CN103674224B - The micro-method for testing vibration of a kind of solar wing driving mechanism - Google Patents
The micro-method for testing vibration of a kind of solar wing driving mechanism Download PDFInfo
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- CN103674224B CN103674224B CN201310446691.3A CN201310446691A CN103674224B CN 103674224 B CN103674224 B CN 103674224B CN 201310446691 A CN201310446691 A CN 201310446691A CN 103674224 B CN103674224 B CN 103674224B
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Abstract
The invention discloses the micro-method for testing vibration of a kind of solar wing driving mechanism, the present invention adopts air floating table to support as air pressure, overcome the impact of ground gravity on solar wing driving mechanism, during test, measurement mechanism is separated with measured test specimen, do not need, on test specimen, optional equipment and sensor are installed, do not affect the dynamic perfromance of test specimen, do not damage test specimen structure, ensure that the security of test specimen; (2) the present invention is by the reasonable Arrangement of eight Common piezoelectricity sensors, thus make the micro-disturbance signal of six-freedom degree that existing one-way piezoelectric force snesor can be utilized to measure, overcome the problem lacking high precision three-dimensional sensor, measuring accuracy is improved greatly.It is inner or outside that disturbing source of the present invention can be placed in measuring system, improves the adaptability of measuring system.
Description
Technical field
The present invention relates to a kind of micro-method for testing vibration, particularly relate to the micro-method for testing vibration of a kind of different gravity environment solar wing driving mechanism, belong to vibration test technical field, can be used for solar wing driving mechanism and simulate vibration characteristics kinetic measurement in-orbit.
Background technology
The principal element affecting satellite shake and attitude stability comprises the disturbing moment etc. of movable part on external force of environment interference, attitude maneuver, star.On satellite imagery device and star movable part cause shake time exist all the time, on the impact of image quality comparatively large, therefore it may be necessary corresponding control device and suppressed or isolate.On star, movable part mainly refers to satellite sun wing driving mechanism, as shown in Figure 1, away from the end of satellite body structure 1, a rotating shaft is installed at the satellite sun wing 3, then in rotating shaft, install ball bearing, rotating shaft and ball-bearing fit realize the satellite sun wing 3 and freely rotate for axle with solar wing driving mechanism (SADA) 2.In order to study at SADA(solar wing driving mechanism) etc. the lower satellite structure of rotary part excitation shake and flexible vibration propagation characteristic, and on the impact that useful load is pointed to, needing the disturbance characteristic to the satellite sun wing (the disturbing spectrum characteristic under various boundary) to measure and data analysis, laying the foundation for launching jitter suppression project study.Current, simulated solar wing driving mechanism in-orbit environment and directly measurement still to can yet be regarded as a kind of feasible and effective scheme.At present, there is not yet the reported in literature about the micro-method for testing vibration of this type of solar wing driving mechanism both at home and abroad.
Summary of the invention
The technical problem to be solved in the present invention is: overcome the deficiencies in the prior art, provides a kind of solar wing driving mechanism micro-method for testing vibration, ensure that the security of test specimen, substantially increase measuring accuracy.
Technical solution of the present invention is: the micro-method for testing vibration of a kind of solar wing driving mechanism, and step is as follows:
(1) micro-vibration test system is set up, this micro-vibration test system comprises support, air floating table, torque sensor, linear bearing, flange, micro-vibration six points of testing tables, Data collection and precessing system, backup system and fictitious load, air floating table is bolted on support, the rotating shaft of air floating table is connected with torque sensor, torque sensor is connected by linear bearing and flange and makes the axis of three in alignment, micro-vibration six points of testing tables to be arranged in support and to be linked together by framework and support, four piezoelectric sensors installed in the horizontal direction and four piezoelectric sensors vertically installed are furnished with in micro-vibration six points of testing tables, backup system is made up of air compressor and delivery pipe, the gases at high pressure that air compressor produces are transported in air floating table by delivery pipe and carry out gravity unloading for air floating table, fictitious load and air floating table are connected and ensure concentric, are used for the impact of test load on solar wing driving mechanism (6) micro-vibration characteristics,
(2) micro-vibration six points of testing tables are demarcated; Scaling method is: be first fixedly mounted on by demarcation dish on micro-vibration six points of testing tables, proof force is adopted to hammer 16 calibration points knocked on demarcation dish into shape, when knocking at every turn, Data collection and precessing system gathers the frequency response function of eight sensors of six points of testing tables respectively, modal theory and Method of generalized inverse calculation is utilized to obtain calibration matrix, obtain the output voltage of eight sensors and power and hammer transformational relation between Input Forces, the i.e. calibration coefficient of micro-vibration six points of testing tables into shape.
