CN110162855B - Dynamic precision analysis and error distribution method for satellite rotation load of remote sensing satellite - Google Patents
Dynamic precision analysis and error distribution method for satellite rotation load of remote sensing satellite Download PDFInfo
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Abstract
A dynamic precision analysis and error distribution method for a satellite rotation load of a remote sensing satellite belongs to the field of spacecraft attitude precision modeling. The method solves the problems that the accuracy of dynamic precision modeling analysis and evaluation of the rotary load on the satellite of the remote sensing satellite connected with the magnetic suspension bearing is low and error distribution cannot be carried out in the conventional method. Analyzing an interference source according to a working mode and a structure composition of a remote sensing satellite system, determining an on-satellite rotation load precision error link and an error composition, and establishing an on-satellite rotation load precision error transfer chain; and then according to a corresponding theoretical formula, establishing a dynamic precision model of the satellite rotation load suitable for multi-error link transmission so as to accurately analyze and evaluate the dynamic precision of the satellite rotation load of the remote sensing satellite connected with the magnetic suspension bearing, and realizing the task requirement of designing and distributing the upper limit of the error of each error link according to the requirement of the rotation load on the ground latitude precision.
Description
Technical Field
The invention belongs to the field of spacecraft attitude precision modeling, and particularly relates to a dynamic precision analysis and error distribution method for a rotary load on a remote sensing satellite.
Background
With the continuous increase of the space task demand, the demands of the pointing precision and the attitude stability of the satellite rotating load are higher and higher. Combining the working mode and the structural characteristics of the remote sensing satellite, under the influence of multi-source interference, the pointing precision and the attitude stability of the on-satellite rotating load are subjected to the cumulative action of multiple error links, such as: track control link errors, attitude control link errors, load spinning link errors and the like.
The establishment of the on-satellite rotation load dynamic precision model suitable for multi-error link transmission to effectively and accurately evaluate the current attitude precision of the rotation load is the early foundation for developing the research work of an advanced control method, can provide error distribution and verification of each link for the design of a control system according to the attitude precision of the rotation load and the conversion result of ground latitude, can provide measurement indexes for evaluating and comparing the performance of the control system, and has high academic value and engineering significance. In addition, the magnetic suspension bearing is a non-contact bearing for realizing rotor floating by utilizing electromagnetic force, has the characteristics of high rotating speed, low power consumption, no abrasion, no need of lubrication and medium-high frequency vibration suppression, and has wide application prospect when being applied to a remote sensing satellite as an on-satellite rotating load connecting component compared with the traditional mechanical bearing.
However, due to the low rigidity characteristic of the magnetic suspension bearing, the connection precision of the on-satellite rotating load is poor, the on-satellite rotating load generates nutation precession relative to a remote sensing satellite platform under the action of multi-source interference, an error link is added, and the accuracy of the on-satellite rotating load dynamic precision modeling analysis and evaluation of the remote sensing satellite connected with the magnetic suspension bearing by using the existing method is low and error distribution cannot be performed, so that higher requirements are provided for the research of the on-satellite rotating load dynamic precision modeling method of the remote sensing satellite connected with the magnetic suspension bearing.
Disclosure of Invention
The invention aims to solve the problems that the accuracy of modeling, analyzing and evaluating the dynamic precision of the rotating load on the remote sensing satellite connected with a magnetic suspension bearing by the conventional method is low and error distribution cannot be carried out.
The technical scheme adopted by the invention for solving the technical problems is as follows: the dynamic precision analysis and error distribution method for the satellite rotation load of the remote sensing satellite comprises the following steps:
step one, determining the structure and the working mode of a remote sensing satellite system, and defining an earth center equatorial coordinate system oxyz and an orbit coordinate system ooxoyozoCenter of mass coordinate system o of rotating load satellite system sxsyszsBody coordinate system o of satellite platformbxbybzbAnd a rotational load body coordinate system opxpypzp；
Determining each error link contained in the remote sensing satellite system, wherein each error link forms a precision error transmission chain, and each error link acts together and is transmitted to the satellite rotation load in an accumulated mode;
step three, respectively calculating the rotating load direction on the actual satellite and the rotating load direction on the expected satellite in the expected orbit coordinate system oo1xo1yo1zo1Projection z ofcAnd zdIn z isdAnd zcThe included angle delta phi is used as an index to estimate the pointing accuracy of the rotating load on the satellite;
step four, calculating the angular velocity vector of the on-satellite rotating load relative to the earth center equatorial coordinate system oxyz in the measured on-satellite rotating load body coordinate system op1xp1yp1zp1Expected value ω of downward projectiondAnd the true value ωcUsing the desired value omegadWith the true value omegacEstimating the three-axis attitude stabilization precision of the on-satellite rotation load by using the difference;
and converting the estimation results of the pointing accuracy and the three-axis attitude stability accuracy of the satellite rotating load into a rotating load ground longitude and latitude index, and distributing and verifying the upper error limit of each error link of the remote sensing satellite system by utilizing the longitude and latitude index.
The invention has the beneficial effects that: according to the method for dynamic precision analysis and error distribution of the satellite rotating load of the remote sensing satellite, interference sources are analyzed according to the working mode and the structure composition of a remote sensing satellite system, the satellite rotating load precision error link and the error composition are determined, and a satellite rotating load precision error transfer chain is established; and then according to a corresponding theoretical formula of the attitude kinematics, constructing a dynamic precision model of the satellite rotation load suitable for multi-error link transmission, reflecting the influence of errors of each link on the precision of the satellite rotation load, so as to improve the accuracy of dynamic precision evaluation and analysis of the satellite rotation load of the remote sensing satellite connected with the magnetic suspension bearing, thereby realizing the task requirements of designing and distributing ground warp latitude precision indexes aiming at a certain determined load and verifying the upper limit of the errors of each error link.
