CN110803304B - Satellite attitude control system - Google Patents
Satellite attitude control system Download PDFInfo
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- CN110803304B CN110803304B CN201910999360.XA CN201910999360A CN110803304B CN 110803304 B CN110803304 B CN 110803304B CN 201910999360 A CN201910999360 A CN 201910999360A CN 110803304 B CN110803304 B CN 110803304B
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- 238000005096 rolling process Methods 0.000 claims abstract description 7
- 238000013016 damping Methods 0.000 claims abstract description 6
- 238000009987 spinning Methods 0.000 claims abstract description 6
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- 238000013461 design Methods 0.000 description 2
- 238000005259 measurement Methods 0.000 description 2
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- 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
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- 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
- B64G1/245—Attitude control algorithms for spacecraft attitude control
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- Radar, Positioning & Navigation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The application relates to a satellite attitude control system which comprises a three-axis magnetometer and a three-axis magnetic torquer. The application also relates to a method for magnetically measuring the attitude of the magnetic control satellite, which comprises the following steps: estimating satellite inertial angular velocity by using the magnetic field intensity and the change rate thereof measured by the magnetometer; taking the estimated value of the inertial angular velocity of the satellite as input, and controlling the satellite to spin around a pitch axis by using a magnetic torquer arranged in the rolling and yawing directions to obtain the change rate of the magnetic field intensity in the pitch axis direction; and damping the non-spinning axis angular velocity by using the magnetic torquer installed in the pitch direction, with the change rate of the magnetic field strength in the pitch direction as an input.
Description
Technical Field
The application relates to the technical field of spaceflight, in particular to a satellite attitude control system.
Background
In the existing satellite attitude control technology, a plurality of attitude sensors such as a sun sensor, a star sensor, a gyroscope, a magnetometer and the like are often configured at the same time for determining the attitude of the satellite and used as the input of a controller, meanwhile, a thruster or a flywheel is mostly adopted as a main execution mechanism, and a magnetic torquer is mostly used as an auxiliary execution mechanism and used for angular momentum unloading, so that the configuration of the multi-sensor multi-execution mechanism is easy to cause the cost of the attitude control system to be overhigh.
The magnetic torquer generates magnetic moment after being electrified, the magnetic torquer interacts with a geomagnetic field to generate moment to control the attitude of the satellite, and the magnetic torquer is fixedly installed, has no vibration and high reliability, and is a common actuating mechanism for satellite attitude control. The magnetometer measures the local magnetic field intensity and is matched with the magnetic torquer to form a low-cost and high-reliability satellite attitude control system. At present, research on a magnetic measurement and control satellite attitude control system only adopting a magnetometer and a magnetic torquer is less. Under the condition of only configuring a magnetometer and a magnetic torquer, the design of a satellite attitude control scheme for realizing a satellite attitude control task has important engineering practical significance. However, at present, magnetometers and magnetic torquers are used for bias momentum satellite nutation damping and precession control and momentum wheel angular momentum unloading, and few researches are made on a minimum mode attitude control system only composed of the magnetometers and the magnetic torquers.
Therefore, there is a need in the art to develop a novel, simple and low-cost method for controlling the attitude of an omni-directional antenna communication satellite operating in a sun-synchronized morning and evening orbit.
Disclosure of Invention
The present application is directed to a satellite attitude control system.
The application also aims to provide a method for magnetically measuring the attitude of the magnetic control satellite.
In order to achieve the above object, the present application provides the following technical solutions.
In a first aspect, the present application provides a satellite attitude control system comprising a three-axis magnetometer and a three-axis magnetotorquer.
In another aspect, the present application provides a method for magnetically measuring an attitude of a magnetically controlled satellite, comprising the steps of:
(1) estimating satellite inertial angular velocity by using the magnetic field intensity and the change rate thereof measured by the magnetometer;
(2) taking the estimated value of the inertial angular velocity of the satellite as input, and controlling the satellite to spin around a pitch axis by using a magnetic torquer arranged in the rolling and yawing directions; and
(3) and damping the non-spinning shaft angular velocity by using the magnetic torquer arranged in the pitching direction and taking the change rate of the magnetic field strength in the pitching direction as input.
Compared with the prior art, the satellite attitude control method has the advantages of being novel, simple and convenient and low in cost.
Drawings
FIG. 1 shows the magnetic field strength B of the present applicationbAnd (5) decomposing schematic diagrams under a satellite body coordinate system.
Detailed Description
The technical solution of the present application will be clearly and completely described below with reference to the accompanying drawings and the embodiments of the present application.
In one aspect of the present application, there is provided a satellite attitude control system consisting of only a three-axis magnetometer and a three-axis magnetotorquer; the three-axis magnetometer is used for measuring the magnetic field intensity under the satellite body coordinate system, and obtaining the magnetic field intensity change rate under the satellite body coordinate system through difference to be used as the input of magnetic control; the three-axis magnetic torquer provides magnetic control magnetic moment, and the sun orientation is realized when the satellite runs on a sun synchronous morning and evening orbit.