(3) lower end of solar wing driving mechanism to be arranged on calibrated micro-vibration six points of testing tables and to make solar wing driving mechanism and micro-vibration six points of testing tables keep the centre of form to overlap, the upper end of solar wing driving mechanism is fixed by flange, and fictitious load is arranged on the top of air floating table;
(4) inflated to air floating table by backup system, air floating table is started working, the gravity of fictitious load is unloaded;
(5) test starts, control solar wing driving mechanism and produce small sample perturbations and rotating shaft disturbance, eight piezoelectric sensors that micro-vibration six points of testing tables are installed are measured small sample perturbations, and torque sensor and the disturbance of grating angular displacement sensor countershaft are measured;
(6) Data collection and precessing system is measured piezoelectric sensor small sample perturbations signal, the torque signal of torque sensor measurement and the angular displacement signal of grating angular displacement sensor measurement carry out acquisition process, obtain micro-vibration characteristics of solar wing driving mechanism.
The present invention compared with prior art has the following advantages:
(1) the present invention adopts air floating table to support as air pressure, overcome the impact of ground gravity on solar wing driving mechanism, during test, measurement mechanism is separated with measured test specimen, do not need, on test specimen, optional equipment and sensor are installed, do not affect the dynamic perfromance of test specimen, do not damage test specimen structure, ensure that the security of test specimen.
(2) the present invention is by the reasonable Arrangement of eight Common piezoelectricity sensors, thus make the micro-disturbance signal of six-freedom degree that existing one-way piezoelectric force snesor can be utilized to measure, overcome the problem lacking high precision three-dimensional sensor, measuring accuracy is improved greatly.
(3) disturbing source of the present invention can be placed in measuring system inside or outside, improves the adaptability of measuring system.
Accompanying drawing explanation
Fig. 1 is test flow chart of the present invention;
Fig. 2 is the composition structural drawing of test macro;
Fig. 3 is the calibration principle figure of the micro-vibration of the present invention six points of testing tables.
Embodiment
As shown in Figure 1, first the present invention sets up micro-vibration test system as shown in Figure 2, this test macro comprises: support 1, air floating table 2, torque sensor 3, linear bearing 4, flange 5, micro-vibration six points of testing tables 7, Data collection and precessing system 8, backup system 9 and fictitious load 10, air floating table 2 is bolted on support 1, the rotating shaft 201 of air floating table 2 is connected with torque sensor 3, torque sensor 3 is connected by linear bearing 4 and flange 5 and is made the axis of three in alignment, linear bearing 4 is for transmitting the moment of torsion in solar wing driving mechanism 6Z direction, micro-vibration six points of testing tables 7 to be arranged in support 1 and to be linked together by framework 11 and support 1, the lower end of solar wing driving mechanism 6 is fixed by micro-vibration six points of testing tables 7, the upper end of solar wing driving mechanism 6 is fixed by flange 5, four piezoelectric sensors installed in the horizontal direction and four piezoelectric sensors vertically installed are furnished with in micro-vibration six points of testing tables 7, when solar wing driving mechanism produces vibration, piezoelectric sensor produces voltage signal, for measuring the vibration that solar wing driving mechanism 6 produces, the output signal of eight piezoelectric sensors is carried out acquisition process by Data collection and precessing system 8 and is converted into three microvibration force signals and three microvibration torque signals, for analyzing the vibration characteristics of the installation interface of solar wing driving mechanism, backup system 9 is made up of air compressor and delivery pipe, the gases at high pressure that air compressor produces are transported in air floating table 2 by delivery pipe, ensure that the air pressure in air floating table 2 enough supports gravity unloading, fictitious load 10 and air floating table 2 are connected and ensure concentric, for applying acting force to solar wing driving mechanism after air floating table 2 gravity unloading.When carrying out vibration test, air floating table by torque sensor, linear bearing, flange successively in line and connected by bolt, be connected with the solar wing driving mechanism output shaft of the rigidity six points of testing tables being arranged in frame bottom simultaneously, fictitious load reaches ground by air floating table, supporting structure, solar wing driving mechanism is not had an impact, solar wing driving mechanism is as disturbing source, installation interface disturbance is recorded by micro-vibration six points of testing tables, and rotating shaft output disturbance is recorded by torque sensor and the grating angular displacement sensor that is arranged in air floating table.Data acquisition equipment is LMSTestlab and the computer equipment of Belgian LMS company, backup system is air compressor, for gravity uninstalling system provides source of the gas support, test by loading the fictitious load of different stage, can the disturbance characteristic of simulated solar wing driving mechanism environment in-orbit.