Drawings
FIG. 1 is a flow chart of a dynamic precision analysis and error distribution method for a satellite rotating load of a remote sensing satellite according to the invention;
FIG. 2 is a schematic diagram of the structure and mode of operation of the remote sensing satellite system of the present invention;
FIG. 3 is a schematic diagram of the present invention for coordinate system definition;
FIG. 4 is a schematic diagram of the error components and error components of the present invention;
FIG. 5 is a schematic diagram of a satellite rotational load pointing accuracy error transfer chain;
in fig. 5: o isO1Representing a desired orbital coordinate system, OO2Representing the measured orbital coordinate system, Os1Measured remote sensing satellite system centroid coordinate system, Op1Body coordinate system, Z, representing the measured rotational load on the satellitedRepresenting the desired orientation of the optical axis, Z representing the actual orientation of the optical axis, theta0Representative of track control errors, θbRepresenting nutation error, theta, of the remote sensing satellite platformwzRepresenting bearing shaft nutation error, thetazjRepresenting precession error of the bearing shaft, thetapRepresenting the nutation error of the satellite rotational load, thetazRepresenting spin angle error of the satellite rotational load, thetazxRepresenting a pointing error;
FIG. 6 is a schematic diagram of the pointing accuracy of the rotating load on the satellite and the conversion of the latitude to the ground warp;
in FIG. 6, L represents the subsatellite point and o represents the geocentric point.
Detailed Description
The first embodiment is as follows: this embodiment will be described with reference to fig. 1. The method for dynamic precision analysis and error distribution of the rotary load on the satellite of the remote sensing satellite comprises the following steps:
Step one, determining the structure and the working mode of a remote sensing satellite system, and defining an earth center equator coordinate system oxyz and an orbit coordinate system ooxoyozoAnd a rotating load satellite system centroid coordinate system osxsyszsBody coordinate system o of satellite platformbxbybzbAnd a rotational load body coordinate system opxpypzp；
Determining each error link contained in the remote sensing satellite system, wherein each error link forms a precision error transmission chain, and each error link acts together and is transmitted to a satellite rotating load in an accumulated manner;
step three, respectively calculating the rotating load direction on the actual satellite and the rotating load direction on the expected satellite in the expected orbit coordinate system oo1xo1yo1zo1Projection z ofcAnd zdIn z isdAnd zcThe included angle delta phi is used as an index to estimate the pointing accuracy of the rotating load on the satellite;
step four, calculating the angular velocity vector of the on-satellite rotating load relative to the earth center equatorial coordinate system oxyz in the measured on-satellite rotating load body coordinate system op1xp1yp1zp1Expected value ω of downward projectiondAnd the true value ωcUsing the desired value omegadWith the true value omegacEstimating the three-axis attitude stabilization precision of the on-satellite rotation load by using the difference;
and converting the estimation results of the pointing accuracy and the three-axis attitude stability accuracy of the satellite rotating load into a rotating load ground longitude and latitude index, and distributing and verifying the upper error limit of each error link of the remote sensing satellite system by utilizing the longitude and latitude index.
Aiming at the problem of modeling of dynamic precision of the satellite rotating load of the remote sensing satellite affected by multi-source interference under the connection of the magnetic suspension bearing, the embodiment establishes a satellite rotating load pointing precision and attitude stability model according to attitude kinematics from the working mode and structural characteristics of the remote sensing satellite, and can embody the influence of each error link in the model, thereby providing index distribution and index verification of each link for the design of a control system and providing evaluation indexes for comparing the performance of a control method.
The second embodiment is as follows: the first difference between the present embodiment and the specific embodiment is: the specific process of the step one is as follows:
as shown in fig. 2, the structure of the remote sensing satellite system comprises a remote sensing satellite platform subsystem, an on-satellite load subsystem and a rotary joint; the remote sensing satellite platform subsystem further comprises: the remote sensing satellite platform comprises a platform three-axis orthogonal flywheel and platform bilateral solar sailboards; the on-board load subsystem further comprises: an on-board rotating load and a load internal flywheel; the rotary joint is a magnetic suspension bearing (and is provided with a mechanical bearing and a driving motor);
the working mode of the remote sensing satellite system is as follows: the remote sensing satellite platform keeps the ground orientation through a platform three-axis orthogonal wheel control system, and the solar sailboard single-axis drive at the two sides of the platform keeps the sun orientation so as to provide energy required by the satellite; the rotary joint is provided with a magnetic suspension bearing for connection, and a mechanical bearing is reserved as an emergency protection scheme; the on-satellite rotating load is controlled by a load inner flywheel to keep rotating at a constant speed, and when the angular momentum of the load inner flywheel is saturated, a magnetic torquer and a driving motor are required to provide an unloading scheme;
The coordinate system is defined as follows: as shown in fig. 3;
establishing an earth center equatorial coordinate system (inertial coordinate system) oxyz by taking the earth center as a coordinate origin o, wherein the x axis of the earth center equatorial coordinate system points to a vernalization point at the moment J2000 from the earth center in a J2000 earth plane equatorial plane, the z axis is a normal line of the J2000 earth plane equatorial plane and points to the north pole direction, and the y axis, the x axis and the z axis form a right-handed system;
using the mass center of the remote sensing satellite system as the origin of coordinates ooEstablishing a track coordinate system ooxoyozoX of said orbital coordinate systemoThe axis being in the orbital plane and pointing in the direction of advance, z, of the remote sensing satellite systemoThe axis is directed from the centroid of the remote sensing satellite system to the centroid, yoAxis is same as xoAxis, zoThe shaft constitutes a right-hand system;
noting the expected orbital coordinate system oo1xo1yo1zo1The origin of coordinates of the desired orbital coordinate system is co-orbital coordinate system ooxoyozoThe desired orbital coordinate system o in the absence of orbital control errorso1xo1yo1zo1Three-axis co-orbital coordinate system ooxoyozoThe three axes of (A) are superposed; recording the measured orbital coordinate system as oo2xo2yo2zo2Measured orbital coordinate system oo2xo2yo2zo2Coordinate origin of the same rail seatMark system ooxoyozoOf the measured orbit coordinate system o in the absence of orbit determination errorso2xo2yo2zo2Three-axis co-orbital coordinate system ooxoyozoThe three axes of (A) are superposed;
