CN112455726B - Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method - Google Patents

Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method Download PDF

Info

Publication number
CN112455726B
CN112455726B CN202011434349.8A CN202011434349A CN112455726B CN 112455726 B CN112455726 B CN 112455726B CN 202011434349 A CN202011434349 A CN 202011434349A CN 112455726 B CN112455726 B CN 112455726B
Authority
CN
China
Prior art keywords
satellite
attitude
coordinate system
quaternion
angular velocity
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
Application number
CN202011434349.8A
Other languages
Chinese (zh)
Other versions
CN112455726A (en
Inventor
曲友阳
刘萌萌
李峰
童鑫
戴路
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Chang Guang Satellite Technology Co Ltd
Original Assignee
Chang Guang Satellite Technology Co Ltd
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Chang Guang Satellite Technology Co Ltd filed Critical Chang Guang Satellite Technology Co Ltd
Priority to CN202011434349.8A priority Critical patent/CN112455726B/en
Publication of CN112455726A publication Critical patent/CN112455726A/en
Application granted granted Critical
Publication of CN112455726B publication Critical patent/CN112455726B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

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
    • B64G1/28—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect
    • B64G1/285—Guiding or controlling apparatus, e.g. for attitude control using inertia or gyro effect using momentum wheels

Landscapes

  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Abstract

A multi-point imaging fast maneuvering control method for a low-orbit optical remote sensing satellite belongs to the technical field of aerospace, and achieves that the attitude of the satellite is controlled to a desired value in the shortest time under the condition that the moment of a satellite reaction flywheel and the angular momentum thereof are limited, so that an optical load has a target in an imaging state. The method comprises the following steps: performing kinematic dynamics modeling on the low-orbit optical remote sensing satellite; a satellite attitude planning method; transforming the attitude planning result from the orbit coordinate system to an inertial coordinate system; establishing a satellite error kinematics kinetic equation by combining the attitude planning information; the attitude of the satellite is divided into an inner loop and an outer loop, a controller is established, and the multipoint imaging rapid maneuvering control of the low-orbit optical remote sensing satellite is realized. The method has the advantages that the method has the rapidity and the high-precision attitude stability of attitude maneuver, the satellite attitude can be converged to the planned attitude stably in the attitude maneuver process, the control method is simple in structure and easy to realize, and the method can be applied to engineering practice.