Support 1 is the spatial structure be made up of four foundation plates, four upper plates, four vertical pillars.Air floating table 2 is arranged on the center of support 1 upper plate.Air floating table 2 includes a grating angular displacement sensor 202, measures corner displacement and the moment of torsion in solar wing driving mechanism 6Z direction with torque sensor synchro measure; The output signal of grating angular displacement sensor 202 carries out acquisition process by Data collection and precessing system 8, and the model of grating angular displacement sensor is RESM20USA200, external diameter 200mm, internal diameter 180mm, pitch 20m, groove number 31488.Torque sensor adopts model XSM/A-HIMT2A3B3V0N, the torque sensor of range 200NM, precision 0.1%FS; Four piezoelectric sensors vertically installed, four piezoelectric sensors installed in the horizontal direction are piezo ceramic element, tighten, to improve the measuring accuracy of piezoelectric sensor in the scope that must allow in intensity as far as possible.A cavity is had for installing solar wing driving mechanism 6 in micro-vibration six points of testing tables 7.Fictitious load 10 is bolted to connection by steel tubular beam and square steel and forms.
(2) micro-vibration six points of testing tables are demarcated:
According to Modal Analysis Theory, N degree of freedom linear system is had:
X(ω)=H(ω)F(ω)(1)
In above formula, X (ω) for response spectra vector (voltage signal), dimension be N × 1; H (ω) is frequency response function matrix, and dimension is N × N; F (ω) is loading spectrum vector, and dimension is N × 1.Generally, real system is continuous structure particularly, and its number of degrees of freedom, N is very large, can not record the load that the response in all degree of freedom is subject to solve structure.Under normal circumstances, because charge number P to be determined can not be very large, so wish to determine load to be identified by few response data of trying one's best, namely identify load by the partial response of structure.Suppose that charge number undetermined is P, the measuring point number of response is L, and both are all less than total number of degrees of freedom, N of system, therefore: X (ω)
l × 1=H (ω)
l × Pf (ω)
p × 1(2)
As long as by the known frequency spectrum determining frequency response function matrix and response vector of above formula, just can solve and carry spectrum, and then obtain the time-domain signal of load with inverse fourier transform, the load identification inverse matrix of frequency response namely in Modal Analysis Theory.If charge number P undetermined in formula (2) is i.e. L=P equal to the measuring point number L of response, then frequency response function matrix H (ω) is square formation, and now loading spectrum vector F (ω) can be tried to achieve by following formula:
F(ω)=H
-1(ω)X(ω)(3)
If the measuring point number L of charge number P undetermined and response is unequal, normally L >=P, then frequency response function matrix H (ω) is no longer just square formation, and must ask generalized inverse to frequency response function, like this, the formula of load identification is:
F(ω)=[H
H(ω)H(ω)]
-1H
H(ω)X(ω)(4)
In formula, the conjugate transpose of subscript H representing matrix.Usual dynamic response X (ω) becomes more readily available, and frequency response function matrix H (ω) really rule be not easy, because the reflection of each element is relation between the point of excitation of various discrete and response point in H (ω) matrix, response point is mutually different for each different point of excitation and point of excitation for the frequency response function between each different response point, that is, H (ω) in above formula and F (ω) is closely connected together, if do not know each component F in load vectors F (ω)
j(ω) active position, also just cannot determine each element in H (ω).Therefore, inverse matrix of frequency response can only be used for the dynamic load identifying known action position, then can not identify in this way for active position the unknown or time dependent situation.
For some disturbing source (as momenttum wheel), its disturbing force application point can not accurately be determined.Two problems can be run into: 1) how to define disturbing force application point during inverse matrix of frequency response at this moment before application described in face; 2) because the disturbing force application point for definition often cannot direct imposed load, how transfer matrix demarcation is carried out to this application point.
