CN112478200A - Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft - Google Patents

Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft Download PDF

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CN112478200A
CN112478200A CN202011362675.2A CN202011362675A CN112478200A CN 112478200 A CN112478200 A CN 112478200A CN 202011362675 A CN202011362675 A CN 202011362675A CN 112478200 A CN112478200 A CN 112478200A
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spacecraft
combined
coordinate system
attitude
axis
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CN112478200B (en
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段广仁
赵琴
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Harbin Institute of Technology
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    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/242—Orbits and trajectories
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
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Abstract

A posture and orbit coupling control method for identifying all quality parameters of a combined spacecraft belongs to the technical research field of spacecraft control. The invention solves the problems that the efficiency of the existing method for identifying the quality parameters and controlling the attitude and orbit coupling is low and all the quality parameters can not be identified. The attitude and orbit coupling control method can estimate all quality parameters of the combined spacecraft by using the self-adaptive law and can realize attitude and orbit coupling control through the controller. The experimental result shows that the method provided by the invention can ensure that all quality parameters of the combined spacecraft converge to true values, and can realize tracking of the attitude and the orbit trajectory, and the relative position, the relative speed, the relative attitude and the relative angular speed of the combined spacecraft can converge to zero when about 1000s, so that the efficiency of quality parameter identification and attitude-orbit coupling control is improved, and all quality parameters can be identified. The invention can be applied to the identification of the quality parameters of the combined spacecraft and the attitude and orbit coupling control.

Description

Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft
Technical Field
The invention belongs to the technical field of spacecraft control research, and particularly relates to a posture and orbit coupling control method for identifying all quality parameters of a combined spacecraft.
Background
The space on-orbit service can be used for prolonging the service life of the spacecraft, enhancing the performance of the spacecraft and the like, and has important economic value. The space non-cooperative target generally refers to an on-orbit spacecraft which is unknown in morphological structure, operation state, characteristic parameters and the like or is not matched in measurement and behavior, and comprises a fault spacecraft, space garbage, a hostile spacecraft and the like. The mass parameters of the spacecraft include mass, moment of inertia and centroid position. The quality parameters of the non-cooperative target are unknown, so that the quality parameters of the combined spacecraft formed after the non-cooperative target is captured by the service spacecraft are unknown, whether the quality parameters of the combined spacecraft can be identified or not can be identified, and whether the follow-up space task can be accurately completed or not can be directly concerned.
The orbit and the attitude of the spacecraft have complex coupling relations which mainly come from the coupling relations between the attitude dynamics and the orbit dynamics of the spacecraft and the coupling of the attitude and the orbit control system caused by the installation of an actuating mechanism. The attitude and orbit coupling control of the spacecraft has the outstanding advantages of high efficiency, strong maneuverability, high control precision and the like. Attitude and orbit control should not be split apart, but should be treated as a whole.
Aiming at the quality parameter identification and attitude and orbit coupling control of the combined spacecraft, although an effective research framework is provided at present, the existing solution still has a promotion space, and the concrete expression is that the existing method firstly directly designs an estimator to identify all quality parameters, and then designs a controller to carry out attitude and orbit coupling control on the combined spacecraft, or even if the quality parameter identification and the attitude and orbit coupling control can be carried out at the same time, all quality parameters can not be identified. This can lead to inefficient identification and control, or the inability to identify all of the quality parameters, which ultimately affects the rapid and accurate completion of subsequent space missions by the assembled spacecraft. Therefore, it is significant to design an attitude and orbit coupling controller capable of identifying all quality parameters of the combined spacecraft to ensure that the combined spacecraft can complete space tasks quickly and accurately.
Disclosure of Invention
The invention aims to solve the problems that the efficiency of identifying quality parameters and controlling attitude and orbit coupling is low and all the quality parameters cannot be identified in the conventional method, and provides an attitude and orbit coupling control method for identifying all the quality parameters of a combined spacecraft.
