CN109029467A - A kind of spacecraft high-precision angular movement measurement method based on rotator type gyro biorthogonal configuration - Google Patents
A kind of spacecraft high-precision angular movement measurement method based on rotator type gyro biorthogonal configuration Download PDFInfo
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- CN109029467A CN109029467A CN201810947412.4A CN201810947412A CN109029467A CN 109029467 A CN109029467 A CN 109029467A CN 201810947412 A CN201810947412 A CN 201810947412A CN 109029467 A CN109029467 A CN 109029467A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
- G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation
Abstract
The present invention relates to a kind of spacecraft high-precision angular movement measurement methods based on rotator type gyro biorthogonal configuration.The analytical relation comprising spacecraft attitude angular speed and angular acceleration information is obtained using bonding force square suffered by gyrorotor radial direction, in the case where not ignoring the inertia coupling terms as caused by spacecraft angular movement and cross-couplings item, pass through the information fusion operation of multiple gyros in biorthogonal configuration, disappear inertia coupling terms and cross-couplings item, to combine to obtain the analytical expression of spacecraft angular speed and angular acceleration.Due to remaining inertia coupling terms and cross-couplings item, measurement accuracy is only dependent upon gyro error and configuration installation error, can't reduce with the increase of spacecraft dynamic range.The invention effectively overcomes the conspicuous contradiction in traditional attitude angular rate and angular acceleration measurement method between measurement accuracy and dynamic range.The invention belongs to technical field of inertial, can be applied to the high-precision high bandwidth measurement of spacecraft attitude angular speed and angular acceleration.
Description
Technical field
The present invention relates to a kind of spacecraft high-precision angular movement measurement method based on rotator type gyro biorthogonal configuration,
Suitable for the occasion using rotator type gyro configuration as spacecraft attitude angular movement information measurement.
Technical background
Rotor gyro mainly includes flexible gyroscope, magnetic floating gyro, liquid floated gyroscope, electrostatic gyroscope, three floating gyros, two floating tops
Spiral shell etc. is the first choice of current high accuracy inertial navigation system because of its measurement accuracy with higher.Rotator type gyro usually has one
A or two freedom degrees angular speed sensitive capability realizes spacecraft triaxial attitude angle using the orthogonal installation of three gyros
Speed measurement is a kind of general equipment arrangement of inertial navigation system.It is single in order to realize in this mounting means and method for solving
The attitude angular rate measurement capability of gyro usually ignores inertia coupling terms and cross-couplings item caused by spacecraft angular movement.
Under low dynamic condition, the inertia coupling terms and cross-couplings item ignored are events for gyro coupling terms, this
The influence that kind is ignored to the measuring precision is lower.But under high dynamic condition, shared by inertia coupling terms and cross-coupling
Large percentage, it is this ignore necessarily cause the measurement accuracy of spacecraft attitude angular speed to reduce, and dynamic higher, measurement accuracy
It is lower.The method for solving of posture angular acceleration usually carries out differential to attitude angular rate and obtains, this can not only introduce differential and make an uproar
Sound, and the solution of posture angular acceleration is also inevitably present the contradiction between precision and bandwidth.
Gyro Precision is improved by signal processing methods such as digital filterings, is a kind of method of relative maturity.How to utilize
More gyro informations carry out high-acruracy survey, are current a research hotspot and difficult point, existing research is concentrated mainly on more tops
In the data fusion of spiral shell information.Although what this method can improve attitude angular rate to a certain extent measures or estimates precision,
But it is inevitably present detection accuracy and detects the conspicuous contradiction between dynamic range.
Summary of the invention
Technology of the invention solves the problems, such as: angular speed and angular acceleration measurement accuracy for high dynamic spacecraft are asked
Topic proposes that one kind is based on turning under conditions of not ignoring the inertia coupling terms as caused by spacecraft angular movement and cross-couplings item
The high-precision high dynamic attitude angle motion measuring method of minor gyro biorthogonal configuration.The measurement accuracy of the method for the present invention only takes
Certainly in gyro error and configuration installation error, do not reduced with the increase of spacecraft dynamic range.The invention effectively overcomes
Spacecraft angle speed greatly improved in contradiction in existing method between attitude angular rate measurement accuracy and measurement dynamic range
The measurement accuracy of rate and angular acceleration within the scope of different frequency bands is the super steady super quiet and super quick posture and vibration control of spacecraft
System is laid a good foundation.