For above-mentioned two problems, solution of the present invention is:
(1) the disturbing source mounting disc that processing one is relatively firm, mounting disc inherently frequency is greater than detection frequency more than 3 times, at this moment can think that mounting disc relative system other parts are rigid body;
(2) all disturbing forces are equivalent to six the external force load of application point in mounting disc, i.e. three translation power, two moments of flexure and moments of torsion;
(3) setting as being rigidly connected between demarcation power point of excitation in mounting disc and its central point O, directly demarcation power being applied to the different application point of this dish and different directions and demarcating.
According to foregoing description, load transfer first will be demarcated to the mounting disc centre of form, that is:
In above formula, the n in matrix subscript represents the number of times loaded in test; F represents the load of equivalence to the centre of form, and dimension is 6 × n; The load of F ' expression actual loaded is a diagonal matrix, in matrix
equal the load value loaded for i-th i time; C represents the transition matrix between the loading matrix of actual loaded and equivalent load matrix, and dimension is 6 × n.Pass between equivalent load and force sensor signals is:
W
6×8(ω)T
8×n(ω)=F
6×n(ω)(6)
N wherein in matrix subscript represents the number of times loaded in test; W is the inverse of system frequency response function matrix; T is the response signal of eight force snesor, and dimension is 8 × n; F is the equivalent load obtained in formula (5).From in formula (6), when response matrix T exists the inverse time, have:
W
6×8(ω)=F
6×n(ω)T
-1 n×8(ω)(7)
Consider that the passage of response signal only has 8, in order to improve measuring accuracy, the number of load(ing) point should be greater than response channel number, i.e. n>8, at this moment response matrix T is no longer a square formation, but the matrix of a row full rank, application Generalized Inverses Theory, has:
W
6 × 8(ω)=F
6 × n(ω) T
h(ω) [T (ω) T
h(ω)]
-1(8) (5) formula is substituted into (8), has:
The generalized inverse of the frequency response function matrix H (ω) in the T (ω) difference expression (4) that the matrix W (ω) of formula (9) being tried to achieve in test and test record and response X (ω), just can the equivalent external applied load of certainty annuity, realize center equivalence and demarcate the force signal that the voltage signal obtained by sensor is converted to actual needs, namely
F
6×1(ω)=W
6×8(ω)T
8×1(ω)(10)
Scaling method: in order to pumping signal F (ω) can be determined from testing the response signal T (ω) that obtain, first should try to achieve corresponding calibration matrix W (ω), integral calibrating matrix is actually the frequency response function matrix between the response signal of eight piezoelectric force transducers and the load acting on equivalent center point.In this test, equivalent center is the central point of the loading disk upper surface of eight component sensor devices, heart point directly applies three translation power and three moments are had any problem hereinto, the present invention is on the basis of rigid body at hypothesis loading disk, the calibrating table that a rigidity is very high is installed, and selects 16 load(ing) points as shown in Figure 3.According to space force system level theory, 16 load(ing) points selected are utilized can equivalent to go out to act on three translation power and three moments of loading disk geometric center.
| Lx/m | Ly/m | Lz/m |
| 0.042 | 0.042 | 0.04 |
Power F is demarcated by 16 in test
1~ F
16obtain the calibration matrix W (ω) of system, the expression formula of the Matrix C in the transition matrix between the loading matrix of the expression actual loaded in test and equivalent load matrix and formula (5) is:
C
1=[010-Lz0-Lx]
T
C
2=[010-Lz00]
T
C
3=[010-Lz0Lx]
T
C
4=[-1000-Lz-Ly]
T
C
5=[-1000-Lz0]
T
C
6=[-1000-LzLy]
T
C
7=[0-10Lz0-Lx]
T
C
8=[0-10Lz00]
T
C
9=[0-10Lz0Lx]
T
C
10=[1000Lz-Ly]
T
C
11=[1000Lz0]
T
C
12=[1000LzLy]
T
C
13=[00-1Ly-Lx0]
T
C
14=[00-1LyLx0]
T
C
15=[00-1-LyLx0]
T
C
16=[00-1-Ly-Lx0]
T
C in above formula
irepresent the transition matrix between load and center equivalent load loaded for i-th time.
The non-detailed description of the present invention is known to the skilled person technology.