using the mass center of the remote sensing satellite system as the origin of coordinates o sEstablishing a centroid coordinate system o of the remote sensing satellite systemsxsyszs(connected to the body), x of the remote sensing satellite system centroid coordinate systemsThe axis points to the axial direction, y of the remote sensing satellite platformsAxis and zsThe axis is located in the axial section of the remote sensing satellite platform, and ysAxis, zsAxis is same as xsThe shaft constitutes a right-hand system; under the condition of no attitude control error, the three axes of the centroid coordinate system of the remote sensing satellite system are simultaneously measured to form an orbit coordinate system oo2xo2yo2zo2The three axes of (A) are superposed;
measured remote sensing satellite system centroid coordinate system os1xs1ys1zs1And measured remote sensing satellite system centroid coordinate system os1xs1ys1zs1Coordinate origin of the system is the same as the centroid coordinate system o of the remote sensing satellite systemsxsyszsThe measured centroid coordinate system o of the remote sensing satellite system when no attitude measurement error existss1xs1ys1zs1Three-axis same-remote sensing satellite system centroid coordinate system osxsyszsThe three axes of (A) are superposed;
using the mass center of the remote sensing satellite platform as the origin of coordinates obEstablishing a body coordinate system o of the remote sensing satellite platformbxbybzb(connected with the body) and the three-axis direction of the body coordinate system of the remote sensing satellite platform is the same as the centroid coordinate system o of the remote sensing satellite systemsxsyszsPoint to the three axes of (1);
coordinate system of measured remote sensing satellite platform body is ob1xb1yb1zb1And measured remote sensing satellite platform body seatMark system ob1xb1yb1zb1Coordinate origin of the satellite system is identical to the body coordinate system o of the remote sensing satellite platform bxbybzbMeasured remote sensing satellite platform body coordinate system ob1xb1yb1zb1Three-axis pointing and measuring remote sensing satellite system centroid coordinate system os1xs1ys1zs1Point to the three axes of (a);
using the center of mass of the satellite rotating load as the origin o of coordinatespEstablishing a satellite rotation load body coordinate system opxpypzp(connected with the body) when the connection error of the rotary joint is zero and the self-rotation angle of the rotary load on the satellite is zero, the coordinate system o of the rotary load on the satellite bodypxpypzpThree-axis pointing and measuring remote sensing satellite system centroid coordinate system os1xs1ys1zs1Point to the three axes of (1);
recording the coordinate system of the measured on-satellite rotating load body as op1xp1yp1zp1And measured satellite rotation load body coordinate system op1xp1yp1zp1Coordinate origin of the satellite system is identical to the body coordinate system o of the remote sensing satellite platformpxpypzpOf the measured satellite rotational load body coordinate system o without measurement errorsp1xp1yp1zp1Three axes of the same coordinate system opxpypzpThe three axes of (a) coincide.
The interference sources for the rotational loads on the satellite were analyzed as follows:
according to the working mode and the structural characteristics of the remote sensing satellite, the interference types are divided into three types, namely orbit interference, attitude interference and spin interference.
The track disturbance includes: the space interference force borne by the on-track running influences the track control precision; the orbit perturbation interference affects the orbit control precision, and specifically comprises earth gravitational field perturbation interference, sun gravitational field perturbation interference, sunlight pressure perturbation interference and moon gravitational field perturbation interference.
The attitude disturbance includes: load dynamic and static unbalance interference, mass center deviation and spindle deflection inevitably occur in the processes of production and installation of rotating load, namely residual dynamic and static unbalance amount exists, the unbalance amount is not offset by the existing mass balancing technology, unbalance amount generates unbalance interference torque when the load spins, and the unbalance torque is transmitted to a remote sensing satellite platform through a magnetic suspension bearing to influence attitude control precision;
model parameter time-varying interference, namely, the moment of inertia of the whole satellite changes along with the rotation of the load due to the difference of the radial main inertia of the load, so that the time-varying property of the model parameters is reflected, an attitude interference torque is generated, and the attitude interference torque is transmitted to a remote sensing satellite platform through a magnetic suspension bearing to influence the attitude control precision;
the load axial offset angular momentum interferes, when the load flywheel does not completely compensate the axial angular momentum of the load, the residual axial angular momentum of the load generates an attitude interference torque which is transmitted to the remote sensing satellite platform through the magnetic suspension bearing to influence the attitude control precision;
the satellite platform vibration interference, the remote sensing satellite attitude motion and the solar panel driving rotation excite the solar panel to vibrate elastically in a wide frequency band for multiple times, and when the deviation of the center of mass of the flywheel is considered, the three-axis orthogonal flywheel in the remote sensing satellite platform can also generate medium-high frequency vibration interference, so that the attitude control precision is influenced.
The space environment interference moment acts on the remote sensing satellite system to influence the attitude control precision, and the interference influence of gravity gradient moment, geomagnetic moment and solar radiation moment is specifically considered.
In order to isolate a high-frequency vibration transmission path in a satellite platform, a magnetic suspension bearing needs to be controlled at a lower rigidity level, so that the phenomenon of nutation precession of a load relative to the platform is also generated under the action of load dynamic and static imbalance interference and model parameter time-varying interference, and the load connection precision is influenced.
Spin interference: the friction resistance of the magnetic bearing interferes to prevent the bearing rotating shaft from spinning, so that the spinning precision of the load is influenced; the magnetic bearing magnetic resistance interferes the torque, and the magnetic resistance interference of the bearing rotating shaft moving in the magnetic suspension bearing magnetic field influences the load spinning precision; the cogging torque of the motor is interfered, the driving motor is not electrified, and the cogging torque interference of the driving motor influences the load spinning precision due to the influence of the permanent magnet poles; the motor electromagnetic resistance moment is driven, and the load spinning precision is influenced by the magnetic field resistance interference on the movement of the bearing rotating shaft in the motor magnetic field.