The technical scheme adopted by the invention for solving the technical problems is as follows: a posture and orbit coupling control method for identifying all quality parameters of a combined spacecraft comprises the following steps:
establishing a relative attitude and orbit coupling kinetic equation of the combined spacecraft;
step two, obtaining an intermediate variable epsilon related to the estimation error of the quality parameter according to the relative attitude and orbit coupling kinetic equation established in the step onejThe expression of (1);
selecting and storing thruster output data and regression matrix data;
and step four, calculating the control input of the combined spacecraft control system and identifying all quality parameters by using the data stored in the step three according to whether the sufficient condition of quality parameter identification is met.
The invention has the beneficial effects that: the invention provides an attitude and orbit coupling control method for identifying all quality parameters of a combined spacecraft. The experimental result shows that the method provided by the invention can ensure that all quality parameters of the combined spacecraft converge to true values, and can realize tracking of the attitude and the orbit trajectory, and the relative position, the relative speed, the relative attitude and the relative angular speed of the combined spacecraft can converge to zero when about 1000s, so that the efficiency of quality parameter identification and attitude-orbit coupling control is improved, and all quality parameters can be identified.
Drawings
FIG. 1 is a schematic view of the installation of the combined spacecraft and actuator;
in the figure, F1Force generated for the thruster 1, F2Force generated for the thruster 2, F3Force generated for the thruster 3, F4Force generated for the thruster 4, F5Force generated for the thruster 5, F6The force generated for the thruster 6;
FIG. 2 is a diagram of relative positions of an assembled spacecraft for tracking;
FIG. 3 is a relative velocity plot for combined spacecraft tracking;
FIG. 4 is a diagram of relative attitude for combined spacecraft tracking;
FIG. 5 is a graph of relative angular velocity of the combined spacecraft tracking;
FIG. 6 is a diagram of quality identification results;
FIG. 7 is a drawing of J in the moment of inertia matrix11An identification result graph in 0-1 s and 500-1500 s;
FIG. 8 is a J in the moment of inertia matrix22An identification result graph in 0-1 s and 500-1500 s;
FIG. 9 is a J in the moment of inertia matrix33An identification result graph in 0-1 s and 500-1500 s;
FIG. 10 is a drawing of J in the moment of inertia matrix23An identification result graph in 0-1 s and 500-1500 s;
FIG. 11 is a graph of J in the moment of inertia matrix13An identification result graph in 0-1 s and 500-1500 s;
FIG. 12 is a drawing of J in the moment of inertia matrix12An identification result graph in 0-1 s and 500-1500 s;
FIG. 13 is a centroid position ρxAn identification result graph in 0-2 s and 500-1500 s;
FIG. 14 is a centroid position ρyAn identification result graph in 0-2 s and 500-1500 s;
FIG. 15 is a centroid position ρzThe recognition result is shown in 0-2 s and 500-1500 s.
Detailed Description
The first embodiment is as follows: this embodiment will be described with reference to fig. 1. In this embodiment, a method for controlling attitude and orbit coupling for identifying all quality parameters of an integrated spacecraft includes the following steps:
establishing a relative attitude and orbit coupling kinetic equation of the combined spacecraft;
step two, obtaining an intermediate variable epsilon related to the estimation error of the quality parameter according to the relative attitude and orbit coupling kinetic equation established in the step onejThe expression of (1);
selecting and storing thruster output data and regression matrix data;
and step four, calculating the control input of the combined spacecraft control system and identifying all quality parameters by using the data stored in the step three according to whether the sufficient condition of quality parameter identification is met.
Specifically, the invention has the following innovation points:
based on a self-adaptive control method, a control scheme for a combined spacecraft formed after capturing a non-cooperative target is provided, and the method not only can realize attitude and orbit trajectory tracking, but also can identify all quality parameters.
The second embodiment is as follows: the first difference between the present embodiment and the specific embodiment is: the combined spacecraft is composed of a service spacecraft, a mechanical arm and a non-cooperative target.
Other steps and parameters are the same as those in the first embodiment.