Technical solution of the invention:
Inertia coupling terms and cross-couplings item as caused by spacecraft angular movement in not ignoring rotor radial output torque
Under the premise of, it can measure to high bandwidth attitude angular velocity and the angle of spacecraft with high precision using 4 gyros of biorthogonal configuration
Acceleration, specifically includes the following steps:
(1) kinetic model of in-orbit gyrorotor is established
The torque of in-orbit n-th of gyrorotor It may be expressed as:
Wherein, n=1,2,3,4； Respectively represent the angular speed in gyro reference frame relative inertness space； Respectively represent the angular acceleration in gyro reference frame relative inertness space；Ir, IzRespectively represent gyrorotor diameter
To the rotary inertia with axial direction；Ω represents gyrorotor angular velocity of rotation； N-th of gyrorotor is represented to sit around stator
Mark system OXfYfZfMiddle OXf, OYfThe yaw rate of axis； N-th of gyrorotor is represented around stator coordinate OXfYfZfIn
OXf, OYfThe deflection angular acceleration of axis；iαn, iβnN-th of gyroscopic couple device coil drive rotor is represented around OXf, OYfAxis deflection
Electric current；kxAnd kyGyroscopic couple device is respectively indicated in OXf, OYfTorque coefficient on axis.
(2) gyro sensitivity torque variable is definedWith
Wherein,It can be carried out once or twice not by the gyrorotor displacement obtained to measurement
Complete differential obtains.
(3) spacecraft three-axis attitude angular speed is solved
Wherein,
(4) spacecraft three-axis attitude angular speed is solved
From (1) Shi Ke get, gyroscopic couple expression formula mainly includes containingInertia coupling terms, contain Intersection
Coupling terms, Yi JihanWithGyro coupling terms.In conventional methods where, it in order to solve attitude angular rate, directly neglects
Inertia coupling terms and cross-couplings item only remain gyro coupling terms, thus necessarily cause attitude angular rate measurement accuracy with
The increase of spacecraft dynamic frequency and reduce.In contrast, it can be obtained from the solution procedure of formula (2), (3), (4), side of the present invention
Method passes through biorthogonal configuration in the case where not ignoring the inertia coupling terms as caused by spacecraft angular movement and cross-couplings item
In multiple gyros information fusion operation, disappear inertia coupling terms and cross-couplings item, to combine to obtain spacecraft angular speed
And the analytical expression of angular acceleration.
Therefore, the measurement accuracy of the invention is only dependent upon gyro error and configuration installation error, and with spacecraft attitude machine
Dynamic dynamic range is unrelated, can greatly improve attitude angular velocity measurement accuracy of the spacecraft under high dynamic condition.Institute of the present invention
The rotator type gyro stated includes the energy such as flexible gyroscope, magnetic floating gyro, liquid floated gyroscope, electrostatic gyroscope, three floating gyros, two floating gyros
Enough obtain the rotator type gyroscope of gyrorotor displacement and torquer current information.
Biorthogonal configuration of the invention is made of 4 rotator type gyros, is divided into two groups, two gyros in each group
Straight line where rotary shaft is nominally directed toward spatially is mutually perpendicular to, can coplanar also antarafacial；Two gyro rotary shafts between two groups
Straight line where nominal direction is parallel to each other or collinearly, straight line where the nominal direction of the another two gyro between two groups is also mutually flat
It is capable or conllinear.
