CN103708044B - A kind of saturated sliding mode variable structure control method for satellite rapid attitude maneuver - Google Patents
A kind of saturated sliding mode variable structure control method for satellite rapid attitude maneuver Download PDFInfo
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Abstract
The invention discloses a kind of saturated sliding mode variable structure control method for satellite rapid attitude maneuver, step 1: the current pose obtaining satellite, and draw the attitude error of current pose relative to targeted attitude of satellite; Step 2: the fundamental frequency determining bandwidth, damping coefficient, actuating unit maximum torque, torque adjusting coefficient, flexible appendage, calculates the controling parameters of satellite rapid attitude maneuver; Step 3: the saturated sliding-mode surface calculating satellite rotating direction, calculates pitching and yaw direction linear sliding mode face; Step 4: calculate control torque; Step 5: the control command being obtained corresponding actuating unit by the control torque described in step 4, and instruction is sent to corresponding actuating unit, control this actuating unit and fast reserve is carried out to satellite attitude.Method of the present invention is simple and reliable, and operand is little, can realize fast reserve and the fast and stable of satellite, make the time shorten of attitude maneuver, with the obvious advantage with other Measures compare.
Description
Technical Field
The invention relates to a satellite attitude control method, in particular to a saturated sliding mode variable structure control method for satellite rapid attitude maneuver.
Background
With the improvement of the requirements of satellite users, in order to shorten the revisit period of many satellites, the attitude control system is required to not only realize high-precision and high-stability control, but also have the capability of fast attitude maneuver. In the process of satellite attitude maneuver, the acceleration, uniform speed and deceleration sections of the satellite basically depend on the capability of an actuating mechanism and have little relation with an algorithm; however, the rapidity of convergence stability in the rapid stabilization phase completely depends on the algorithm, and in the prior art, the algorithm is not reliable, and the calculation amount is too large, so that the attitude maneuver speed of the satellite is slow and the reliability is not high.
Disclosure of Invention
The invention aims to provide a saturated sliding mode variable structure control method for satellite rapid attitude maneuver, which is simple and reliable, has small operand, can realize rapid maneuver and rapid stability of a satellite, shortens the time of attitude maneuver, and has obvious advantages compared with other methods.
In order to achieve the purpose, the invention is realized by the following technical scheme:
a saturated sliding mode variable structure control method for satellite rapid attitude maneuver comprises the following steps:
step 1: acquiring the current attitude of the satellite through attitude measurement, and obtaining the attitude error of the current attitude of the satellite relative to the target attitude;
step 2: determining the bandwidth, the damping coefficient, the maximum moment of an actuating mechanism, the moment adjusting coefficient and the fundamental frequency of a flexible accessory, and calculating the control parameters of the rapid attitude maneuver of the satellite;
and step 3: calculating a saturated sliding mode surface in the rolling direction of the satellite, and calculating linear sliding mode surfaces in the pitching and yawing directions;
and 4, step 4: calculating to obtain a control moment according to the calculation results of the steps 1-3;
and 5: and 4, obtaining a control instruction of the corresponding executing mechanism by the control moment in the step 4, sending the instruction to the corresponding executing mechanism, and controlling the executing mechanism to quickly maneuver the satellite attitude.
The substeps of step 1 are:
step 1.1: measuring satellite inertial angular velocity information through the gyroscope combination, and measuring and calculating through the star sensor to obtain an attitude quaternion of the satellite relative to an orbital system;
step 1.2: and (4) calculating the attitude angle error and the attitude angular speed error of the current attitude of the satellite relative to the target attitude according to the information measured in the step 1.1.
The control parameters of step 2 include: boundary layer thickness and sliding mode face coefficient.
The bandwidth in step 2 is 1/10 of the fundamental frequency of the flexible accessory.
The flexible accessory is a solar sailboard of a satellite.
The method for calculating the control torque in the step 4 comprises the following steps: and calculating to obtain the control moment by adopting a method of replacing a sign function with a saturation function according to comparison between the attitude angular velocity error of the satellite and a threshold value and comparison between the sliding mode surface size and the boundary layer size.
The actuating mechanism in the step 5 is a flywheel or a single-frame control moment gyro group.
Compared with the prior art, the invention has the following advantages:
the method is simple and reliable, has small calculation amount, can realize the rapid maneuver and rapid stability of the satellite, shortens the maneuver time of the attitude, and has obvious advantages compared with other methods.
Drawings
Fig. 1 is a flowchart of a saturated sliding mode variable structure control method for satellite rapid attitude maneuver according to the present invention.