The third concrete implementation mode: the first difference between the present embodiment and the specific embodiment is: the specific process of the step one is as follows:
establishing an earth-centered inertial coordinate system O-xiyiziBody coordinate system O of combined spacecraftp-xpypzpVirtual coordinate system Ot-xtytztAnd a reference coordinate system Oo-xoyozo；
The geocentric inertial coordinate system O-xiyiziWith the geocentric O as the origin, xiThe axis being in the equatorial plane and pointing towards the vernal equinox, ziThe axis coincides with the earth's axis of rotation and points in the north, y, directioniAxis and xiAxis and ziThe axis meets the right hand rule;
body coordinate system O of combined spacecraftp-xpypzpWith the mass center of the combined spacecraft as the origin Op，xpAxis, ypAxis and zpThe axis coinciding with the principal axis of inertia, x, of the serving spacecraftpAxis, ypAxis and zpThe axes meet the right-hand rule to form a rectangular coordinate system;
the above-mentionedVirtual coordinate system Ot-xtytztWith the desired position of the combined spacecraft as the origin Ot，xtAxis, ytAxis and ztThe axes and the three axes are mutually perpendicular and satisfy the right hand rule, xtAxis, ytAxis and ztThe orientation of the axis represents the desired attitude of the combined spacecraft;
said reference coordinate system Oo-xoyozoSelecting any point with known position in the combined spacecraft as an origin Oo，xoAxis, yoAxis, zoThe axes are respectively equal to xpAxis, ypAxis, zpThe axes are parallel and point to the same;
for the convenience of calculation, O in FIG. 1 is selectedoThe position is the origin;
reference coordinate system Oo-xoyozoFixedly connected to the combined spacecraft from an origin OoTo OpFormed vector rho ═ OoOp＝[ρxρy ρz]TThe position coordinates of the combined spacecraft centroid in the reference coordinate system are obtained;
representing combined spacecraft by using corrected Rodrigue parameter sigma and by using combined spacecraft body coordinate system Op-xpypzpTo a virtual coordinate system Ot-xtytztOmega is the relative angular velocity between the current angular velocity and the expected angular velocity of the combined spacecraft in the combined spacecraft body coordinate system Op-xpypzpIn the body coordinate system O of the combined spacecraftp-xpypzpIn the method, the relative attitude dynamics and kinematic equations of the combined spacecraft are as follows:
Figure BDA0002804454220000041
wherein,
Figure BDA0002804454220000045
is the first derivative of sigma, G (sigma) is the intermediate variable,
Figure BDA0002804454220000042
the superscript T denotes transposition, I3Is a 3 × 3 identity matrix, S (·) denotes a ═ a for an arbitrary vector1 a2 a3]TAre all provided with
Figure BDA0002804454220000043
Figure BDA0002804454220000046
Is the first derivative of the omega and is,
Figure BDA0002804454220000047
Figure BDA0002804454220000048
for the angular velocity of the combined spacecraft relative to the inertial space in a combined spacecraft body coordinate system Op-xpypzpIs represented by the coordinates in (1) or (b),
Figure BDA0002804454220000049
as a combined spacecraft relative to the earth's center inertial frame O-xiyiziIn a coordinate system Ot-xtytztIs represented by the coordinates in (1) or (b),
Figure BDA00028044542200000410
as a coordinate system Ot-xtytztTo the coordinate system Op-xpypzpJ is a rotational inertia matrix of the combined spacecraft,
Figure BDA0002804454220000044
τ is control force, CrAnd nrAre all intermediate variables;
intermediate variable CrThe expression of (a) is:
Figure BDA00028044542200000411
intermediate variable nrThe expression of (a) is:
Figure BDA00028044542200000412
Figure BDA00028044542200000414
is composed of
Figure BDA00028044542200000413
The first derivative of (a);
p represents a relative position vector of the current position and the expected position of the combined spacecraft, v represents a relative velocity vector of the current velocity and the expected velocity of the combined spacecraft, and the combined spacecraft is positioned in a coordinate system Op-xpypzpThe relative orbit dynamics model of the combined spacecraft is as follows:
Figure BDA0002804454220000051
wherein,
Figure BDA0002804454220000054
is the first derivative of p and is,
Figure BDA0002804454220000055
is the first derivative of v, m is the mass of the assembled spacecraft, μ is the Earth's gravitational constant, rpThe distance from the current position of the combined spacecraft to the earth mass center, f is control force,
Figure BDA0002804454220000056
position vector of the expected position of the spacecraft of the combination relative to the geocentric in a coordinate system Ot-xtytztIs represented by the coordinates in (1) or (b),
Figure BDA0002804454220000057
is composed of
Figure BDA0002804454220000058
The second derivative of (a);
Figure BDA0002804454220000059
Figure BDA00028044542200000510
as a position vector of the current position of the combined spacecraft relative to the earth center in a coordinate system Op-xpypzpCoordinate representation of (1);
the combo spacecraft is driven by six thrusters mounted on the service spacecraft. Each thruster can generate a bidirectional force parallel to the coordinate axis of the combined spacecraft body, and the arrow direction in the figure 1 is the positive direction of the force;
assuming that the service spacecraft is in a cuboid shape, the length, width and height of the service spacecraft are respectively L1、L2、L3Then the control input u is:
Figure BDA0002804454220000052
wherein F is a vector of a force generated by the thruster, and F ═ F1 F2 F3 F4 F5 F6]TThe control input matrix a (ρ) is:
Figure BDA0002804454220000053
where ρ isxFor the combined spacecraft centroid in x of the reference coordinate systemoComponent in the axial direction, pyFor the combined spacecraft centroid in the reference coordinate systemoComponent in the axial direction, pzZ of combined spacecraft centroid in reference coordinate systemoA component in the axial direction;
selecting system states
Figure BDA00028044542200000511
Wherein x is1And x2Is a system state variable, x1＝[pT σT]T,x2＝[vT ωT]TObtaining a relative attitude and orbit coupling dynamic equation of the combined spacecraft according to the formulas (1), (2) and (3), wherein the equation is as follows:
Figure BDA0002804454220000061
wherein,
Figure BDA0002804454220000062
is x1The first derivative of (a) is,
Figure BDA0002804454220000063
is x2First derivative of (A), intermediate variables Lambda, C1、C2The expressions for M and n are respectively:
Figure BDA0002804454220000064
Figure BDA0002804454220000065
other steps and parameters are the same as those in one of the first to second embodiments.
The fourth concrete implementation mode: the third difference between the present embodiment and the specific embodiment is that: the specific process of the second step is as follows:
defining: for any vector b ═ b1 b2 b3]TIs provided with
Figure BDA0002804454220000066
Let theta equal to [ J11 J22 J33 J23 J13 J12]TRepresenting a vector consisting of unknown elements in the matrix of the moment of inertia of the spacecraft in combination, [ m θ ═ m θ [ ]T ρT]TTo a combined spacecraftP is the coordinate of the centroid position of the combined spacecraft,
Figure BDA0002804454220000068
is the estimated value corresponding to the theta,
Figure BDA0002804454220000069
is an estimated value corresponding to m,
Figure BDA00028044542200000610
is an estimated value corresponding to the value of theta,
Figure BDA00028044542200000611
for the estimated value corresponding to p,
Figure BDA00028044542200000612
for quality parameter estimation errors, let xjAnd FjX and F, respectively, at time tjThe value of (2) is obtained according to the formula (4):
Ψ1Θ＝A1F (5)
therein, Ψ1Is a regression matrix, Ψ1＝[N1+N2+N3 -H(F)]，N1、N2、N3And H (F) are all intermediate variables,
Figure BDA0002804454220000067
and is provided with
Figure BDA0002804454220000071
Figure BDA0002804454220000072
Figure BDA0002804454220000073
Order to
Figure BDA0002804454220000074
Therein, Ψ1,jIs a regression matrix Ψ1At tjValue of time epsilonjIs an intermediate variable. EpsilonjIs epsilon at tjThe value of the moment.
Other steps and parameters are the same as those in one of the first to third embodiments.