Description

Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method
Technical Field
The patent belongs to the technical field of aerospace, and particularly relates to a low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method.
Background
With the rapid development of the low-orbit satellite constellation industry, the low-orbit optical remote sensing satellite becomes a pillar in the remote sensing information industry due to the high-quality and large-breadth imaging capability of the low-orbit optical remote sensing satellite. The maneuvering capability and the stability of the satellite attitude control system directly influence the imaging time and the imaging quality of the satellite, and are one of the important indexes of the optical remote sensing satellite system. In the field of satellite control engineering, the stability and the rapidity of a system are mutually contradictory, and the common control method is difficult to realize both the stability and the rapidity. Therefore, the research on the quick maneuvering control method of the low-orbit optical remote sensing satellite multi-point imaging is of great significance.
Creame and the like adopt a BCB type path planning scheme for an attitude maneuver path in a moon imaging task of a small American moon detector, but the control precision of the attitude maneuver path is difficult to meet the requirement of an optical remote sensing satellite. Michael et al propose a minimum time maneuver control method in a james weber space telescope control system, the attitude planning of which is continuous and smooth, and although good attitude control accuracy is ensured, maneuver time is prolonged. In addition, a plurality of schemes for determining the optimal path of the actual maneuvering of the satellite through a multi-objective optimization algorithm exist. The method can not realize high-precision quick maneuvering control of the optical remote sensing satellite under the condition that the moment and the angular momentum of a flywheel reacting with the satellite are limited.
Disclosure of Invention
In order to solve the problems in the prior art, the invention provides a low-orbit optical remote sensing satellite multipoint imaging rapid maneuvering control method, which realizes that the satellite attitude is controlled to a desired value in the shortest time under the condition that the satellite reaction flywheel moment and the angular momentum thereof are limited, so that the optical load has the target of an imaging state.
The technical scheme adopted by the invention for solving the technical problem is as follows:
a low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method comprises the following steps:
the method comprises the following steps: performing kinematic dynamics modeling on the low-orbit optical remote sensing satellite: definition FIRepresenting an inertial coordinate system, FBRepresenting a satellite body coordinate system; satellite body coordinate system FBRelative to an inertial frame FIIs expressed as
Figure BDA0002827641690000021
Body coordinate system FBRelative inertial frame FIIs expressed as a unit quaternion
Figure BDA0002827641690000022
And satisfy the constraint condition
Figure BDA00028276416900000221
Wherein q is0Is the scalar part of a quaternion Q, Q ═ Q1,q2,q3) Is the vector portion of the quaternion Q,
Figure BDA0002827641690000023
representing the kinematics and dynamics of an n-dimensional real vector space, a satelliteThe equation is:
Figure BDA0002827641690000024
Figure BDA0002827641690000025
in the formula:
Figure BDA0002827641690000026
the moment of inertia of the satellite is represented by a positive definite matrix; i is3Is a 3 × 3 identity matrix; u is the control moment of the reaction flywheel;
Figure BDA0002827641690000027
is the total angular momentum of the reaction flywheel;
Figure BDA0002827641690000028
as an antisymmetric matrix, for arbitrary vectors
Figure BDA0002827641690000029
Satisfy s (x) y ═ x × y, where x represents a vector cross product;
the expected attitude of the satellite is defined as the attitude direction of the coordinate system of the satellite body relative to the inertial coordinate system, and quaternion is carried out through the expected attitude
Figure BDA00028276416900000210
Represents; q. q.sd0Quaternion Q for the desired attitudedScalar part of qdQuaternion Q for the desired attitudedThe vector portion of (1); the attitude tracking error is defined as the error quaternion:
Figure BDA00028276416900000211
in the formula:
Figure BDA00028276416900000212
Figure BDA00028276416900000213
representing quaternion multiplication, qe0Is an error quaternion QeScalar part of qeIs an error quaternion QeThe vector portion of (1); the angular velocity tracking error is:
ωe＝ω-R(Qe)ωd
in the formula: omegadIs the desired angular velocity of the satellite; rotation matrix R (Q)e) The following relationships exist:
Figure BDA00028276416900000214
Figure BDA00028276416900000215
and satisfies the constraint condition | | | R (Q)e)||＝1；
Step two: the satellite attitude planning method comprises the following steps: definition FoRepresenting an orbital coordinate system, a satellite body coordinate system FBRelative to the orbital coordinate system FoIs expressed as
Figure BDA00028276416900000216
Body coordinate system FBRelative orbit coordinate system FoUsing unit quaternion
Figure BDA00028276416900000217
