CN112478200B - 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 PDFInfo
- Publication number
- CN112478200B CN112478200B CN202011362675.2A CN202011362675A CN112478200B CN 112478200 B CN112478200 B CN 112478200B CN 202011362675 A CN202011362675 A CN 202011362675A CN 112478200 B CN112478200 B CN 112478200B
- Authority
- CN
- China
- Prior art keywords
- spacecraft
- combined
- coordinate system
- axis
- combined spacecraft
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 230000008878 coupling Effects 0.000 title claims abstract description 36
- 238000010168 coupling process Methods 0.000 title claims abstract description 36
- 238000005859 coupling reaction Methods 0.000 title claims abstract description 36
- 238000000034 method Methods 0.000 title claims abstract description 29
- 239000011159 matrix material Substances 0.000 claims description 38
- 230000014509 gene expression Effects 0.000 claims description 11
- 239000000126 substance Substances 0.000 claims description 10
- 230000008569 process Effects 0.000 claims description 8
- 238000013461 design Methods 0.000 claims description 5
- 239000002131 composite material Substances 0.000 claims description 3
- 230000006978 adaptation Effects 0.000 claims description 2
- 230000017105 transposition Effects 0.000 claims description 2
- 238000011160 research Methods 0.000 abstract description 3
- 238000010586 diagram Methods 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000009434 installation Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000002457 bidirectional effect Effects 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 238000005259 measurement Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 230000000877 morphologic effect Effects 0.000 description 1
- 238000012795 verification Methods 0.000 description 1
Images
Classifications
-
- 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
- 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
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
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
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 spacecraft and the actuator of the assembly;
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, F5A force generated by 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 view 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 view of 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 matrix23Differentiation at 0-1 s and 500-1500 sIdentifying a result graph;
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 ρxThe recognition result chart is set at 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;
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;
for convenience of calculation, selectingO in FIG. 1oThe position is an 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 speed between the current angular speed and the expected angular speed of the combined spacecraft in a 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:
wherein the content of the first and second substances,is the first derivative of sigma, G (sigma) is the intermediate variable,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
Is the first derivative of the omega and is, 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),as a combined spacecraft relative to the earth's center inertial frame O-xiyiziIn a coordinate system Ot-xtytztThe coordinate of (2) is represented by,as a coordinate system Ot-xtytztTo the coordinate system Op-xpypzpJ is a rotational inertia matrix of the combined spacecraft,τ is control force, CrAnd nrAre all intermediate variables;
intermediate variable CrThe expression of (a) is:intermediate variable nrThe expression of (a) is: is composed ofThe first derivative of (a);
representing the current position of the combined spacecraft relative to the expected position by pA position vector, v represents the relative velocity vector of the current velocity and the expected velocity of the combined spacecraft, and is positioned in a coordinate system Op-xpypzpThe relative orbit dynamics model of the combined spacecraft is as follows:
wherein the content of the first and second substances,is the first derivative of p and is,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,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),is composed ofThe second derivative of (a);
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:
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:
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 statesWherein 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 relative attitude and orbit coupling dynamic equation is as follows:
wherein the content of the first and second substances,is x1The first derivative of (a) is,is x2First derivative of (A), intermediate variables Lambda, C1、C2The expressions for M and n are respectively:
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:
Let theta equal to [ J11 J22 J33 J23 J13 J12]TRepresenting a vector composed 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,is the estimated value corresponding to the theta,is an estimated value corresponding to m,is an estimated value corresponding to the value of theta,is rho pairThe value of the corresponding estimate is,for quality parameter estimation errors, let xjAnd FjX and F, respectively, at time tjThe value of (b) 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,
and is provided with
Order to
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:
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 inPerforming 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:
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 stackIn (1).
If the current time satisfies the condition in (7), it is recorded as tj+1Time and store 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 stackAt most q pairs of data are allowed to be stored, when heapStackWhen the storage is full, replaceThe oldest stored data pair.
Other steps and parameters are the same as those in one of the first to fifth embodiments.