The fifth concrete implementation mode: the fourth difference between this embodiment and the specific embodiment is that: the specific process of the third step is as follows:
Figure BDA0002804454220000076
to be a stack, t isjRegression matrix data Ψ corresponding to time1,jAnd thruster output data FjAs a data pair (Ψ)1,j,Fj) Is stored in
Figure BDA0002804454220000077
Performing the following steps;
if the current time returns to the matrix data psi1And tjRegression matrix data Ψ corresponding to time1,jThere is a difference and the following inequality is satisfied:
Figure BDA0002804454220000075
where | l | · | | is a norm, and κ is a given positive number, then it will be noted as tj+1Taking the regression matrix data and the thruster output data at the current moment as tj+1Data pair of time instants (Ψ)1,j+1,Fj+1) Deposit onto a stack
Figure BDA0002804454220000078
In (1).
If the current time satisfies the condition in (7), it is recorded as tj+1Coexistence of time of dayStoring the relevant data; if the current time does not satisfy the condition in (7), waiting for the next time to arrive, and if the next time satisfies the formula (7), recording the next time as tj+1(ii) a If the next time does not meet (7), whether the next time can be recorded as t is calculatedj+1. That is, when (7) is satisfied, the subscript + 1.
Other steps and parameters are the same as in one of the first to fourth embodiments.
The sixth specific implementation mode: the fifth embodiment is different from the fifth embodiment in that: the stack
Figure BDA0002804454220000086
Allowing up to q pairs of data to be stored, when stacked
Figure BDA0002804454220000087
When the storage is full, replace
Figure BDA0002804454220000088
The oldest stored data pair.
Other steps and parameters are the same as those in one of the first to fifth embodiments.
The seventh embodiment: the sixth embodiment is different from the sixth embodiment in that: the specific process of the step four is as follows:
virtual control of design composite spacecraft
Figure BDA0002804454220000089
Wherein K1For diagonal positive definite matrix, alpha1And alpha2Virtual control of the stationary relative position and the relative attitude, respectively, the derivative of alpha being
Figure BDA0002804454220000081
When the sufficient condition of quality parameter identification is met, the control input of the combination spacecraft tracking position and attitude trajectory by using the data stored in the step three is as follows:
Figure BDA0002804454220000082
wherein, K2In order to be a positive definite matrix,
Figure BDA0002804454220000083
Figure BDA0002804454220000084
Figure BDA0002804454220000085
wherein,
Figure BDA00028044542200000810
is rhoxIs determined by the estimated value of (c),
Figure BDA00028044542200000811
is rhoyIs determined by the estimated value of (c),
Figure BDA00028044542200000812
is rhozAn estimated value of (d);
all quality parameters are identified by the following adaptation law:
Figure BDA0002804454220000091
wherein the intermediate variable Ψ2＝[Ν1+N2'+N3' -H(F)]Γ and K3Are all positive definite matrixes;
when the sufficient condition for identifying the quality parameters is not met, identifying all the quality parameters by adopting the following self-adaptive law;
Figure BDA0002804454220000092
wherein,
Figure BDA0002804454220000095
is composed of
Figure BDA0002804454220000096
The first derivative of (a).
And after the real-time verification that the sufficient conditions for identifying the quality parameters are met, switching the self-adaptive law (10) to (9).
Other steps and parameters are the same as those in one of the first to sixth embodiments.
The specific implementation mode is eight: the seventh embodiment is different from the seventh embodiment in that: the sufficient conditions for identifying the quality parameters are as follows:
rank (Ω) 10, where the intermediate variable matrix
Figure BDA0002804454220000093
rank () denotes the rank of the matrix in parentheses.
When the sufficiency condition is satisfied, all quality parameters can be identified. Remarking: the spacecraft can meet the sufficient conditions in the conventional maneuvering.
Other steps and parameters are the same as those in one of the first to seventh embodiments.
The specific implementation method nine: the eighth embodiment is different from the eighth embodiment in that: said coordinate system Ot-xtytztTo the coordinate system Op-xpypzpOf the rotation matrix
Figure BDA0002804454220000097
The expression of (a) is:
Figure BDA0002804454220000094
other steps and parameters are the same as those in one to eight of the embodiments.