Expressing and satisfying quaternion constraint conditions; the attitude of the satellite multi-point imaging task is that the satellite continuously performs side-sway maneuver; obtaining a satellite body coordinate system F through attitude planningBRelative orbit coordinate system FoDesired attitude quaternion
Figure BDA00028276416900000218
Desired angular velocity
Figure BDA00028276416900000219
And expected angular acceleration
Figure BDA00028276416900000220
The attitude planning is realized by an improved differentiator, and the specific form is as follows:
Figure BDA0002827641690000031
in the formula:
Figure BDA0002827641690000032
to the planned desired yaw angle;
Figure BDA0002827641690000033
to the planned desired yaw angular velocity; t is the controller step length;
Figure BDA0002827641690000034
for a planned desired yaw angular acceleration;
Figure BDA0002827641690000035
the calculation formula is as follows:
Figure BDA0002827641690000036
in the formula: r is the maximum maneuvering angular acceleration of the satellite side swing shaft;
Figure BDA0002827641690000037
is a prescribed maximum maneuvering angular velocity; h is a smoothing factor; a is an intermediate calculation variable, and the specific form is as follows:
Figure BDA0002827641690000038
Figure BDA0002827641690000039
in the formula: thetavFor the final desired roll angle,
obtaining the expected yaw angle according to the plan
Figure BDA00028276416900000310
Desired angular velocity
Figure BDA00028276416900000311
And desired angular acceleration
Figure BDA00028276416900000312
Calculating the quaternion of expected attitude of the satellite body coordinate system relative to the orbit coordinate system
Figure BDA00028276416900000313
Desired angular velocity
Figure BDA00028276416900000314
And expected angular acceleration
Figure BDA00028276416900000315
Comprises the following steps:
Figure BDA00028276416900000316
Figure BDA00028276416900000317
Figure BDA00028276416900000318
step three: transforming the attitude planning result from the orbital coordinate system to an inertial coordinate system:
calculating according to the second step to obtain the quaternion of the expected attitude of the satellite body coordinate system relative to the orbit coordinate system
Figure BDA00028276416900000319
Desired angular velocity
Figure BDA00028276416900000320
And expected angular acceleration
Figure BDA00028276416900000321
Quaternion Q of expected attitude of satellite body coordinate system relative to inertial coordinate systemdDesired angular velocity ωdAnd desired angular acceleration adComprises the following steps:
Figure BDA0002827641690000041
Figure BDA0002827641690000042
Figure BDA0002827641690000043
in the formula: qoIIs the attitude quaternion of the orbit coordinate system relative to the inertia coordinate system;
step four: and establishing a satellite error kinematic kinetic equation by combining the attitude planning information:
according to the satellite attitude kinematics and the kinetic equation in the step one and by combining the attitude planning results in the step two and the step three, the error kinematics and the kinetic model of the satellite can be obtained as follows:
Figure BDA0002827641690000044
Figure BDA0002827641690000045
step five: the attitude control of the satellite is divided into an angle loop and an angular velocity loop, and control law design is carried out, wherein a controller of the satellite angle loop is as follows:
Figure BDA0002827641690000046
definition of
Figure BDA0002827641690000047
The controller of the satellite angular velocity loop is then:
Figure BDA0002827641690000048
the invention has the beneficial effects that: the quick maneuvering control method for the low-orbit optical remote sensing satellite multi-point imaging can realize quick and high-precision maneuvering control, greatly reduce maneuvering time of the satellite during orbit imaging, and improve the utilization rate of the optical remote sensing satellite. Compared with a classical control method, the method has the advantages that the method has the rapidity and the high-precision attitude stability of attitude maneuver, the satellite attitude can be converged to the planned attitude stably in the attitude maneuver process, and the control method is simple in structure, easy to implement and capable of being applied to engineering practice.
Drawings
FIG. 1 is a schematic structural diagram of a low-orbit optical remote sensing satellite multipoint imaging rapid maneuvering control system.
FIG. 2 is a diagram of an attitude planning effect in an embodiment of a low-orbit optical remote sensing satellite multi-point imaging rapid maneuver control method of the present invention.
FIG. 3 is a diagram of the effect of multipoint imaging of a satellite in an embodiment of the multipoint imaging fast maneuvering control method for the low-orbit optical remote sensing satellite of the invention.
FIG. 4 is a diagram of an attitude control tracking convergence trajectory in an embodiment of the low-orbit optical remote sensing satellite multi-point imaging rapid maneuver control method of the present invention.
FIG. 5 is a track diagram of attitude control error convergence in an embodiment of the low-orbit optical remote sensing satellite multi-point imaging fast maneuver control method of the present invention.
Table 1 shows the parameters of the attitude planning for the satellite multi-point imaging task mode, with different imaging points corresponding to the yaw angles of 15 °, -10 °, and 0 °.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and examples.