The seventh concrete implementation mode: 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 spacecraftWherein K1For diagonal positive definite matrix, alpha1And alpha2Virtual control of the stationary relative position and the relative attitude, respectively, the derivative of alpha being
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:
wherein the content of the first and second substances,is rhoxIs determined by the estimated value of (c),is rhoyIs determined by the estimated value of (c),is rhozAn estimated value of (d);
all quality parameters are identified by the following adaptation law:
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;
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 matrixrank () 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 above 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 a coordinate system Op-xpypzpOf the rotation matrixThe expression of (a) is:
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,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. By using a thrusterThe combined spacecraft maneuvered to a circular orbit with an orbit radius of 6628km and achieved 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 (2)
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; the combined spacecraft consists of a service spacecraft, a mechanical arm and a non-cooperative target;
the specific process for establishing the relative attitude and orbit coupling kinetic equation of the combined spacecraft comprises the following steps:
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;
the spacecraft body coordinate system O of the combination bodyp-xpypzpWith the mass center of the combined spacecraft as the origin Op,xpAxis, ypAxis and zpThe axis coincides with the principal axis of inertia of the serving spacecraft, xpAxis, 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:
wherein the content of the first and second substances,is the first derivative of sigma, G (sigma) is the intermediate variable,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
Is the first derivative of the omega and is, 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),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),as a coordinate system Ot-xtytztTo the coordinate system Op-xpypzpOf the rotation matrixJ is the rotational inertia matrix of the combined spacecraft, τ is the control force, CrAnd nrAre all intermediate variables;
intermediate variable CrThe expression of (a) is:intermediate variable nrThe expression of (a) is: is composed ofThe 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:
wherein the content of the first and second substances,is the first derivative of p and is,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,is a combined bodyPosition vector of the expected position of the spacecraft relative to the geocentric is in a coordinate system Ot-xtytztIs represented by the coordinates in (1) or (b),is composed ofThe second derivative of (a);
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:
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:
where ρ isxX of combined spacecraft centroid in 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 statesWherein 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:
wherein the content of the first and second substances,is x1The first derivative of (a) is,is x2First derivative of (A), intermediate variables Lambda, C1、C2The expressions for M and n are respectively:
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);
the specific process of the second step is as follows:
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,is the estimated value corresponding to the theta,is an estimated value corresponding to m,is an estimated value corresponding to the value of theta,for the estimated value corresponding to p,estimate error for quality parameter, let xjAnd FjX and F, respectively, at time tjThe value of (b) 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,
and is provided with
Order to
Therein, Ψ1,jIs a regression matrix Ψ1At tjValue of the time of day epsilonjIs an intermediate variable;
selecting and storing thruster output data and regression matrix data;
the specific process of the third step is as follows:
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 inPerforming 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:
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 (psi)1,j+1,Fj+1) Deposit onto a stackPerforming the following steps;
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 specific process of the step four is as follows:
virtual control of design composite spacecraftWherein K1Is a diagonal positive definite matrix, alpha1And alpha2Respectively, the derivative of alpha is
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:
wherein the content of the first and second substances,is rhoxIs determined by the estimated value of (c),is rhoyIs determined by the estimated value of (c),is rhozAn estimated value of (d);
all quality parameters are identified by the following adaptation law:
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;
the sufficient conditions for identifying the quality parameters are as follows:
2. The attitude and orbit coupling control method for the identification of all the quality parameters of the composite spacecraft as claimed in claim 1, wherein said stack is composed ofAllowing up to q pairs of data to be stored, when stackedWhen the storage is full, replaceThe oldest stored data pair.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN202011362675.2A CN112478200B (en) | 2020-11-27 | 2020-11-27 | Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN202011362675.2A CN112478200B (en) | 2020-11-27 | 2020-11-27 | Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft |
Publications (2)
Publication Number | Publication Date |
---|---|
CN112478200A CN112478200A (en) | 2021-03-12 |
CN112478200B true CN112478200B (en) | 2022-06-14 |
Family
ID=74936488
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
CN202011362675.2A Active CN112478200B (en) | 2020-11-27 | 2020-11-27 | Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft |
Country Status (1)
Country | Link |
---|---|
CN (1) | CN112478200B (en) |