The third concrete implementation mode: this embodiment will be described with reference to fig. 4. The second embodiment is different from the first embodiment in that: the specific process of the second step is as follows:
Determining error links contained in the remote sensing satellite system as follows: a track error link, a platform attitude error link, a magnetic suspension bearing connection error link and an on-satellite rotating load spinning error link;
the track error link consists of two parts: part of the control error of the track, influenced by track disturbances, resulting in a desired track coordinate system oo1xo1yo1zo1And a track coordinate system ooxoyozoThere is a relative deviation between them; the other part is orbit determination error, which is influenced by orbit model error, satellite-time error, initial value error and calculation error to cause orbit coordinate system ooxoyozoAnd a measured orbital coordinate system oo2xo2yo2zo2There is a relative deviation between them;
the platform attitude error link consists of two parts: one part is attitude control error, which is influenced by attitude interference to cause a system o in the remote sensing satellite platformbxbybzbAnd a measured orbital coordinate system oo2xo2yo2zo2There is a relative deviation between them; the other part is attitude measurement error including star sensor error and fiber optic gyroscope error, which causes a system o of the remote sensing satellite platformbxbybzbAnd measured remote sensing satellite platform body system ob1xb1yb1zb1Relative deviation exists between the two measuring elements, wherein the error of the measuring element is formed by internal calibration error, space environment interference, aging of the measuring element, noise interference and mounting error of the measuring element;
The magnetic suspension bearing has two connection error linksThe part constitutes: one part is the axle gap run-out error, which causes the satellite rotation load body coordinate system opxpypzpRelative measurement remote sensing satellite platform body system ob1xb1yb1zb1Radial nutation precession attitude deviation exists; the other part is measurement error, which is influenced by a measuring element to cause a satellite rotation load body coordinate system opxpypzpAnd measured on-satellite rotational load body coordinate system op1xp1yp1zp1Radial attitude deviation exists between the two;
the spin error link of the satellite rotating load consists of two parts, one part is a rotating speed control error, which causes a coordinate system o of the satellite rotating load bodypxpypzpRelative measurement remote sensing satellite platform body system ob1xb1yb1zb1Axial rotation speed errors exist; the other part is a measurement error which is influenced by the measurement errors of the micrometering gyroscope and the star sensor to cause a body coordinate system o of the satellite rotating loadpxpypzpAnd measured on-satellite rotational load body coordinate system op1xp1yp1zp1Axial rotation speed deviation exists between the two parts;
the track error link, the platform attitude error link, the magnetic suspension bearing connection error link and the satellite rotation load spinning error link are connected in series to form a precision error transmission chain, and all the error links act together and are transmitted to the satellite rotation load in an accumulated mode. The attitude orientation error of the satellite rotation load and the three-axis attitude stability error are reflected.
The fourth concrete implementation mode is as follows: this embodiment will be described with reference to fig. 5. The third difference between the present embodiment and the specific embodiment is that: the specific process of the third step is as follows:
track error link results in a track coordinate system ooxoyozoRelative desired orbital coordinate system oo1xo1yo1zo1With errors, the orbital coordinate system o is described by zxy sequenceoxoyozoRelative desired orbital coordinate system oo1xo1yo1zo1Attitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: desired orbital coordinate system oo1xo1yo1zo1Around zo1Angle of rotation of the shaft
To intermediate system i1Intermediate system i1Around itself (i.e. intermediate system i)1) X-axis rotational angle of delta theta1To intermediate system i2Intermediate system i2Y around an orbital coordinate systemoAxial rotation through delta psi1To the orbital coordinate system ooxoyozo；
Δθ1And delta phi1All represent angle errors with no actual physical meaning;
errors between the measured orbit and the true orbit used for attitude control of the satellite, i.e. the measured orbital coordinate system oo2xo2yo2zo2Relative orbital coordinate system ooxoyozoWith errors, describing the measured orbital coordinate system o by zxy ordero2xo2yo2zo2Relative orbital coordinate system ooxoyozoAttitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: orbital coordinate system ooxoyozoAround zoAngle of rotation of the shaft
To intermediate system i1', intermediate system i1' around itself (i.e. intermediate system i) 1') x-axis over angle Δ θ2To intermediate system i2', intermediate system i2' around the measured orbital coordinate system yo2Angle delta psi of shaft rotation2To the measured orbital coordinate system oo2xo2yo2zo2；
Δθ2And delta phi2All represent angle errors; there is no actual physical meaning;
system o of remote sensing satellite platform body for measuring by combining remote sensing satellite platform control with attitude measurement errorb1xb1yb1zb1Relative measured orbital coordinate system oo2xo2yo2zo2Error exists, and the system o of the remote sensing satellite platform is described by using zxy sequenceb1xb1yb1zb1Relative measured orbital coordinate system oo2xo2yo2zo2Attitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: measured orbital coordinate system oo2xo2yo2zo2Around zo2Angle of rotation of the shaft
To intermediate system i1", intermediate system i1"around itself (i.e. intermediate system i)1") x-axis over angle Δ θ3To intermediate system i2", intermediate system i2"y around the system of remote sensing satellite platformb1Angle delta psi of shaft rotation3To the remote sensing satellite platform body system ob1xb1yb1zb1；
Δθ3And delta phi3Respectively representing yaw angle, roll angle and pitch angle;
the magnetic suspension bearing causes attitude deviation between a bearing rotating shaft and a bearing stator, the bearing rotating shaft is fixedly connected with an onboard rotating load, the bearing stator is fixedly connected with a remote sensing satellite platform, and then a measured onboard rotating load body coordinate system o is combined with onboard rotating load spinning precision and attitude measurement error p1xp1yp1zp1Relative measurement remote sensing satellite platform body system ob1xb1yb1zb1With errors, the global coordinate system o of the measured satellite rotation load is described by zyxp1xp1yp1zp1Relative measurement remote sensing satellite platform body system ob1xb1yb1zb1Attitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: measured remote sensing satellite platform body system ob1xb1yb1zb1Around zb1Angle delta psi of shaft rotation4To intermediate system i1", intermediate system i1"wind aroundSelf y-axis rotation angle delta theta4To intermediate system i2", intermediate system i2' x around measured satellite rotation load body coordinate systemp1Angle of rotation of the shaft
To the measured on-satellite rotational load body coordinate system op1xp1yp1zp1；Δψ4And Δ θ4All represent an angle error which is represented by,
representing the expected rotation angle of the rotating load on the satellite relative to the remote sensing satellite platform,
representing the rotation angle deviation of the on-satellite rotation load relative to the remote sensing satellite platform;
respectively calculating the rotating load direction on the actual satellite and the rotating load direction on the expected satellite in the expected orbit coordinate system oo1xo1yo1zo1Projection z ofcAnd zd；
Specifying actual on-board rotational load pointing in measured on-board rotational load body coordinate system op1xp1yp1zp1Projection of lower is zp1′＝[0,0,1]TThen the rotational load on the satellite is expected to be directed at the desired orbital coordinate system oo1xo1yo1zo1Projection z ofdComprises the following steps:
the actual on-board rotating load is directed at the desired orbital coordinate system o o1xo1yo1zo1Projection of lower zcComprises the following steps:
wherein, the first and the second end of the pipe are connected with each other,
representing the orbital coordinate system ooxoyozoRelative to a desired orbital coordinate system oo1xo1yo1zo1The attitude transformation matrix of (a) is,
orbital coordinate system o representing measurementso2xo2yo2zo2Relative to the orbital coordinate system ooxoyozoThe attitude transformation matrix of (a) is,
system o of remote sensing satellite platform body for representing measurementb1xb1yb1zb1Relative to the measured orbital coordinate system oo2xo2yo2zo2The attitude transformation matrix of (a) is,
body coordinate system o of satellite rotation load representing measurementp1xp1yp1zp1System o of remote sensing satellite platform relative to measurementb1xb1yb1zb1The attitude transformation matrix of (2);
combining the formulas (2) to (6) to obtain the projection z of the actual on-satellite rotation load pointing direction under the expected orbit coordinate systemcSolving two vectors zdAnd zcThe included angle Δ Φ between is:
ΔΦ＝arccos(zd Tzc) (7)
obtaining an established on-satellite rotation load pointing accuracy model; wherein: arccos (·) represents an inverse cosine computation;
and estimating the pointing accuracy of the rotating load on the satellite by using the included angle delta phi.