A low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method comprises the following steps:
the method comprises the following steps: performing kinematic dynamics modeling on the low-orbit optical remote sensing satellite: definition FIRepresenting an inertial coordinate system, FBRepresenting a satellite body coordinate system; satellite body coordinate system FBRelative to an inertial frame FIIs expressed as
Figure BDA0002827641690000051
Body coordinate system FBRelative inertial frame FIIs expressed as a unit quaternion
Figure BDA0002827641690000052
And satisfy the constraint condition
Figure BDA0002827641690000053
Wherein q is0Is the scalar part of a quaternion Q, Q ═ Q1,q2,q3) Is the vector portion of the quaternion Q,
Figure BDA0002827641690000054
representing an n-dimensional real vector space, the kinematic and kinetic equations of the satellite are:
Figure BDA0002827641690000055
Figure BDA0002827641690000056
in the formula:
Figure BDA0002827641690000057
the moment of inertia of the satellite is represented by a positive definite matrix; i is3Is a 3 × 3 identity matrix; u is the control moment of the reaction flywheel;
Figure BDA0002827641690000058
is the total angular momentum of the reaction flywheel;
Figure BDA0002827641690000059
as an antisymmetric matrix, for arbitrary vectors
Figure BDA00028276416900000510
Satisfy s (x) y ═ x × y, where x represents a vector cross product;
the expected attitude of the satellite is defined as the attitude direction of the coordinate system of the satellite body relative to the inertial coordinate system, and quaternion is carried out through the expected attitude
Figure BDA00028276416900000511
Represents; q. q.sd0Quaternion Q for the desired attitudedScalar part of qdQuaternion Q for the desired attitudedThe vector portion of (1); the attitude tracking error is defined as the error quaternion:
Figure BDA00028276416900000512
in the formula:
Figure BDA00028276416900000513
Figure BDA00028276416900000514
representing quaternion multiplication, qe0Is an error quaternion QeScalar part of qeIs an error quaternion QeThe vector portion of (1); the angular velocity tracking error is:
ωe＝ω-R(Qe)ωd
in the formula: omegadIs the desired angular velocity of the satellite; rotation matrix R (Q)e) The following relationsComprises the following steps:
Figure BDA0002827641690000061
Figure BDA0002827641690000062
and satisfies the constraint condition | | | R (Q)e)||＝1；
Step two: the satellite attitude planning method comprises the following steps:
the multi-point imaging task of the low-orbit optical remote sensing satellite requires the attitude to simultaneously meet the requirements of rapidity and high precision of a stable section in the maneuvering process. Compared with step control input, after the attitude planning is added, the maneuvering process can be controlled according to the maximum torque to ensure rapidity, and after maneuvering is finished, the attitude planning can enable the satellite attitude to be converged to a stable state quickly and stably.
Definition FoRepresenting an orbital coordinate system, a satellite body coordinate system FBRelative to the orbital coordinate system FoIs expressed as
Figure BDA0002827641690000063
Body coordinate system FBRelative orbit coordinate system FoUsing unit quaternion
Figure BDA0002827641690000064
Expressing and satisfying quaternion constraint conditions; the attitude of the satellite multi-point imaging task is that the satellite continuously performs side-sway maneuver; the attitude planning is to obtain a satellite body coordinate system F through a corresponding planning algorithmBRelative orbit coordinate system FoDesired attitude quaternion
Figure BDA0002827641690000065
Desired angular velocity
Figure BDA0002827641690000066
And expected angular acceleration
Figure BDA0002827641690000067
The attitude planning is realized by an improved differentiator, and the specific form is as follows:
Figure BDA0002827641690000068
in the formula:
Figure BDA0002827641690000069
to the planned desired yaw angle;
Figure BDA00028276416900000610
to the planned desired yaw angular velocity; t is the controller step length;
Figure BDA00028276416900000611
for a planned desired yaw angular acceleration;
Figure BDA00028276416900000612
the calculation formula is as follows:
Figure BDA00028276416900000613
in the formula: r is the maximum maneuvering angular acceleration of the satellite side swing shaft;
Figure BDA00028276416900000614
is a prescribed maximum maneuvering angular velocity; h is a smoothing factor; a is an intermediate calculation variable, and the specific form is as follows:
Figure BDA0002827641690000071
Figure BDA0002827641690000072
in the formula: thetavFor the final desired roll angle,
obtaining the expected yaw angle according to the plan
Figure BDA0002827641690000073
Desired angular velocity
Figure BDA0002827641690000074
And desired angular acceleration
Figure BDA0002827641690000075
Calculating the quaternion of expected attitude of the satellite body coordinate system relative to the orbit coordinate system
Figure BDA0002827641690000076
Desired angular velocity
Figure BDA0002827641690000077
And expected angular acceleration
Figure BDA0002827641690000078
Comprises the following steps:
Figure BDA0002827641690000079
Figure BDA00028276416900000710
Figure BDA00028276416900000711
step three: transforming the attitude planning result from the orbital coordinate system to an inertial coordinate system:
calculating according to the second step to obtain the quaternion of the expected attitude of the satellite body coordinate system relative to the orbit coordinate system
Figure BDA00028276416900000712
Desired angular velocity
Figure BDA00028276416900000713
And expected angular acceleration
Figure BDA00028276416900000714
Quaternion Q of expected attitude of satellite body coordinate system relative to inertial coordinate systemdDesired angular velocity ωdAnd desired angular acceleration adComprises the following steps:
Figure BDA00028276416900000715
Figure BDA00028276416900000716
Figure BDA00028276416900000717
in the formula: qoIIs the attitude quaternion of the orbit coordinate system relative to the inertia coordinate system;
step four: and establishing a satellite error kinematic kinetic equation by combining the attitude planning information:
according to the satellite attitude kinematics and the kinetic equation in the step one and by combining the attitude planning results in the step two and the step three, the error kinematics and the kinetic model of the satellite can be obtained as follows:
Figure BDA00028276416900000718
Figure BDA00028276416900000719
step five: the attitude of the satellite is divided into an inner loop and an outer loop, a controller is established, and the multi-point imaging rapid maneuvering control of the low-orbit optical remote sensing satellite is realized:
the controller is designed based on the idea of nonlinear feedback and model reference control. According to the satellite error kinematics and the kinetic equation, the kinematics equation and the kinetic equation have different time scales, so that the attitude control of the satellite is divided into an angle loop and an angular velocity loop to carry out controller design, the block diagram of the control system is shown in figure 1, the sign in the figure represents the sign polarity of a signal,
Figure BDA0002827641690000086
representing a signal summer. In the figure, a controller structure combining inner and outer loop control and nonlinear feedback control in the patent is described in a structural block diagram form, and the specific flow is as follows: after the expected attitude is subjected to attitude planning and coordinate system transformation, the generated expected attitude quaternion is used for an angle loop controller, and the expected angular velocity and the expected angular acceleration are used for a nonlinear feedback controller. The output of the angular loop controller is the virtual expected angular velocity of the angular velocity loop
Figure BDA0002827641690000081
And finally, inputting the sum u of the control quantities of the angular velocity loop controller and the nonlinear feedback controller into a flywheel to generate a control moment to act on the satellite, wherein the attitude dynamics and the attitude kinematics in the graph respectively feed back the real-time angular velocity and the attitude quaternion of the satellite.
The controller of the satellite angle loop is as follows:
Figure BDA0002827641690000082
definition of
Figure BDA0002827641690000083
The control quantity generated by the satellite angular velocity loop controller in combination with the nonlinear feedback controller is:
Figure BDA0002827641690000084
and when the satellite multi-point imaging task mode is adopted, different imaging points correspond to different side swing angles. The present embodiment performs attitude planning by respectively swinging to 15 °, -10 °, and 0 °, and ensures that each imaging task has an imaging time greater than 15 s. Table 1 shows the parameters relevant to the examples.
Figure BDA0002827641690000085
Figure BDA0002827641690000091
TABLE 1
The attitude planning curves of the yaw angle, the yaw angular velocity and the yaw angular acceleration are shown in fig. 2, and the information of the yaw angle, the yaw angular velocity and the yaw angular acceleration planned for realizing the multipoint imaging fast maneuver control is given in the diagram. The corresponding maneuvering and imaging effects of the satellite are schematically shown in fig. 3, which shows a trajectory diagram of an optical axis of an optical remote sensing satellite camera using rapid maneuvering control for three different shooting points by the satellite.
The attitude controller of the patent can guarantee stable and accurate attitude tracking control according to the attitude planning result of multi-point imaging. The attitude control tracking convergence trajectory chart is shown in fig. 4, and is a satellite yaw angle tracking convergence trajectory chart and a satellite yaw angular velocity tracking convergence trajectory chart, respectively. It can be observed from the figure that the actual roll angle curve and the actual angular velocity curve are basically completely overlapped with the planned roll angle and the planned angular velocity curve, thereby further demonstrating that the controller designed by the patent realizes good dynamic tracking on the planned roll angle and the planned roll angular velocity. The attitude control error convergence trajectory graph is shown in fig. 5, which is a satellite yaw angle error convergence trajectory graph and a satellite yaw angular velocity error convergence trajectory graph, and shows that the overall control error is small and the convergence rate is high in the maneuvering process of the satellite in multipoint imaging.