The following examples were used to demonstrate the beneficial effects of the present invention:
the first embodiment is as follows:
the attitude and orbit coupling control method for identifying all quality parameters of the combined spacecraft is prepared according to the specific implementation mode:
the real mass, the moment of inertia and the mass center of the combined spacecraft are as follows: m is 1000kg, and m is 1000kg,
Figure BDA0002804454220000101
ρ＝[1.5 1.5 2.5]Tand m is selected. Size L of the service spacecraft1＝L2＝L34 m. The controller parameter is selected to be K1＝0.008I6，K2＝4I6，K3＝0.04I10. And (4) maneuvering the combined spacecraft to a circular orbit with the orbit radius of 6628km by using a thruster, and realizing the ground orientation.
According to the steps from one to four, the control input and the self-adaptation law of the combined spacecraft can be obtained. It can be seen from fig. 2-5 that the relative position, the relative velocity, the relative attitude, and the relative angular velocity of the combined spacecraft converge to zero at about 1000s, that is, the combined spacecraft realizes the tracking of the reference trajectory.
Meanwhile, the mass, the moment of inertia and the position of the mass center of the combined spacecraft can be seen to converge to the true values through the images 6-15, and thus, the identification of all mass parameters is realized.
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 (8)

1. A posture and orbit coupling control method for identifying all quality parameters of a combined spacecraft is characterized by comprising the following steps:
establishing a relative attitude and orbit coupling kinetic equation of the combined spacecraft;
step two, obtaining an intermediate variable epsilon related to the estimation error of the quality parameter according to the relative attitude and orbit coupling kinetic equation established in the step onejThe expression of (1);
selecting and storing thruster output data and regression matrix data;
and step four, calculating the control input of the combined spacecraft control system and identifying all quality parameters by using the data stored in the step three according to whether the sufficient condition of quality parameter identification is met.
2. An attitude and orbit coupling control method for the identification of all the quality parameters of the spacecraft assembly according to claim 1, wherein the spacecraft assembly is composed of a service spacecraft, a robot arm and a non-cooperative target.
3. The attitude and orbit coupling control method for identifying all the quality parameters of the combined spacecraft according to claim 2, wherein the specific process of the first step is as follows:
establishing an earth-centered inertial coordinate system O-xiyiziBody coordinate system O of combined spacecraftp-xpypzpVirtual coordinate system Ot-xtytztAnd a reference coordinate system Oo-xoyozo；
The geocentric inertial coordinate system O-xiyiziWith the geocentric O as the origin, xiThe axis being in the equatorial plane and pointing towards the vernal equinox, ziThe axis coincides with the earth's axis of rotation and points in the north, y, directioniAxis and xiAxis and ziThe axis meets the right hand rule;
body coordinate system O of combined spacecraftp-xpypzpWith the mass center of the combined spacecraft as the origin Op，xpAxis, ypAxis and zpThe axis coinciding with the principal axis of inertia, x, of the serving spacecraftpAxis, ypAxis and zpThe axes meet the right-hand rule to form a rectangular coordinate system;
said virtual coordinate system Ot-xtytztWith the desired position of the combined spacecraft as the origin Ot，xtAxis, ytAxis and ztThe axes and the three axes are mutually perpendicular and satisfy the right hand rule, xtAxis, ytAxis and ztThe orientation of the axis represents the desired attitude of the combined spacecraft;
said reference coordinate system Oo-xoyozoSelecting any point with known position in the combined spacecraft as an origin Oo，xoAxis, yoAxis, zoThe axes are respectively equal to xpAxis, ypAxis, zpThe axes are parallel and point to the same;
representing combined spacecraft by using corrected Rodrigue parameter sigma and by using combined spacecraft body coordinate system Op-xpypzpTo a virtual coordinate system Ot-xtytztOmega is the relative angular velocity between the current angular velocity and the expected angular velocity of the combined spacecraft in the combined spacecraft body coordinate system Op-xpypzpIn the body coordinate system O of the combined spacecraftp-xpypzpIn the method, the relative attitude dynamics and kinematic equations of the combined spacecraft are as follows:
Figure FDA0002804454210000021
wherein,
Figure FDA0002804454210000022
is the first derivative of sigma, G (sigma) is the intermediate variable,
Figure FDA0002804454210000023
the superscript T denotes transposition, I3Is a 3 × 3 identity matrix, S (-) denotesMeaning vector a ═ a1 a2 a3]TAre all provided with
Figure FDA0002804454210000024
Figure FDA0002804454210000025