The fifth concrete implementation mode: the fourth difference between this embodiment and the specific embodiment is that: the specific process of the step four is as follows:
combining the on-satellite rotation load pointing accuracy model established in the third step, utilizing the angular velocity vector of the on-satellite rotation load relative to the earth center equatorial coordinate system oxyz to measure the on-satellite rotation load body coordinate system o p1xp1yp1zp1Expected value ω of downward projectiondAnd true value ωcThe difference is used for measuring the three-axis attitude stabilization precision of the satellite rotating load;
measured remote sensing satellite platform system o without considering attitude errorb1xb1yb1zb1And a desired orbital coordinate system oo1xo1yo1zo1Coincident and measured on-satellite rotation load body coordinate system op1xp1yp1zp1And the satellite rotation load body coordinate system opxpypzpCoincident and measured on-satellite rotation load body coordinate system op1xp1yp1zp1Relative measurement remote sensing satellite platform body system ob1xb1yb1zb1With only axial rotation, in which case the system o of the remote sensing satellite platform is measuredb1xb1yb1zb1Angular velocity relative to the equatorial coordinate system oxyz and the desired orbital coordinate system oo1xo1yo1zo1The angular speeds of the relative geocentric equatorial coordinate system oxyz are consistent, and the measured remote sensing satellite platform body system ob1xb1yb1zb1Angular velocity relative to the equatorial coordinate system oxyz in the desired orbital coordinate system oo1xo1yo1zo1Projection of the lower is [ 0-omega ]0 0]TWherein: omega0Representing orbital angular velocity, the angular velocity vector of the rotating load on the satellite relative to the equatorial coordinate system oxyz is in the desired orbital coordinate system oo1xo1yo1zo1Lower projection omegad1Expressed (without attitude error) as:
wherein the content of the first and second substances,
representing a scalar quantity
Taking a time derivative;
when the attitude error is considered, the sitting-in angular velocity vector of the on-satellite rotating load relative to the equatorial coordinate system oxyz is obtained by combining equation (2) Mark system op1xp1yp1zp1Expected value ω of downward projectiondComprises the following steps:
wherein the content of the first and second substances,
is to
The transpose of (a) is performed,
is to
The transpose of (a) is performed,
is to
The transpose of (a) is performed,
is to
Transposing; meanwhile, under the real condition of considering attitude error, the measured remote sensing satellite platform body coordinate system ob1xb1yb1zb1Angular velocity relative to the equatorial coordinate system oxyz in coordinate system ob1xb1yb1zb1Lower projection omegabComprises the following steps: omegab＝[Δωbx -ω0+Δωby Δωbz]TWherein Δ ωbx、Δωby、ΔωbzRespectively indicates that the angular speed error of the coordinate system of the remote sensing satellite platform body is ob1xb1yb1zb1Is ofxb1、yb1、zb1And (4) projecting under the shaft.
In the formula: omegabThe method can be obtained from the measurement result of the fiber-optic gyroscope on the remote sensing satellite platform;
measured satellite rotation load body coordinate system op1xp1yp1zp1Relative measurement remote sensing satellite platform body coordinate system ob1xb1yb1zb1With attitude errors, from the coordinate system ob1xb1yb1zb1Sequentially switched by delta phi through zyx4、Δθ4、
To a coordinate system op1xp1yp1zp1Noting the attitude angular rate of
(wherein:
is to delta phi4The derivative of (a) of (b),
is to Δ θ4The derivative of (a) of (b),
is to
Derivative of (d); the angular velocity vector of the rotating load on the satellite relative to the equatorial coordinate system oxyz of the earth is in the coordinate system op1xp1yp1zp1True value ω of the lower projectioncComprises the following steps:
intermediate variable C in equation (10)x、CyAnd CzRespectively expressed as:
the expected value ω is obtained by combining the formulas (9) and (10)dWith the true value omegacThe difference Δ ω is:
Δω＝ωc-ωd (12)
using desired value omega dWith true value omegacAnd estimating the three-axis attitude stabilization precision of the on-satellite rotation load by the difference delta omega. Namely, the three-axis attitude stabilization precision model of the formula (12) is obtained.
According to the formula (7) and the formula (12), the pointing accuracy and the three-axis attitude stability of the rotating load on the remote sensing satellite can be effectively evaluated.