Claims (1)

1. A low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method is characterized by comprising the following steps:
the method comprises the following steps: performing kinematic dynamics modeling on the low-orbit optical remote sensing satellite: definition FIRepresenting an inertial coordinate system, FBRepresenting a satellite body coordinate system; satellite body coordinate system FBRelative to an inertial frame FIIs expressed as
Figure FDA0003512672110000011
Satellite body coordinate system FBRelative inertial frame FIIs expressed as a unit quaternion
Figure FDA0003512672110000012
And satisfy the constraint condition
Figure FDA0003512672110000013
Wherein q is0Is a scalar part of unit quaternion Q, Q ═ Q1,q2,q3) Is the vector portion of the unit quaternion Q,
Figure FDA0003512672110000014
representing an n-dimensional real vector space, the kinematic and kinetic equations of the satellite are:
Figure FDA0003512672110000015
in the formula:
Figure FDA0003512672110000016
the moment of inertia of the satellite is represented by a positive definite matrix; i is3Is a 3 × 3 identity matrix; u is the control moment of the reaction flywheel;
Figure FDA0003512672110000017
is the total angular momentum of the reaction flywheel;
Figure FDA0003512672110000018
as an antisymmetric matrix, for an arbitrary vector x,
Figure FDA0003512672110000019
satisfy s (x) y ═ x × y, where x represents a vector cross product;
the expected attitude of the satellite is defined as the attitude direction of the coordinate system of the satellite body relative to the inertial coordinate system, and quaternion is carried out through the expected attitude
Figure FDA00035126721100000110
Represents; q. q.sd0Quaternion Q for the desired attitudedScalar part of qdQuaternion Q for the desired attitudedThe vector portion of (1); the attitude tracking error is defined as the error quaternion:
Figure FDA00035126721100000111
in the formula:
Figure FDA00035126721100000112
Figure FDA00035126721100000113
representing quaternion multiplication, qe0Is an error quaternion QeScalar part of qeIs an error quaternion QeThe vector portion of (1); the angular velocity tracking error is:
ωe＝ω-R(Qe)ωd
in the formula: omegadIs the desired angular velocity of the satellite; rotation matrix R (Q)e) The following relationships exist:
Figure FDA00035126721100000114
Figure FDA00035126721100000115
and satisfies the constraint condition | | | R (Q)e)||＝1；
Step two: the satellite attitude planning method comprises the following steps: definition FoRepresenting an orbital coordinate system, a satellite body coordinate system FBRelative to the orbital coordinate system FoIs expressed as
Figure FDA00035126721100000116
Satellite body coordinate system FBRelative orbit coordinate system FoUsing unit quaternion
Figure FDA0003512672110000021
Expressing and satisfying unit quaternion constraint conditions; the attitude of the satellite multi-point imaging task is that the satellite continuously performs side-sway maneuver; obtaining a satellite body coordinate system F through attitude planningBRelative orbit coordinate system FoDesired attitude quaternion
Figure FDA0003512672110000022
Desired angular velocity
Figure FDA0003512672110000023
And expected angular acceleration
Figure FDA0003512672110000024
The attitude planning is realized by an improved differentiator, and the specific form is as follows:
Figure FDA0003512672110000025
in the formula:
Figure FDA0003512672110000026
to the planned desired yaw angle;
Figure FDA0003512672110000027
to the planned desired yaw angular velocity; t is the controller step length;
Figure FDA0003512672110000028
for a planned desired yaw angular acceleration;
Figure FDA0003512672110000029
the calculation formula is as follows:
Figure FDA00035126721100000210
in the formula: r is the maximum maneuvering angular acceleration of the satellite side swing shaft;
Figure FDA00035126721100000211
is a prescribed maximum maneuvering angular velocity; h is a smoothing factor; a is an intermediate calculation variable, and the specific form is as follows:
Figure FDA00035126721100000212
Figure FDA00035126721100000213
in the formula: thetavFor the final desired roll angle,
obtaining the expected yaw angle according to the plan
Figure FDA00035126721100000214
Desired angular velocity
Figure FDA00035126721100000215
And desired angular acceleration
Figure FDA00035126721100000216
Calculating the quaternion of expected attitude of the satellite body coordinate system relative to the orbit coordinate system
Figure FDA00035126721100000217
Desired angular velocity
Figure FDA00035126721100000218
And expected angular acceleration
Figure FDA00035126721100000219
Comprises the following steps:
Figure FDA00035126721100000220
Figure FDA00035126721100000221
Figure FDA00035126721100000222
step three: transforming the attitude planning result from the orbital coordinate system to an inertial coordinate system:
calculating according to the second step to obtain the quaternion of the expected attitude of the satellite body coordinate system relative to the orbit coordinate system
Figure FDA0003512672110000031
Desired angular velocity
Figure FDA0003512672110000032
And expected angular acceleration
Figure FDA0003512672110000033
Quaternion Q of expected attitude of satellite body coordinate system relative to inertial coordinate systemdDesired angular velocity ωdAnd desired angular acceleration adComprises the following steps:
Figure FDA0003512672110000034
Figure FDA0003512672110000035
Figure FDA0003512672110000036
in the formula: qoIIs the attitude quaternion of the orbit coordinate system relative to the inertia coordinate system;
step four: and establishing a satellite error kinematic kinetic equation by combining the attitude planning information:
according to the satellite attitude kinematics and the kinetic equation in the step one and by combining the attitude planning results in the step two and the step three, the error kinematics and the kinetic model of the satellite can be obtained as follows:
Figure FDA0003512672110000037
Figure FDA0003512672110000038
step five: the attitude control of the satellite is divided into an angle loop and an angular velocity loop, and control law design is carried out, wherein a controller of the satellite angle loop is as follows:
Figure FDA0003512672110000039
definition of
Figure FDA00035126721100000310
The controller of the satellite angular velocity loop is then:
Figure FDA00035126721100000311
wherein K1For the first control gain, K2Respectively, second control gains.
CN202011434349.8A 2020-12-10 2020-12-10 Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method Active CN112455726B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202011434349.8A CN112455726B (en) 2020-12-10 2020-12-10 Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202011434349.8A CN112455726B (en) 2020-12-10 2020-12-10 Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method