Claims (4)
1. the micro-method for testing vibration of solar wing driving mechanism, is characterized in that step is as follows:
(1) micro-vibration test system is set up, this micro-vibration test system comprises support (1), air floating table (2), torque sensor (3), linear bearing (4), flange (5), micro-vibration six points of testing tables (7), Data collection and precessing system (8), backup system (9) and fictitious load (10), air floating table (2) is bolted on support (1), rotating shaft (201) the same to torque sensor (3) of air floating table (2) is connected, torque sensor (3) is connected by linear bearing (4) and flange (5) and is made the axis of three in alignment, micro-vibration six points of testing tables (7) are arranged in support (1) and are also linked together by framework (11) and support (1), four piezoelectric sensors installed in the horizontal direction and four piezoelectric sensors vertically installed are furnished with in micro-vibration six points of testing tables (7), backup system (9) is made up of air compressor and delivery pipe, the gases at high pressure that air compressor produces are transported in air floating table (2) by delivery pipe and carry out gravity unloading for air floating table (2), fictitious load (10) is connected with air floating table (2) and ensures concentric, is used for the impact of test load on solar wing driving mechanism (6) micro-vibration characteristics,
(2) micro-vibration six points of testing tables (7) are demarcated;
(3) lower end of solar wing driving mechanism (6) is arranged on calibrated micro-vibration six points of testing tables (7) go up and make solar wing driving mechanism (6) and micro-vibration six points of testing tables (7) keep the centre of form to overlap, the upper end of solar wing driving mechanism (6) is fixed by flange (5), and fictitious load (10) is arranged on the top of air floating table (2);
(4) inflated to air floating table (2) by backup system (9), air floating table (2) is started working, the gravity of fictitious load (10) is unloaded;
(5) test starts, control solar wing driving mechanism (6) and produce small sample perturbations and rotating shaft disturbance, upper eight piezoelectric sensors installed of micro-vibration six points of testing tables (7) are measured small sample perturbations, and torque sensor (3) and grating angular displacement sensor (202) countershaft disturbance are measured;
(6) torque signal that Data collection and precessing system (8) is measured piezoelectric sensor small sample perturbations signal, torque sensor are measured and the angular displacement signal measured of grating angular displacement sensor (2) carry out acquisition process, obtain micro-vibration characteristics of solar wing driving mechanism.
2. the micro-method for testing vibration of a kind of solar wing driving mechanism according to claim 1, it is characterized in that: the scaling method of described step (2) is: first demarcation dish is fixedly mounted on micro-vibration six points of testing tables (7), proof force is adopted to hammer 16 calibration points knocked on demarcation dish into shape, when knocking at every turn, Data collection and precessing system (8) gathers the frequency response function of eight sensors of six points of testing tables (7) respectively, modal theory and Method of generalized inverse calculation is utilized to obtain calibration matrix, obtain the output voltage of eight sensors and power and hammer transformational relation between Input Forces into shape, the i.e. calibration coefficient of micro-vibration six points of testing tables (7).
3. the micro-method for testing vibration of a kind of solar wing driving mechanism according to claim 1, is characterized in that: described air floating table (2) is arranged on the center of support (1) upper plate.
4. the micro-method for testing vibration of a kind of solar wing driving mechanism according to claim 1, is characterized in that: described linear bearing (4) is for transmitting the moment of torsion of solar wing driving mechanism (6) Z-direction.
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| CN110850845B (en) * | 2019-11-13 | 2020-09-25 | 上海航天控制技术研究所 | Space station solar wing simulated load test system |
| CN112444365B (en) * | 2020-11-30 | 2023-08-29 | 哈尔滨工业大学 | Satellite solar wing substrate unfolding low-frequency mode testing method |
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| CN101246063A (en) * | 2008-03-21 | 2008-08-20 | 北京航空航天大学 | Spacecraft spacing tiny perturbation load measuring system |
| CN101592518A (en) * | 2009-06-24 | 2009-12-02 | 北京航空航天大学 | A kind of high-precision microvibration measuring system |
| CN101881651A (en) * | 2010-06-18 | 2010-11-10 | 北京航空航天大学 | Small-size portable high-accuracy disturbance measuring system |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN101246063A (en) * | 2008-03-21 | 2008-08-20 | 北京航空航天大学 | Spacecraft spacing tiny perturbation load measuring system |
| CN101592518A (en) * | 2009-06-24 | 2009-12-02 | 北京航空航天大学 | A kind of high-precision microvibration measuring system |
| CN101881651A (en) * | 2010-06-18 | 2010-11-10 | 北京航空航天大学 | Small-size portable high-accuracy disturbance measuring system |
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