According to theoretical results, the precision model is subjected to the accumulative action of error links, and the precision model embodies strong dynamics in consideration of various interference action characteristics; in addition, mutual coupling characteristics exist between the three-axis attitude stabilization precision and the pointing precision. As shown in fig. 6, the load pointing accuracy and the triaxial attitude stabilization accuracy of predictive analysis by using the accuracy model can be converted into load ground-warp latitude accuracy, so that the upper error limit allowed by each error link is designed and distributed according to a certain determined load ground-warp latitude accuracy requirement, whether the error design scheme of each link meets the ground-warp latitude requirement of the load is verified, the performance of each control system can be evaluated, and the method has high academic value and engineering significance.
The above-described calculation examples of the present invention are merely to explain the calculation model and the calculation flow of the present invention in detail, and are not intended to limit the embodiments of the present invention. It will be apparent to those skilled in the art that other variations and modifications of the present invention can be made based on the above description, and it is not intended to be exhaustive or to limit the invention to the precise form disclosed, and all such modifications and variations are possible and contemplated as falling within the scope of the invention.
Claims (5)
1. The method for dynamic precision analysis and error distribution of the satellite rotation load of the remote sensing satellite is characterized by comprising the following steps of:
step one, determining the structure and the working mode of a remote sensing satellite system, and defining an earth center equatorial coordinate system oxyz and an orbit coordinate system ooxoyozoCenter of mass coordinate system o of rotating load satellite systemsxsyszsSatellite platform body coordinate system obxbybzbAnd a rotational load body coordinate system opxpypzp；
Determining each error link contained in the remote sensing satellite system, wherein each error link forms a precision error transmission chain, and each error link acts together and is transmitted to the satellite rotation load in an accumulated mode;
step three, respectively calculating the rotating load direction on the actual satellite and the rotating load direction on the expected satellite in the expected orbit coordinate system oo1xo1yo1zo1Projection z ofcAnd zdIn z isdAnd zcThe included angle delta phi is used as an index to estimate the pointing accuracy of the rotating load on the satellite;
step four, calculating the angular velocity vector of the on-satellite rotating load relative to the earth center equatorial coordinate system oxyz in the measured on-satellite rotating load body coordinate system op1xp1yp1zp1Expected value ω of downward projectiondAnd the true value ωcUsing the desired value omegadWith the true value omegacEstimating the three-axis attitude stabilization precision of the on-satellite rotation load by using the difference;
and converting the estimation results of the pointing accuracy and the three-axis attitude stability accuracy of the satellite rotating load into a rotating load ground longitude and latitude index, and distributing and verifying the upper error limit of each error link of the remote sensing satellite system by utilizing the longitude and latitude index.
2. The method for dynamic precision analysis and error distribution of the rotating load on the satellite of the remote sensing satellite according to claim 1, wherein the specific process of the first step is as follows:
the structure of the remote sensing satellite system comprises a remote sensing satellite platform subsystem, an on-satellite load subsystem and a rotary joint; the remote sensing satellite platform subsystem comprises: the remote sensing satellite platform comprises a platform three-axis orthogonal flywheel and platform bilateral solar sailboards; the on-board load subsystem further comprises: an on-board rotating load and a load internal flywheel; the rotary joint is a magnetic suspension bearing;
the working mode of the remote sensing satellite system is as follows: the remote sensing satellite platform keeps the earth orientation through a platform three-axis orthogonal wheel control system, solar sailboard single-axis driving on two sides of the platform keeps the sun orientation, a rotary joint is provided with a magnetic suspension bearing for connection, a rotary load on the satellite keeps constant-speed rotation through control of a load inner flywheel, and a magnetic torquer and a driving motor are required to provide an unloading scheme when angular momentum of the load inner flywheel is saturated;
the coordinate system is defined as follows:
establishing an earth center equatorial coordinate system oxyz by taking the earth center as a coordinate origin o, wherein the x axis of the earth center equatorial coordinate system points to a vernalization point at the moment J2000 from the earth center in a J2000 earth plane equatorial plane, the z axis is a normal line of the J2000 earth plane equatorial plane and points to the north pole direction, and the y axis, the x axis and the z axis form a right-handed system;
Using the centroid of the remote sensing satellite system as the origin of coordinates ooEstablishing a track coordinate system ooxoyozoX of said orbital coordinate systemoThe axis being in the orbital plane and pointing in the direction of advance, z, of the remote sensing satellite systemoThe axis is directed from the centroid of the remote sensing satellite system to the centroid, yoAxis is same as xoAxis, zoThe shaft constitutes a right-hand system;
noting the expected orbital coordinate system oo1xo1yo1zo1Coordinate origin of desired orbital coordinate system is co-orbitalMark system ooxoyozoThe desired orbital coordinate system o in the absence of orbital control errorso1xo1yo1zo1Three-axis co-orbital coordinate system ooxoyozoThe three axes of (A) are superposed; recording the measured orbital coordinate system as oo2xo2yo2zo2Measured orbital coordinate system oo2xo2yo2zo2Co-orbital coordinate system ooxoyozoOf the measured orbit coordinate system o in the absence of orbit determination errorso2xo2yo2zo2Three-axis co-orbital coordinate system ooxoyozoThe three axes of (A) are superposed;
using the mass center of the remote sensing satellite system as the origin of coordinates osEstablishing a centroid coordinate system o of the remote sensing satellite systemsxsyszsX of the centroid coordinate system of the remote sensing satellite systemsThe axis points to the axial direction, y of the remote sensing satellite platformsAxis and zsThe axis is located in the axial section of the remote sensing satellite platform, and ysAxis, zsAxis is same as xsThe shaft constitutes a right-hand system; under the condition of no attitude control error, the three axes of the centroid coordinate system of the remote sensing satellite system are simultaneously measured to form an orbit coordinate system o o2xo2yo2zo2The three axes of (A) are superposed;
measured remote sensing satellite system centroid coordinate system os1xs1ys1zs1And measured remote sensing satellite system centroid coordinate system os1xs1ys1zs1Coordinate origin of the system is the same as the centroid coordinate system o of the remote sensing satellite systemsxsyszsThe measured centroid coordinate system o of the remote sensing satellite system when no attitude measurement error existss1xs1ys1zs1Three-axis same-remote sensing satellite system centroid coordinate system osxsyszsThe three axes of (A) are superposed;
uses the mass center of the remote sensing satellite platform as a coordinate sourcePoint obEstablishing a body coordinate system o of the remote sensing satellite platformbxbybzbThree-axis orientation of the body coordinate system of the remote sensing satellite platform is same as the centroid coordinate system o of the remote sensing satellite systemsxsyszsPoint to the three axes of (1);
coordinate system of measured remote sensing satellite platform body is ob1xb1yb1zb1And measured remote sensing satellite platform body coordinate system ob1xb1yb1zb1Coordinate origin of the satellite system is identical to the body coordinate system o of the remote sensing satellite platformbxbybzbMeasured remote sensing satellite platform body coordinate system ob1xb1yb1zb1Three-axis pointing and same-measurement centroid coordinate system o of remote sensing satellite systems1xs1ys1zs1Point to the three axes of (1);
using the center of mass of the satellite rotating load as the origin o of coordinatespEstablishing a satellite rotation load body coordinate system opxpypzpWhen the connection error of the rotary joint is zero and the self-rotation angle of the satellite rotary load is zero, the satellite rotary load body coordinate system o pxpypzpThree-axis pointing and measuring remote sensing satellite system centroid coordinate system os1xs1ys1zs1Point to the three axes of (a);
recording the coordinate system of the measured on-satellite rotating load body as op1xp1yp1zp1And measured satellite rotation load body coordinate system op1xp1yp1zp1Coordinate origin of the satellite system is identical to the body coordinate system o of the remote sensing satellite platformpxpypzpOf the measured satellite rotational load body coordinate system o without measurement errorsp1xp1yp1zp1Three axes of the same coordinate system opxpypzpThe three axes of (a) coincide.