Inventive principle of the invention is: obtaining spacecraft attitude using bonding force square relational expression suffered by rotor gyro radial direction
The parsing relationship of the parameters such as angular speed, angular acceleration and rotor displacement, torquer electric current；It is merged using the information of gyro configuration,
Do not ignoring cross-coupling caused by the inertia coupling terms as caused by spacecraft angular acceleration and spacecraft attitude angular speed product
In the case where, the plus and minus calculation between current information amount, the ingenious inertia coupling terms that disappear are displaced by multiple gyrorotors
And cross-couplings item, the analytical expression of spacecraft attitude angular speed and angular acceleration is obtained, to realize spacecraft attitude angle
The high-precision high dynamic of motion information measures.
Since the method for the present invention has been effectively retained systematic error caused by cross-couplings item and inertia coupling terms, survey
Accuracy of measurement is only dependent upon gyroscope itself error and configuration installation error, and angular velocity measurement precision can't be dynamic with spacecraft
The increase of state range and reduce, remain at higher level.
The solution of the present invention and existing scheme ratio, major advantage are: the measurement accuracy of attitude angular rate and angular acceleration
It can't be reduced with the increase of spacecraft dynamic range, effectively overcome measurement accuracy and dynamic range in conventional method
Between contradiction；The angular acceleration of available parsing expresses formula, avoids conventional method because of introduced system of differentiating
Noise.
Detailed description of the invention
Fig. 1 specific embodiment figure；
Fig. 2 " ten " word configuration installation diagram；
The orthogonal configuration installation diagram of Fig. 3 tri-；
0.1Hz disturbs the attitude angular rate measurement of comparison result of lower distinct methods when Fig. 4 gyro free error and installation error
Analogous diagram；
The attitude angular rate measurement of comparison result that 1Hz disturbs lower distinct methods when Fig. 5 gyro free error and installation error is imitated
True figure；
0.1Hz disturbs the attitude angular rate measurement of comparison result of lower distinct methods when Fig. 6 has gyro error and installation error
Analogous diagram；
The attitude angular rate measurement of comparison result that 1Hz disturbs lower distinct methods when Fig. 7 has gyro error and installation error is imitated
True figure.
Specific embodiment
Specific embodiments of the present invention are as shown in Figure 1, specific implementation step is as follows:
(1) kinetic model of in-orbit single gyrorotor is established
As shown in Fig. 2, G1、G2、G3、G44 rotator type gyros, are divided into two groups, G1And G2One group, G3And G4It is one group, each
Straight line where the rotary shaft of two gyros in group is nominally directed toward spatially is mutually perpendicular to and coplanar；G between two groups1And G3Two
Straight line where a gyro rotary shaft is nominally directed toward is conllinear, the G between two groups2And G4Straight line is also total where the nominal direction of two gyros
Line constitutes " ten " word configuration, it is clear that this is a kind of special shape of biorthogonal configuration；XbYbZbIt is satellite body coordinate system, G1
And G3Angular momentum be nominally directed toward and ObZbAnd-ObZbOverlapping of axles, G2And G4Angular momentum be nominally directed toward and ObXbAnd-ObXbAxis
It is overlapped.
The radial output torque of in-orbit n-th of gyrorotor It may be expressed as:
Wherein, n=1,2,3,4； Respectively represent the angular speed in gyro reference frame relative inertness space； Respectively represent the angular acceleration in gyro reference frame relative inertness space；Ir, IzRespectively represent gyrorotor diameter
To the rotary inertia with axial direction；Ω represents gyrorotor angular velocity of rotation； N-th of gyrorotor is represented around stator coordinate
It is OXfYfZfMiddle OXf, OYfThe yaw rate of axis； N-th of gyrorotor is represented around stator coordinate OXfYfZfIn
OXf, OYfThe deflection angular acceleration of axis；iαn, iβnN-th of gyroscopic couple device coil drive rotor is represented around OXf, OYfAxis deflection
Electric current；kxAnd kyGyroscopic couple device is respectively indicated in OXf, OYfTorque coefficient on axis.
(2) gyro sensitivity torque variable is definedWith
According to (5) formula, gyro sensitivity torque variable can defineWithIt is as follows:
Wherein,It can be carried out once or twice not by the gyrorotor displacement obtained to measurement
Complete differential obtains.