Publications (2)

Publication Number Publication Date
CN112455726A CN112455726A (en) 2021-03-09
CN112455726B true CN112455726B (en) 2022-04-19

Family

ID=74801113

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202011434349.8A Active CN112455726B (en) 2020-12-10 2020-12-10 Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method

Country Status (1)

Country Link
CN (1) CN112455726B (en)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114115305B (en) * 2021-11-01 2022-10-04 武汉大学 Control system design method of high-precision remote sensing small satellite with quick attitude maneuvering
CN113968362B (en) * 2021-11-16 2023-07-07 长光卫星技术股份有限公司 Satellite in-orbit autonomous triaxial rapid maneuvering control method
CN115649491B (en) * 2022-11-11 2024-03-12 长光卫星技术股份有限公司 Low orbit optical remote sensing satellite staring imaging control method suitable for multi-source interference

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103293957A (en) * 2013-05-22 2013-09-11 上海新跃仪表厂 Satellite attitude maneuver method for performing routing planning relative to moving coordinate system
WO2015188409A1 (en) * 2014-06-13 2015-12-17 北京航天控制仪器研究所 Method and system for controlling antenna of mobile communication application system based on double quaternions in mems inertial navigation

Family Cites Families (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0790542B1 (en) * 1995-08-11 2000-01-19 DaimlerChrysler AG Method for controlling the attitude of a three-axis stabilized, earth oriented bias momentum spacecraft
FR2826470B1 (en) * 2001-06-26 2003-09-19 Astrium Sas METHOD AND DEVICE FOR STEERING THE ATTITUDE AND GUIDANCE OF A SATELLITE BY A GYRODYNES CLUSTER
US8386098B2 (en) * 2009-08-03 2013-02-26 David A. Bailey Method for a deeper search in a time-limited image satellite planning environment
EP3140613B1 (en) * 2014-05-05 2024-04-03 Hexagon Technology Center GmbH Surveying system
CN106094854B (en) * 2016-07-15 2019-03-01 中国人民解放军战略支援部队航天工程大学 Electromagnetism formation satellite attitude and track relation control method
CN107479566A (en) * 2017-08-25 2017-12-15 中国科学院长春光学精密机械与物理研究所 Flexible satellite attitude control method based on three-stage path planning
CN109018441A (en) * 2018-09-13 2018-12-18 上海航天控制技术研究所 A kind of satellite any attitude mobile process drift angle tracking and controlling method
CN109823572B (en) * 2019-01-30 2020-09-01 武汉大学 Actuating mechanism configuration and control method for reciprocating and rapid swinging of agile satellite attitude
CN110435930B (en) * 2019-08-26 2020-08-25 长光卫星技术有限公司 Low-orbit optical satellite uniform-deceleration push-broom attitude planning method
CN110816897A (en) * 2019-10-31 2020-02-21 北京控制工程研究所 Multi-mode switching control method based on CMG system