3. The method for dynamic precision analysis and error distribution of the rotating load on the remote sensing satellite according to claim 2, wherein the specific process of the second step is as follows:
determining each error link contained in the remote sensing satellite system as follows: a track error link, a platform attitude error link, a magnetic suspension bearing connection error link and an on-satellite rotating load spinning error link;
the track error link consists of two parts: partly control errors of the track, resulting in a desired track coordinate system oo1xo1yo1zo1And a track coordinate system ooxoyozoThere is a relative deviation between them; another part is the determination error of the track, resulting in a track coordinate system ooxoyozoAnd a measured orbital coordinate system oo2xo2yo2zo2There is a relative deviation between them;
the platform attitude error link consists of two parts: one part is attitude control error, which is influenced by attitude interference to cause a system o in the remote sensing satellite platform bxbybzbAnd a measured orbital coordinate system oo2xo2yo2zo2There is a relative deviation between them; the other part is attitude measurement errors including star sensor errors and fiber optic gyroscope errors, which cause a system o of the remote sensing satellite platformbxbybzbAnd measured remote sensing satellite platform body system ob1xb1yb1zb1There is a relative deviation between them;
the magnetic suspension bearing connection error link consists of two parts: one part is the axle gap run-out error, which causes the satellite rotation load body coordinate system opxpypzpRelative measurement remote sensing satellite platform body system ob1xb1yb1zb1Radial nutation precession attitude deviation exists; another part is measurement error, resulting in a satellite rotation load body coordinate system opxpypzpAnd measured on-satellite rotational load body coordinate system op1xp1yp1zp1Has radial attitude deviation therebetweenA difference;
the spin error link of the satellite rotating load consists of two parts, one part is a rotating speed control error, which causes a coordinate system o of the satellite rotating load bodypxpypzpRelative measurement remote sensing satellite platform body system ob1xb1yb1zb1Axial rotation speed errors exist; another part is measurement error, resulting in a satellite rotation load body coordinate system opxpypzpAnd measured on-satellite rotational load body coordinate system op1xp1yp1zp1Axial rotation speed deviation exists between the two parts;
the track error link, the platform attitude error link, the magnetic suspension bearing connection error link and the satellite rotation load spinning error link are connected in series to form a precision error transmission chain, and all the error links act together and are transmitted to the satellite rotation load in an accumulated mode.
4. The method for dynamic precision analysis and error distribution of the rotary load on the satellite of the remote sensing satellite according to claim 3, wherein the specific process of the third step is as follows:
orbital coordinate system ooxoyozoRelative desired orbital coordinate system oo1xo1yo1zo1With errors, the orbital coordinate system o is described by zxy sequenceoxoyozoRelative desired orbital coordinate system oo1xo1yo1zo1Attitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: desired orbital coordinate system oo1xo1yo1zo1Around zo1Angle of rotation of the shaft
To intermediate system i1Intermediate system i1Through an angle delta theta about its x-axis1To intermediate system i2Intermediate system i2Y around an orbital coordinate systemoAxial rotation through delta psi1To the orbital coordinate system ooxoyozo；
Measured orbital coordinate system oo2xo2yo2zo2Relative orbital coordinate system ooxoyozoWith errors, describing the measured orbital coordinate system o by zxy ordero2xo2yo2zo2Relative orbital coordinate system ooxoyozoAttitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: orbital coordinate system ooxoyozoAround zoAngle of rotation of the shaft
To intermediate system i1', intermediate system i1' Angle of rotation about its x-axis Delta theta2To intermediate system i2', intermediate system i2' around the measured orbital coordinate system yo2Angle delta psi of shaft rotation2To the measured orbital coordinate system oo2xo2yo2zo2；
Measured remote sensing satellite platform body system o b1xb1yb1zb1Relative measurementOrbital coordinate system of magnitude oo2xo2yo2zo2Error exists, and the system o of the remote sensing satellite platform is described by using zxy sequenceb1xb1yb1zb1Relative measured orbital coordinate system oo2xo2yo2zo2Attitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: measured orbital coordinate system oo2xo2yo2zo2Around zo2Angle of rotation of the shaft
To intermediate system i1", intermediate system i1"Angle of rotation about its x-axis Δ θ3To intermediate system i2", intermediate system i2"y around the system of remote sensing satellite platformb1Angle delta psi of shaft rotation3To the remote sensing satellite platform body system ob1xb1yb1zb1；
Measured satellite rotation load body coordinate system op1xp1yp1zp1Relative measurement remote sensing satellite platform body system ob1xb1yb1zb1With errors, the global coordinate system o of the measured satellite rotation load is described by zyxp1xp1yp1zp1Relative measurement remote sensing satellite platform body system ob1xb1yb1zb1Attitude error of, rotational Euler angle of
Is recorded as:
euler angle
Is defined as: measured remote sensing satellite platform body system ob1xb1yb1zb1Around zb1Angle delta psi of shaft rotation4To intermediate system i1", intermediate system i1"' over-rotation about its y-axis by an angle Δ θ4To intermediate system i2", intermediate system i2' x around measured satellite rotation load body coordinate systemp1Angle of rotation of the shaft