Families Citing this family (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN113485407A (en) * | 2021-08-14 | 2021-10-08 | 苏州吉天星舟空间技术有限公司 | Attitude and orbit coupling control method for identifying all quality parameters of spacecraft |
CN115057005B (en) * | 2022-06-07 | 2023-04-25 | 哈尔滨工业大学 | Large annular space structure attitude and rail shape integrated distributed control method |
Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN102620886A (en) * | 2012-03-27 | 2012-08-01 | 南京航空航天大学 | Two-step in-orbit recognition rotary inertia estimation method for combined spacecraft |
CN103235598A (en) * | 2013-05-14 | 2013-08-07 | 北京理工大学 | Method for regulating propeller direction to point to combined-body spacecraft centroid |
CN104252574A (en) * | 2014-07-17 | 2014-12-31 | 西北工业大学 | Space tethered capturing system based non-cooperative target quality identification method |
CN107529388B (en) * | 2013-02-22 | 2016-09-07 | 上海新跃仪表厂 | A kind of noncooperative target assembly Spacecraft Attitude Control method |
CN106502261A (en) * | 2016-12-26 | 2017-03-15 | 西北工业大学 | The identification control integral method of star of receiving reconstruct fault satellites attitude control function |
CN110146224A (en) * | 2019-05-22 | 2019-08-20 | 哈尔滨工业大学 | A method of identification assembly spacecraft mass, centroid position and inertial tensor |
CN110844121A (en) * | 2019-10-22 | 2020-02-28 | 西北工业大学深圳研究院 | Cooperative game control method for cooperative transportation of on-orbit assembly spacecraft |
Family Cites Families (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US7110915B2 (en) * | 2003-05-06 | 2006-09-19 | Edward Wilson | Multiple concurrent recursive least squares identification with application to on-line spacecraft mass-property identification |
-
2020
- 2020-11-27 CN CN202011362675.2A patent/CN112478200B/en active Active
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN102620886A (en) * | 2012-03-27 | 2012-08-01 | 南京航空航天大学 | Two-step in-orbit recognition rotary inertia estimation method for combined spacecraft |
CN107529388B (en) * | 2013-02-22 | 2016-09-07 | 上海新跃仪表厂 | A kind of noncooperative target assembly Spacecraft Attitude Control method |
CN103235598A (en) * | 2013-05-14 | 2013-08-07 | 北京理工大学 | Method for regulating propeller direction to point to combined-body spacecraft centroid |
CN104252574A (en) * | 2014-07-17 | 2014-12-31 | 西北工业大学 | Space tethered capturing system based non-cooperative target quality identification method |
CN106502261A (en) * | 2016-12-26 | 2017-03-15 | 西北工业大学 | The identification control integral method of star of receiving reconstruct fault satellites attitude control function |
CN110146224A (en) * | 2019-05-22 | 2019-08-20 | 哈尔滨工业大学 | A method of identification assembly spacecraft mass, centroid position and inertial tensor |
CN110844121A (en) * | 2019-10-22 | 2020-02-28 | 西北工业大学深圳研究院 | Cooperative game control method for cooperative transportation of on-orbit assembly spacecraft |
Non-Patent Citations (1)
Title |
---|
组合航天器的姿态控制与结构鲁棒控制分配(英文);黄秀韦等;《控制理论与应用》;20180908;第35卷(第10期);1447-1457 * |
Also Published As
Publication number | Publication date |
---|---|
CN112478200A (en) | 2021-03-12 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
CN110673620B (en) | Four-rotor unmanned aerial vehicle air line following control method based on deep reinforcement learning | |
CN107490965B (en) | Multi-constraint trajectory planning method for space free floating mechanical arm | |
CN112478200B (en) | Attitude and orbit coupling control method for identifying all quality parameters of combined spacecraft | |
Spletzer et al. | Cooperative localization and control for multi-robot manipulation | |
CN110347170B (en) | Reusable carrier reentry segment robust fault-tolerant guidance control system and working method | |
CN109426147B (en) | Adaptive gain adjustment control method for combined spacecraft after satellite acquisition | |
CN110550240B (en) | Method for intercepting multi-star cooperative game | |
CN109856972B (en) | Robust fault-tolerant tracking control method for unmanned helicopter | |
CN111624878B (en) | Integral sliding mode acquisition method and system for autonomous water surface robot trajectory tracking | |
Lu et al. | Adaptive coordinated control of uncertain free-floating space manipulators with prescribed control performance | |
CN107992062B (en) | Spatial high-dynamic target high-precision attitude tracking control method based on hybrid actuating mechanism | |
CN115639841A (en) | Unmanned aerial vehicle cluster formation control system and control method based on robust containment | |
Zhang et al. | Output-feedback super-twisting control for line-of-sight angles tracking of non-cooperative target spacecraft | |
Yang et al. | Trajectory planning of dual-arm space robots for target capturing and base manoeuvring | |
CN109623812B (en) | Mechanical arm trajectory planning method considering spacecraft body attitude motion | |
Jin et al. | Observer-based fixed-time tracking control for space robots in task space | |
Zhao et al. | Concurrent learning adaptive finite-time control for spacecraft with inertia parameter identification under external disturbance | |
CN111813140A (en) | High-precision trajectory tracking control method for quad-rotor unmanned aerial vehicle | |
Du et al. | Learning to control a free-floating space robot using deep reinforcement learning | |
Yi et al. | Anti-disturbance control of a quadrotor manipulator with tiltable rotors based on integral sliding mode control | |
Shi et al. | Modeling and simulation of space robot visual servoing for autonomous target capturing | |
CN116277036B (en) | Rapid fault-tolerant vibration suppression control method for flexible-base and flexible-arm space robot | |
CN113220007A (en) | Flexible spacecraft finite time attitude cooperative control method for executing mechanism faults | |
CN109918706B (en) | Generalized dynamics-based satellite-antenna coupling system path planning algorithm | |
CN109507875B (en) | Euler rotary satellite attitude maneuver hierarchical saturation PID control method |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PB01 | Publication | ||
PB01 | Publication | ||
SE01 | Entry into force of request for substantive examination | ||
SE01 | Entry into force of request for substantive examination | ||
GR01 | Patent grant | ||
GR01 | Patent grant |