In order to make full use of the dynamic information of system, new variable is defined according to (5) formulaWith
DefinitionWithThe installation transformed matrix for respectively representing spacecraft attitude angular speed and n-th of gyro, then have,
For G1, transformed matrix is installed and is represented byTherefore it is obtained in conjunction with (7) formula and (8) formula,
Similarly, for G1, G2And G3, corresponding installation transformed matrix may be expressed as:
Therefore,
(3) spacecraft three-axis attitude angular speed is solved
It can be obtained according to the above analysis, the survey of the equal feasible system attitude angular rate of the opposite gyro of any two angular momentum
Amount.With G1And G3For,
Such attitude angular rateWithIt can give essence the case where not ignoring inertia coupling terms and cross-couplings item
Really solve,
Similarly, in conjunction with G2And G4It can be accurately obtained
(4) spacecraft three-axis attitude angular speed is solved
(14) and (15) are substituted into (9) formula and (10) formula respectively can obtain spacecraft three-axis attitude angular speed Analytical expression are as follows:
It should be noted that, in conjunction with (9) formula and (10) formula, being neglected for traditional three orthogonal configurations as shown in Figure 3
Under conditions of omitting inertia coupling terms and cross-couplings item, three-axis attitude angular speed can be obtained are as follows:
In order to prove the correctness of the method for the present invention and have superiority, to the method for the present invention and conventional method in identical item
Contrast simulation has been carried out under part.The disturbing moment of spacecraft designs are as follows: Tdx=0.1sin (0.2 π t) Nm or Tdx=
0.1sin (2 π t) Nm, namely Spacecraft During Attitude Maneuver caused by disturbing moment dynamic frequency be respectively 0.1Hz and
1Hz；Gyro error is as follows: the electric current of gyroscopic couple device and the measurement accuracy of rotor displacement are 10-4Magnitude, gyrorotor
Tachometric survey error is not higher than 5 × 10-5Magnitude；Gyro configuration installation error is 20 ".
When not considering gyro error and installation error, figure 4 and figure 5 respectively show disturb bandwidth in 0.1Hz and 1Hz
Under roll attitude angular rate comparing result, wherein abscissa indicates the time, and unit is s, and ordinate indicates angular speed error,
Unit is °/s, angular speed error be measured value is subtracted as true value obtained by.It can be obtained from Fig. 4, in the case where 0.1Hz disturbs bandwidth,
Conventional method and the relative error of the resulting spacecraft attitude angular speed of the method for the present invention are respectively 10-5With 10-15Magnitude.From 5
It is found that the relative error of the method for the present invention remains at 10 when forcing frequency increases to 1Hz from 0.1Hz-15Magnitude；So
And the relative error of conventional method is but from 10-5It is reduced to 10-3Magnitude.
Under the conditions of considering gyro error installation error, it is imitative that the comparison at 0.1Hz and 1Hz is set forth in Fig. 6 and 7
True unit is s as a result, wherein abscissa indicates the time, and ordinate indicates angular speed error, and unit is °/s.It can according to Fig. 6
Know, compared to the situation of gyro free error and installation, when having gyro error and installation, conventional method and the method for the present invention exist
Error under 0.1Hz disturbance increases respectively to 5 × 10-4°/s and 1 × 10-4°/s.When disturbance bandwidth increases to 1Hz, tradition side
The measurement error of method increases to 3 × 10-4°/s, this and result when not having gyro error and installation error as shown in Fig. 5 (b)
Quite.The above results show that compared to gyro error and installation error, the systematic error of conventional method is to influence measurement accuracy
Principal element.As shown in Fig. 7 (b), compared to the situation of gyro free error and installation, the sheet when having gyro error and installation
The angular speed error of inventive method increases to 1 × 10-5°/s.Although this more than be not present gyro error and installation error when
Greatly, but this more than the error of conventional method under the same terms wants small.Therefore, the method for the present invention alleviates attitude angle speed well
Contradiction between rate measurement accuracy and dynamic range.
The content being not described in detail in present specification belongs to the prior art well known to professional and technical personnel in the field.