Is the first derivative of the omega and is,
Figure FDA0002804454210000026
Figure FDA0002804454210000027
for the angular velocity of the combined spacecraft relative to the inertial space in a combined spacecraft body coordinate system Op-xpypzpIs represented by the coordinates in (1) or (b),
Figure FDA0002804454210000028
as a combined spacecraft relative to the earth's center inertial frame O-xiyiziIn a coordinate system Ot-xtytztIs represented by the coordinates in (1) or (b),
Figure FDA0002804454210000029
as a coordinate system Ot-xtytztTo the coordinate system Op-xpypzpJ is the rotational inertia matrix of the combined spacecraft, tau is the control force, CrAnd nrAre all intermediate variables;
intermediate variable CrThe expression of (a) is:
Figure FDA00028044542100000210
intermediate variable nrThe expression of (a) is:
Figure FDA00028044542100000211
Figure FDA00028044542100000212
is composed of
Figure FDA00028044542100000213
The first derivative of (a);
p represents a relative position vector of the current position and the expected position of the combined spacecraft, v represents a relative velocity vector of the current velocity and the expected velocity of the combined spacecraft, and the combined spacecraft is positioned in a coordinate system Op-xpypzpThe relative orbit dynamics model of the combined spacecraft is as follows:
Figure FDA00028044542100000214
wherein,
Figure FDA00028044542100000215
is the first derivative of p and is,
Figure FDA00028044542100000216
is the first derivative of v, m is the mass of the assembled spacecraft, μ is the Earth's gravitational constant, rpThe distance from the current position of the combined spacecraft to the earth mass center, f is control force,
Figure FDA00028044542100000217
position vector of the expected position of the spacecraft of the combination relative to the geocentric in a coordinate system Ot-xtytztIs represented by the coordinates in (1) or (b),
Figure FDA00028044542100000218
is composed of
Figure FDA00028044542100000219
The second derivative of (a);
Figure FDA00028044542100000220
Figure FDA00028044542100000221
as a position vector of the current position of the combined spacecraft relative to the earth center in a coordinate system Op-xpypzpCoordinate representation of (1);
assuming that the service spacecraft is in a cuboid shape, the length, width and height of the service spacecraft are respectively L1、L2、L3Then the control input u is:
Figure FDA0002804454210000031
wherein F is a vector of a force generated by the thruster, and F ═ F1 F2 F3 F4 F5 F6]TThe control input matrix a (ρ) is:
Figure FDA0002804454210000032
where ρ isxFor the combined spacecraft centroid in x of the reference coordinate systemoComponent in the axial direction, pyFor the combined spacecraft centroid in the reference coordinate systemoComponent in the axial direction, pzZ of combined spacecraft centroid in reference coordinate systemoA component in the axial direction;
selecting system states
Figure FDA0002804454210000033
Wherein x is1And x2Is a system state variable, x1＝[pT σT]T,x2＝[vTωT]TObtaining a relative attitude and orbit coupling dynamic equation of the combined spacecraft according to the formulas (1), (2) and (3), wherein the equation is as follows:
Figure FDA0002804454210000034
wherein,
Figure FDA0002804454210000035
is x1The first derivative of (a) is,
Figure FDA0002804454210000036
is x2First derivative of (A), intermediate variables Lambda, C1、C2The expressions for M and n are respectively:
Figure FDA0002804454210000037
Figure FDA0002804454210000038
4. the attitude and orbit coupling control method for identifying all the quality parameters of the composite spacecraft according to claim 3, wherein the specific process of the second step is as follows:
defining: for any vector b ═ b1 b2 b3]TIs provided with
Figure FDA0002804454210000041
Let theta equal to [ J11 J22 J33 J23 J13 J12]TRepresenting a vector consisting of unknown elements in the matrix of the moment of inertia of the spacecraft in combination, [ m θ ═ m θ [ ]T ρT]TIs a vector formed by unknown mass parameters of the combined spacecraft, rho is a centroid position coordinate of the combined spacecraft,
Figure FDA0002804454210000042
is the estimated value corresponding to the theta,
Figure FDA0002804454210000043
is an estimated value corresponding to m,
Figure FDA0002804454210000044
is an estimated value corresponding to the value of theta,
Figure FDA0002804454210000045
for the estimated value corresponding to p,
Figure FDA0002804454210000046
for quality parameter estimation errors, let xjAnd FjX and F, respectively, at time tjThe value of (2) is obtained according to the formula (4):
Ψ1Θ＝A1F (5)
therein, Ψ1Is a regression matrix, Ψ1＝[N1+N2+N3 -H(F)]，N1、N2、N3And H (F) are all intermediate variables,
Figure FDA0002804454210000047
and is provided with
Figure FDA0002804454210000048
Figure FDA0002804454210000049
Figure FDA00028044542100000410
Order to
Figure FDA0002804454210000051
Therein, Ψ1,jIs a regression matrix Ψ1At tjValue of time epsilonjIs an intermediate variable.