In another aspect of the application, the magnetic measurement and control satellite attitude control method is provided, which only utilizes the design technical scheme of the magnetometer and the magnetic torquer, and realizes the control of the satellite attitude. The scheme can realize the directional supply of satellite energy to the sun when the satellite runs on the sun synchronous morning and evening orbit. The system comprises the following steps:
(1) estimating satellite inertial angular velocity by using the magnetic field intensity and the change rate thereof measured by the magnetometer;
(2) taking the estimated value of the inertial angular velocity of the satellite as input, and controlling the satellite to spin around a pitch axis (the pitch axis is the maximum or minimum inertia axis) by using a magnetic torquer arranged in the rolling and yawing directions to obtain the change rate of the magnetic field intensity in the pitch axis direction;
(3) and damping the non-spinning shaft angular velocity by using the magnetic torquer arranged in the pitching direction and taking the change rate of the magnetic field strength in the pitching direction as input.
Specifically, in the step (1), the magnetic field strength and the change rate thereof are measured by a magnetometer, and the satellite inertial angular velocity is estimated by using the magnetic field strength and the change rate. Magnetic field intensity BbThe decomposition under the satellite body coordinate system is shown in fig. 1.
(i) Angular velocity estimation
When the satellite is around the pitch axis YbWhile spinning, neglecting the rotation around the rolling axis XbAnd a yaw axis ZbThe angular velocity of (2), then the angular velocity can be estimated
Wherein, ω isbiIs the angular velocity vector of the system of the satellite relative to the inertial system; omegabix，ωbiy，ωbizAre each omegabiUnder the satellite system Xb，Yb，ZbA component of direction; BETA (BETA)bIs the magnetic field intensity vector under the satellite system, Bx，By，BzIs respectively BETAbUnder the satellite system Xb，Yb，ZbA component of direction;
are respectively Bx，By，BzThe rate of change of (c); alpha is the magnetic field strength BbIn the satellite body system XbOZbProjection of plane and ZbThe included angle of the axes;
is the rate of change of alpha. The superscript T denotes transpose.
(ii) Law of spin control around pitch axis
Tcy＝k*(ωbic-ωbiy)
Wherein, ω isbicTo expect the spin angular velocity, TcyA control moment is expected for the satellite pitch axis; mcxA desired magnetic moment generated for a magnetic torquer installed in a rolling direction; mczThe desired magnetic moment for the magnetic torquer mounted in the yaw direction. Denotes multiplication. Upper label2Indicating squaring. k is an angular velocity gain coefficient, and k > 0.
(iii) Damping control law of angular velocity of non-spinning shaft
Wherein M iscySign (·) represents the sign of a calculation variable for a desired magnetic moment generated by a magnetic torquer attached to the satellite in the pitch direction, and when (·) is positive, sign (·) is 1, and when (·) is negative, sign (·) is-1.
The embodiments described above are intended to facilitate the understanding and appreciation of the application by those skilled in the art. It will be readily apparent to those skilled in the art that various modifications to these embodiments may be made, and the generic principles described herein may be applied to other embodiments without the use of the inventive faculty. Therefore, the present application is not limited to the embodiments herein, and those skilled in the art who have the benefit of this disclosure will appreciate that many modifications and variations are possible within the scope of the present application without departing from the scope and spirit of the present application.
Claims (1)
1. The utility model provides a satellite attitude control system, its characterized in that, the system comprises a triaxial magnetometer and a triaxial magnetic torquer, the triaxial magnetometer measures the magnetic field intensity under the satellite body coordinate system to obtain the magnetic field intensity rate of change under the satellite body coordinate system through the difference, with as the input of magnetic control and triaxial magnetic torquer provides the magnetic control magnetic moment, realizes orienting to the sun when the satellite moves in the synchronous morning and evening orbit of sun, wherein:
(1) estimating satellite inertial angular velocity by using the magnetic field intensity and the change rate thereof measured by the magnetometer;
(2) taking the estimated value of the inertial angular velocity of the satellite as input, and controlling the satellite to spin around a pitch axis by using a magnetic torquer arranged in the rolling and yawing directions to obtain the change rate of the magnetic field intensity in the pitch axis direction; and
(3) the change rate of the magnetic field intensity in the pitching axis direction is used as input, the magnetic torquer arranged in the pitching direction is used for damping the non-spinning axis angular velocity,
the calculation formula of the satellite inertial angular velocity estimation value is as follows:
wherein, ω isbiIs the angular velocity vector of the system of the satellite relative to the inertial system; omegabix，ωbiy，ωbizAre each omegabiUnder the satellite system Xb，Yb，ZbA component of direction; b isbIs the magnetic field intensity vector under the satellite system, Bx，By，BzAre respectively BbUnder the satellite system Xb，Yb，ZbA component of direction;
are respectively Bx，By，BzThe rate of change of (c); alpha is the magnetic field strength BbIn the satellite body system XbOZbProjection of plane and ZbThe included angle of the axes;
is the rate of change of α; superscript T denotes transpose;
the calculation formula for controlling the satellite to spin around the pitch axis through the magnetic torquer is as follows:
Tcy＝k*(ωbic-ωbiy)
wherein, ω isbicTo expect the spin angular velocity, TcyA control moment is expected for the satellite pitch axis; mcxA desired magnetic moment generated for a magnetic torquer installed in a rolling direction; mczA desired magnetic moment for a magnetic torquer mounted in a yaw direction; denotes multiplication; upper label2Expressing to square; k is an angular velocity gain coefficient, and k is more than 0;
wherein the non-spin axis angular velocity is calculated as follows:
wherein M iscySign (·) represents the sign of a calculation variable for a desired magnetic moment generated by a magnetic torquer attached to the satellite in the pitch direction, and when (·) is positive, sign (·) is 1, and when (·) is negative, sign (·) is-1.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
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| CN201910999360.XA CN110803304B (en) | 2018-05-02 | 2018-05-02 | Satellite attitude control system |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201810409800.7A CN108583938B (en) | 2018-05-02 | 2018-05-02 | A kind of omnidirectional antenna telecommunication satellite attitude control system and its method that can be applied to run on sun synchronization morning and evening track |
| CN201910999360.XA CN110803304B (en) | 2018-05-02 | 2018-05-02 | Satellite attitude control system |
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| CN201810409800.7A Division CN108583938B (en) | 2018-05-02 | 2018-05-02 | A kind of omnidirectional antenna telecommunication satellite attitude control system and its method that can be applied to run on sun synchronization morning and evening track |
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| CN110803304A CN110803304A (en) | 2020-02-18 |
| CN110803304B true CN110803304B (en) | 2021-08-10 |
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| CN201910999360.XA Expired - Fee Related CN110803304B (en) | 2018-05-02 | 2018-05-02 | Satellite attitude control system |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN113184222B (en) * | 2021-05-13 | 2022-11-15 | 上海卫星工程研究所 | Magnetic torquer signal processing method and system of satellite attitude and orbit control comprehensive test equipment |
| CN113353292B (en) * | 2021-06-26 | 2022-06-07 | 山东航天电子技术研究所 | Magnetic control non-spinning sun-facing orientation method |
| CN115687847B (en) * | 2022-10-11 | 2023-04-18 | 中国人民解放军63921部队 | Common-scan sensing method for GEO space debris by low-orbit observation platform |
| CN115817856B (en) * | 2022-11-21 | 2023-06-20 | 清华大学 | Method and device for controlling stable attitude of satellite to solar spin based on pure magnetic control mode |
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| CN101554926A (en) * | 2009-05-20 | 2009-10-14 | 上海微小卫星工程中心 | Attitude control system for space vehicle and method thereof |
| CN105966639A (en) * | 2016-05-11 | 2016-09-28 | 上海微小卫星工程中心 | Stable control system and method for satellite spinning around sun |
| CN107444675A (en) * | 2017-08-18 | 2017-12-08 | 上海微小卫星工程中心 | A kind of spacecraft rate damping control method |
| US9914551B2 (en) * | 2014-06-26 | 2018-03-13 | The Boeing Company | Passive timing of asynchronous IMU attitude data |
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| CN101934863B (en) * | 2010-09-29 | 2013-04-03 | 哈尔滨工业大学 | Satellite posture all-round controlling method based on magnetic moment device and flywheel |
| FR2994287B1 (en) * | 2012-08-03 | 2016-11-04 | Thales Sa | DEVICE AND METHOD FOR SATELLITE DESORBITATION |
| CN105676671B (en) * | 2014-11-21 | 2018-05-15 | 上海新跃仪表厂 | A kind of semi-physical simulation test system of Direct to the sun control |
| CN105667838B (en) * | 2016-03-14 | 2017-08-11 | 西北工业大学 | A kind of modularization attitude determination and control devices and methods therefor of skin Nano satellite |
| CN107054697B (en) * | 2017-03-09 | 2019-03-12 | 西北工业大学 | A kind of Nano satellite magnetic torquer space temperature compensation attitude control method |
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Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN101554926A (en) * | 2009-05-20 | 2009-10-14 | 上海微小卫星工程中心 | Attitude control system for space vehicle and method thereof |
| US9914551B2 (en) * | 2014-06-26 | 2018-03-13 | The Boeing Company | Passive timing of asynchronous IMU attitude data |
| CN105966639A (en) * | 2016-05-11 | 2016-09-28 | 上海微小卫星工程中心 | Stable control system and method for satellite spinning around sun |
| CN107444675A (en) * | 2017-08-18 | 2017-12-08 | 上海微小卫星工程中心 | A kind of spacecraft rate damping control method |
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| CN110803304A (en) | 2020-02-18 |
| CN108583938A (en) | 2018-09-28 |
| CN108583938B (en) | 2019-09-17 |
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