Step 1: choosing the rolling X-axis attitude angular velocity measurement data ω in satellite period in orbitX(t) it is analyzed
(Fig. 2).
Step 2: according to the whole star structural dynamical model calculated result in ground, the frequency of X-direction principal oscillation mode is about
0.37Hz, the vibration frequency band that concern is arranged is [0.259,0.481] Hz, is existed using the fertile hereby bandpass filter of 5 rank Barts to satellite
Attitude angular velocity measurement data ω run time railX(t) it is filtered, removing rolls in X-axis attitude angular velocity signal
The phugoid mode of motion and high-frequency vibration component only retain the vibration information in concern frequency band, obtain filtered rolling X-axis attitude angle
Speed omegaX′(t)。
Fig. 3 and Fig. 4 are shown in the time domain comparison and power spectral density comparison of attitude angular velocity measurement data filtering front and back respectively.From
Time domain comparison and frequency domain comparison can be clearly seen, and the signal within [0.259,0.481] Hz frequency band is retained, remaining frequency range
Signal is substantially weakened.
Step 3: according to formulaBy filtered single-axis attitude angular velocity omegaX' (t) is converted into mode
The first derivative of variableWherein, JX=6872.13kgm2It is satellite in the rotary inertia for rolling X-direction, BX=
33.48762m·kg1/2For the rotation coefficient of coup of flexible appendage mode of oscillation.Again to the first derivative of mode variableInto
Row integral, obtains the time series η of mode variableX(t), see Fig. 5.
Step 4: by the formation data D of mode of oscillationX(Fig. 6) and mode variable ηX(t) it is multiplied, can be obtained on flexible appendage
Vibration displacement situation of each node in this period.It is made the difference using the vibration displacement of two node any on flexible appendage available
The relative displacement of any two node on flexible appendage.
Step 5: bringing the formation d of end node intoXThe flexible appendage in (Fig. 6) this section of satellite can be obtained period in orbit
End displacement: mX(t)=dXηX(t).In this example, the formation of end node is dX=0.277kg-1/2, obtained end displacement
As a result see Fig. 7.
Step 6: optional two node, formation are respectively d on flexible appendage1And d2, then relative displacement calculated result is m12
(t)=(d1-d2)ηX(t), as a result see Fig. 8.
One skilled in the art will appreciate that in addition to realizing system provided by the invention in a manner of pure computer readable program code
It, completely can be by the way that method and step be carried out programming in logic come so that provided by the invention other than system, device and its modules
System, device and its modules are declined with logic gate, switch, specific integrated circuit, programmable logic controller (PLC) and insertion
The form of controller etc. realizes identical program.So system provided by the invention, device and its modules may be considered that
It is a kind of hardware component, and the knot that the module for realizing various programs for including in it can also be considered as in hardware component
Structure；It can also will be considered as realizing the module of various functions either the software program of implementation method can be Hardware Subdivision again
Structure in part.
Specific embodiments of the present invention are described above.It is to be appreciated that the invention is not limited to above-mentioned
Particular implementation, those skilled in the art can make various deformations or amendments within the scope of the claims, this not shadow
Ring substantive content of the invention.In the absence of conflict, the feature in embodiments herein and embodiment can any phase
Mutually combination.
Claims (10)
1. a kind of method using the gyro data identification in-orbit vibrational state of satellite flexible appendage, which is characterized in that comprising following
Step:
Attitude angular velocity obtaining step: the i axis in the underdamping free vibration area after obtaining the in-orbit jet closed-loop control of satellite is corresponding
Attitude angular velocity measurement data ωi(t), wherein i axis is X-axis, Y-axis, any of Z axis in rectangular coordinate system in space
Axis；
Filter step: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；
Time series calculates step: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；
Global displacement obtaining step: according to ηi(t) with mode of oscillation formation data Di, each node on flexible appendage is calculated
Global displacement situation；
End displacement calculates step: according to ηi(t) with end node formation data di, the in-orbit vibration of satellite flexible appendage is calculated
Dynamic end displacement mi(t)；
Relative displacement calculates step: according to ηi(t) it with the formation data of any two node on flexible appendage, is calculated flexible attached
The relative displacement m of any two node on part12(t)。
2. the method according to claim 1 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is, in attitude angular velocity obtaining step, ωi(t) it is underdamping free vibration period after the in-orbit gas puff Z-pinch of satellite, leads to
Cross what satellite gyroscope measurement obtained.