To the measured on-satellite rotational load body coordinate system op1xp1yp1zp1；
Representing the expected rotation angle of the rotating load on the satellite relative to the remote sensing satellite platform, 
Representing the rotation angle deviation of the on-satellite rotation load relative to the remote sensing satellite platform;
respectively calculating the rotating load direction on the actual satellite and the rotating load direction on the expected satellite in the expected orbit coordinate system oo1xo1yo1zo1Projection z ofcAnd zd；
Specifying actual on-board rotational load pointing in measured on-board rotational load body coordinate system op1xp1yp1zp1Projection of lower is zp1′＝[0,0,1]TThen the rotational load on the satellite is expected to be directed at the desired orbital coordinate system oo1xo1yo1zo1Projection z ofdComprises the following steps:
the actual on-board rotating load is directed at the desired orbital coordinate system oo1xo1yo1zo1Projection z ofcComprises the following steps:
wherein the content of the first and second substances,
representing the orbital coordinate system ooxoyozoRelative to a desired orbital coordinate system oo1xo1yo1zo1The attitude transformation matrix of (a) is,
orbital coordinate system o representing measurementso2xo2yo2zo2Relative to the orbital coordinate system ooxoyozoThe attitude transformation matrix of (a) is,
system o of remote sensing satellite platform body for representing measurementb1xb1yb1zb1Relative to the measured orbital coordinate system oo2xo2yo2zo2The attitude transformation matrix of (a) is,
body coordinate system o of satellite rotation load representing measurementp1xp1yp1zp1System o of remote sensing satellite platform relative to measurementb1xb1yb1zb1The attitude transformation matrix of (2);
combining the formulas (2) to (6) to obtain the projection z of the actual on-satellite rotation load pointing direction under the expected orbit coordinate systemcSolving two vectors zdAnd zcThe included angle Δ Φ between is:
ΔΦ＝arccos(zd Tzc) (7)
Wherein: arccos (·) denotes an inverse cosine computation;
and estimating the pointing accuracy of the rotating load on the satellite by using the included angle delta phi.
5. The method for dynamic precision analysis and error distribution of the rotating load on the remote sensing satellite according to claim 4, wherein the specific process of the fourth step is as follows:
measured remote sensing satellite platform body system o without considering attitude errorb1xb1yb1zb1And a desired orbital coordinate system oo1xo1yo1zo1Coincident and measured on-satellite rotation load body coordinate system op1xp1yp1zp1And the satellite rotation load body coordinate system opxpypzpCoincident and measured on-satellite rotation load body coordinate system op1xp1yp1zp1Relative measurement remote sensing satellite platform body system ob1xb1yb1zb1System o of remote sensing satellite platform with axial rotation and measurement onlyb1xb1yb1zb1Angular velocity relative to the equatorial coordinate system oxyz and the desired orbital coordinate system oo1xo1yo1zo1The angular speeds of the relative geocentric equatorial coordinate system oxyz are consistent, and the measured remote sensing satellite platform body system ob1xb1yb1zb1Angular velocity relative to the equatorial coordinate system oxyz in the desired orbital coordinate system oo1xo1yo1zo1Projection of the lower is [ 0-omega ]0 0]TWherein: omega0Representing orbital angular velocity, the angular velocity vector of the rotating load on the satellite relative to the equatorial coordinate system oxyz is in the desired orbital coordinate system oo1xo1yo1zo1Lower projection omega d1Expressed as:
wherein, the first and the second end of the pipe are connected with each other,
representing a pair of scalars
Taking a time derivative;
when the attitude error is considered, the angular velocity vector of the on-satellite rotating load relative to the equatorial coordinate system oxyz in the coordinate system o is obtained by combining equation (2)p1xp1yp1zp1Expected value ω of downward projectiondComprises the following steps:
wherein the content of the first and second substances,
is to
The transpose of (a) is performed,
is to
The transpose of (a) is performed,
is to
The transpose of (a) is performed,
is to
Transposing; measured remote sensing satellite platform body coordinate system ob1xb1yb1zb1Angular velocity relative to the equatorial coordinate system oxyz in coordinate system ob1xb1yb1zb1Lower projection omegabComprises the following steps: omegab＝[Δωbx -ω0+Δωby Δωbz]TWherein Δ ωbx、Δωby、ΔωbzRespectively indicates that the angular speed error of the coordinate system of the remote sensing satellite platform body is ob1xb1yb1zb1X of systemb1、yb1、zb1Projection under the shaft;
measured satellite rotation load body coordinate system op1xp1yp1zp1Relative measurement remote sensing satellite platform body coordinate system ob1xb1yb1zb1With attitude errors, from the coordinate system ob1xb1yb1zb1Sequentially switched by delta phi through zyx4、Δθ4、
To a coordinate system op1xp1yp1zp1Noting the attitude angular rate of
The angular velocity vector of the rotating load on the satellite relative to the equatorial coordinate system oxyz of the earth is in the coordinate system op1xp1yp1zp1True value ω of the lower projectioncComprises the following steps:
intermediate variable C in equation (10)x、CyAnd CzRespectively expressed as:
the expected value ω is obtained by combining the formulas (9) and (10)dWith the true value omegacThe difference Δ ω is:
Δω＝ωc-ωd (12)
using desired value omegadWith the true value omegacAnd estimating the three-axis attitude stabilization precision of the on-satellite rotation load by the difference delta omega.
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