Claims (4)
1. a kind of spacecraft high-precision angular movement measurement method based on rotator type gyro biorthogonal configuration, it is characterised in that:
Under the premise of the inertia coupling terms and cross-couplings item of not ignoring gyrorotor, pass through the information of 4 gyros in biorthogonal configuration
Fusion operation, can with the attitude angular rate of accurately measure spacecraft, specifically includes the following steps:
(1) kinetic model of in-orbit gyrorotor is established
The radial torque of in-orbit n-th of gyrorotorIt is represented by
Wherein, n=1,2,3,4；Respectively represent the angular speed in gyro reference frame relative inertness space；Respectively represent the angular acceleration in gyro reference frame relative inertness space；Ir, IzRespectively represent gyrorotor
Radial and axial rotary inertia；Ω represents gyrorotor angular velocity of rotation；N-th of gyrorotor is represented around stator
Coordinate system OXfYfZfMiddle OXf, OYfThe yaw rate of axis；N-th of gyrorotor is represented around stator coordinate OXfYfZf
Middle OXf, OYfThe deflection angular acceleration of axis；iαn, iβnN-th of gyroscopic couple device coil drive rotor is represented around OXf, OYfAxis deflection
Electric current；kxAnd kyGyroscopic couple device is respectively indicated in OXf, OYfTorque coefficient on axis；
(2) gyro sensitivity torque variable is definedWith
Wherein,It can be carried out by the gyrorotor displacement obtained to measurement incomplete once or twice
Differential obtains；
(3) spacecraft three-axis attitude angular speed is solved
Wherein,
2. a kind of spacecraft high-precision angular movement measurement based on rotator type gyro biorthogonal configuration according to claim 1
Method, it is characterised in that: by the information fusion operation between multiple gyros in biorthogonal configuration, can further obtain spacecraft appearance
State angular accelerationAnalytical expression,
Obtained attitude angle acceleration analysis precision is not reduced with the increase of spacecraft dynamic frequency.
3. a kind of spacecraft high-precision angular movement measurement based on rotator type gyro biorthogonal configuration according to claim 1
Method, it is characterised in that the rotator type gyro includes flexible gyroscope, magnetic floating gyro, liquid floated gyroscope, electrostatic gyroscope, three floating tops
Spiral shell, two floating gyros etc. can obtain the rotator type gyroscope of gyrorotor displacement and torquer current information.
4. a kind of spacecraft high-precision angular movement measurement based on rotator type gyro biorthogonal configuration according to claim 1
Method is divided into two groups it is further characterized in that the biorthogonal configuration is made of 4 rotator type gyros, two tops in each group
Straight line where the rotary shaft of spiral shell is nominally directed toward spatially is mutually perpendicular to, can coplanar also antarafacial；Two teetotums between two groups
Straight line where shaft is nominally directed toward is parallel to each other or collinearly, straight line where the nominal direction of the another two gyro between two groups is also mutual
It is parallel or conllinear.
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| CN110068336A (en) * | 2019-04-25 | 2019-07-30 | 中国人民解放军战略支援部队航天工程大学 | A kind of angular movement measurement method based on magnetic suspension control sensitivity gyro parallel configuration |
| CN110823448A (en) * | 2019-11-21 | 2020-02-21 | 中国船舶重工集团公司第七0七研究所 | Method for measuring axial mass center of rotor assembly of gyroscope with two degrees of freedom |
| CN111412930A (en) * | 2020-04-30 | 2020-07-14 | 中国船舶重工集团公司第七0七研究所 | Calibration operation method for installation error of combined attitude measurement device |
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Cited By (3)
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| CN110068336A (en) * | 2019-04-25 | 2019-07-30 | 中国人民解放军战略支援部队航天工程大学 | A kind of angular movement measurement method based on magnetic suspension control sensitivity gyro parallel configuration |
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| CN111412930A (en) * | 2020-04-30 | 2020-07-14 | 中国船舶重工集团公司第七0七研究所 | Calibration operation method for installation error of combined attitude measurement device |
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Application publication date: 20181218 |