5. The attitude and orbit coupling control method for identifying all the quality parameters of the composite spacecraft according to claim 4, wherein the specific process of the third step is as follows:
Figure FDA0002804454210000052
to be a stack, t isjRegression matrix data Ψ corresponding to time1,jAnd thruster output data FjAs a data pair (Ψ)1,j,Fj) Is stored in
Figure FDA0002804454210000053
Performing the following steps;
if the current time returns to the matrix data psi1And tjRegression matrix data Ψ corresponding to time1,jThere is a difference and the following inequality is satisfied:
Figure FDA0002804454210000054
where | l | · | | is a norm, and κ is a given positive number, then it will be noted as tj+1Taking the regression matrix data and the thruster output data at the current moment as tj+1Data pair of time instants (Ψ)1,j+1,Fj+1) Deposit onto a stack
Figure FDA00028044542100000510
In (1).
6. A method as claimed in claim 5, wherein the stack is used for attitude and orbit coupling control of the overall quality parameter identification of the spacecraft assembly
Figure FDA0002804454210000059
Allowing up to q pairs of data to be stored, when stacked
Figure FDA00028044542100000511
When the storage is full, replace
Figure FDA00028044542100000512
The oldest stored data pair.
7. The attitude and orbit coupling control method for identifying all the quality parameters of the composite spacecraft according to claim 6, wherein the specific process of the fourth step is as follows:
virtual control of design composite spacecraft
Figure FDA0002804454210000055
Wherein K1For diagonal positive definite matrix, alpha1And alpha2Virtual control of the stationary relative position and the relative attitude, respectively, the derivative of alpha being
Figure FDA0002804454210000056
When the sufficient condition of quality parameter identification is met, the control input of the combination spacecraft tracking position and attitude trajectory by using the data stored in the step three is as follows:
Figure FDA0002804454210000057
wherein, K2In order to be a positive definite matrix,
Figure FDA0002804454210000058
Figure FDA0002804454210000061
Figure FDA0002804454210000062
wherein,
Figure FDA0002804454210000063
is rhoxIs determined by the estimated value of (c),
Figure FDA0002804454210000064
is rhoyIs determined by the estimated value of (c),
Figure FDA0002804454210000065
is rhozAn estimated value of (d);
all quality parameters are identified by the following adaptation law:
Figure FDA0002804454210000066
wherein the intermediate variable Ψ2＝[Ν1+N2'+N3' -H(F)]Γ and K3Are all positive definite matrixes;
when the sufficient condition for identifying the quality parameters is not met, identifying all the quality parameters by adopting the following self-adaptive law;
Figure FDA0002804454210000067
wherein,
Figure FDA0002804454210000068
is composed of
Figure FDA0002804454210000069
The first derivative of (a).
8. An attitude and orbit coupling control method for identifying all quality parameters of a composite spacecraft as claimed in claim 7, wherein the sufficient condition for identifying the quality parameters is:
rank (Ω) 10, where the intermediate variable matrix
Figure FDA00028044542100000610
rank () denotes the rank of the matrix in parentheses.
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