3. the method according to claim 1 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is that the filter step comprises the steps of:
Concern vibration frequency band setting steps: the concern vibration frequency band of setting i axis direction is [f1,f2]Hz；
Strip step: enable bandpass filter to ωi(t) ω is removed in row filtering processingiIt (t) is more than the periodic motion of setting value in
With high-frequency vibration component, only retains the vibration information closed in vibration injection band, obtain ω 'i(t)。
4. the method according to claim 3 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is, in global displacement obtaining step, by ηi(t) with mode of oscillation formation data DiIt is multiplied, obtains each node on flexible appendage
In the global displacement situation of setting period.
5. the method according to claim 4 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is, also includes principal oscillation mode judgment step: according to the global displacement situation, judging flexible appendage in each vibration side
Whether only has a principal oscillation mode upwards；If so, continuing to execute, end displacement calculates step and relative displacement calculates step；
It executes filter step if it is not, then returning and adjusts concern vibration frequency band.
6. the method according to claim 3 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is, in concern vibration frequency band setting steps, according to the whole star structural dynamical model calculated result in ground, i axis direction is arranged
Concern vibrate frequency band.
7. the method according to claim 1 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is that time series calculates step and comprises the steps of:
First derivative step of converting: according to the following formula to ω 'i(t) it is handled:
In formula:For the first derivative of the corresponding mode variable of i axis；
JiFor satellite i axis direction rotary inertia；
BiFor the rotation coefficient of coup of the corresponding flexible appendage mode of oscillation of i axis；
Integration step: rightIntegral calculation is carried out, η is obtainedi(t)。
8. the method according to claim 1 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is, mi(t) it is obtained by following calculation formula:
mi(t)=diηi(t)。
9. the method according to claim 1 using the gyro data identification in-orbit vibrational state of satellite flexible appendage, special
Sign is, m12(t) it is obtained by following calculation formula:
m12(t)=(d1-d2)ηi(t)
In formula: d1、d2The formation data of two nodes on the corresponding flexible appendage of respectively required relative displacement.
10. a kind of system using the gyro data identification in-orbit vibrational state of satellite flexible appendage, which is characterized in that comprising following
Module:
Attitude angular velocity obtains module: the i axis in the underdamping free vibration area after obtaining the in-orbit jet closed-loop control of satellite is corresponding
Attitude angular velocity measurement data ωi(t), wherein i axis is X-axis, Y-axis, any of Z axis in rectangular coordinate system in space
Axis；
Filter module: to ωi(t) it is filtered and obtains filtered i axis attitude angular velocity data ω 'i(t)；
Time series computing module: according to ω 'i(t) the time series η for obtaining mode variable is calculatedi(t)；
Global displacement obtains module: according to ηi(t) with mode of oscillation formation data Di, each node on flexible appendage is calculated
Global displacement situation；
End displacement computing module: according to ηi(t) with end node formation data di, the in-orbit vibration of satellite flexible appendage is calculated
Dynamic end displacement mi(t)；
Relative displacement computing module: according to ηi(t) it with the formation data of any two node on flexible appendage, is calculated flexible attached
The relative displacement m of any two node on part12(t)。
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910017079.1A CN109612664B (en) | 2019-01-08 | 2019-01-08 | Method and system for identifying on-orbit vibration state of satellite flexible accessory by utilizing gyroscope data |
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| CN201910017079.1A CN109612664B (en) | 2019-01-08 | 2019-01-08 | Method and system for identifying on-orbit vibration state of satellite flexible accessory by utilizing gyroscope data |
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| CN109612664B CN109612664B (en) | 2020-08-25 |
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