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103293957A (en) * 2013-05-22 2013-09-11 上海新跃仪表厂 Satellite attitude maneuver method for performing routing planning relative to moving coordinate system
WO2015188409A1 (en) * 2014-06-13 2015-12-17 北京航天控制仪器研究所 Method and system for controlling antenna of mobile communication application system based on double quaternions in mems inertial navigation

Also Published As

Publication number Publication date
CN112455726A (en) 2021-03-09

Similar Documents

Publication Publication Date Title
CN112455726B (en) Low-orbit optical remote sensing satellite multi-point imaging rapid maneuvering control method
Zhang et al. Constrained coordinated path-following control for underactuated surface vessels with the disturbance rejection mechanism
CN112904728B (en) Mechanical arm sliding mode control track tracking method based on improved approach law
CN103869704B (en) Based on the robot for space star arm control method for coordinating of expansion Jacobian matrix
CN109857100B (en) Composite track tracking control algorithm based on inversion method and fast terminal sliding mode
CN111026160B (en) Trajectory tracking control method for quad-rotor unmanned aerial vehicle
CN106874617B (en) Efficient helicopter maneuvering flight quality grade evaluation method
CN115649491B (en) Low orbit optical remote sensing satellite staring imaging control method suitable for multi-source interference
CN109782759B (en) Approximate decoupling and rapid track following control method of wheeled mobile robot
Zhang et al. Adaptive fuzzy finite-time control for spacecraft formation with communication delays and changing topologies
CN108614560A (en) A kind of mobile robot visual servo guaranteed cost tracking and controlling method
CN108663936B (en) Model does not know spacecraft without unwinding Attitude Tracking finite-time control method
CN111198570B (en) Anti-delay high-precision active disturbance rejection attitude control method based on fixed time differentiator prediction
CN108519736A (en) A kind of USV track followings contragradience sliding-mode control
CN111605737A (en) Spacecraft three-phase control multi-level collaborative planning and agile maneuvering method
Gu et al. Trajectory planning and tracking control of a ground mobile robot: A reconstruction approach towards space vehicle
Kang et al. Inverse dynamics and energy optimal trajectories for a wheeled mobile robot
Alipour et al. Practical fractional-order nonsingular terminal sliding mode control of spacecraft
CN108062024A (en) A kind of mobile robot back-stepping sliding mode control method for considering resistance
Qi et al. Three-dimensional formation control based on filter backstepping method for multiple underactuated underwater vehicles
Zhang et al. A quadratic programming based neural dynamic controller and its application to UAVs for time-varying tasks
CN113867157B (en) Optimal trajectory planning method and device for control compensation and storage device
CN114594785B (en) Unmanned aerial vehicle obstacle avoidance real-time track planning method based on mixed integer second order cone optimization
CN114840947A (en) Three-degree-of-freedom mechanical arm dynamic model with constraint
CN114800521A (en) Three-degree-of-freedom mechanical arm fixed path point motion control system with constraint

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
CB02 Change of applicant information

Address after: No. 1299, Mingxi Road, Beihu science and Technology Development Zone, Changchun City, Jilin Province

Applicant after: Changguang Satellite Technology Co.,Ltd.

Address before: No.1759 Mingxi Road, Gaoxin North District, Changchun City, Jilin Province

Applicant before: CHANG GUANG SATELLITE TECHNOLOGY Co.,Ltd.

CB02 Change of applicant information
GR01 Patent grant
GR01 Patent grant
PE01 Entry into force of the registration of the contract for pledge of patent right

Denomination of invention: A Fast Maneuver Control Method for Multi point Imaging of Low Earth Orbit Optical Remote Sensing Satellites

Granted publication date: 20220419

Pledgee: Jilin Shuangyang Rural Commercial Bank Co.,Ltd. Ecological Square sub branch

Pledgor: Changguang Satellite Technology Co.,Ltd.

Registration number: Y2024220000052

PE01 Entry into force of the registration of the contract for pledge of patent right