Detailed Description
The present invention will now be further described by way of the following detailed description of a preferred embodiment thereof, taken in conjunction with the accompanying drawings.
As shown in fig. 1, a method for controlling a saturated sliding mode variable structure for a satellite fast attitude maneuver includes the following steps:
step 1: acquiring the current attitude of the satellite through attitude measurement, and obtaining the attitude error of the current attitude of the satellite relative to the target attitude; wherein:
step 1.1: and measuring satellite inertial angular velocity information through the gyroscope combination, and measuring and calculating through the star sensor to obtain an attitude quaternion of the satellite relative to the orbital system.
Attitude angle error of satellite relative to target attitude:whereinthe attitude quaternion of the star body relative to the orbital system is obtained through measurement and calculation of a star sensor;is the desired target quaternion.
Step 1.2: and (4) calculating the attitude angle error and the attitude angular speed error of the current attitude of the satellite relative to the target attitude according to the information measured in the step 1.1.
Will be provided withConversion to an attitude matrix，Whereinis thatThe vector portion of (a) is,is thatIs determined by the skew-symmetric matrix of (a),
according to 1-2-3 order, fromAnd (3) calculating an attitude angle error:
wherein,is a matrixRow i and column j.
Attitude angular velocity error of the satellite relative to the target attitude:
wherein,is the star inertia angular velocity measured by the gyro combination,is thatA corresponding matrix of the attitude is formed,is the real-time track angular velocity,and is the desired target attitude angular velocity.
Step 2: determining the bandwidth and the damping coefficient, the maximum torque of the actuating mechanism, the torque adjusting coefficient and the fundamental frequency of the flexible accessory, and calculating the control parameters of the rapid attitude maneuver of the satellite. In this embodiment, since the solar sailboard has the greatest influence on the fast attitude maneuver of the satellite, the solar sailboard of the satellite is selected as the flexible attachment, and the fundamental frequency of the solar sailboard of the satellite is set asSelecting the system bandwidth asSelecting damping of the systemThe maximum control moment in three directions can be determined according to the capacity of the actuator。
Coefficient of slip form surface:
boundary layer thickness:
wherein,the moment adjustment coefficient is 0.1-0.5. Get、、Is the three-axis direction moment of inertia of the satellite.
And step 3: calculating a saturated sliding mode surface in the rolling direction of the satellite, and calculating linear sliding mode surfaces in the pitching and yawing directions;
the YZ direction slip form surface is:
the X-direction slip form surface is as follows:
wherein,
is the maximum angular velocity of the attitude maneuver, the maximum angular momentum in the rolling direction that can be determined by the executing maneuverAnd calculating to obtain:
is a bias threshold of inertia, and can be 0.1-0.3 generally.
Step 4, according to the calculation results of the steps 1-3, comparing the attitude angular velocity error of the satellite with a threshold value, comparing the size of the sliding mode surface with the size of the boundary layer, and calculating by adopting a method of replacing a sign function with a saturation function to obtain a control moment, wherein the threshold value is a fixed value 1 × 10-3°/s。
When in use，Otherwise
When in useWhen the temperature of the water is higher than the set temperature,
when in useWhen the temperature of the water is higher than the set temperature,
wherein,。
and 5: and 4, obtaining a control instruction of the corresponding executing mechanism by the control moment in the step 4, sending the instruction to the corresponding executing mechanism, and controlling the executing mechanism to quickly maneuver the satellite attitude. In the embodiment, the actuating mechanism is a flywheel or a single-frame control moment gyro group. If the flywheel is used as an actuating mechanism, calculating a rotating speed instruction of the flywheel and sending the rotating speed instruction to the flywheel; if the single-frame control moment gyro group is used as an actuating mechanism, a rotating speed instruction of an outer frame of the control moment gyro is calculated and sent to the control moment gyro.
In conclusion, the saturated sliding mode variable structure control method for the satellite rapid attitude maneuver is simple and reliable, has small calculation amount, can realize the rapid maneuver and the rapid stability of the satellite, shortens the time of the attitude maneuver, and has obvious advantages compared with other methods.
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.
Claims (7)
1. A saturated sliding mode variable structure control method for satellite rapid attitude maneuver is characterized by comprising the following steps:
step 1: acquiring the current attitude of the satellite through attitude measurement, and obtaining the attitude error of the current attitude of the satellite relative to the target attitude;
step 2: determining the bandwidth, the damping coefficient, the maximum moment of an actuating mechanism, the moment adjusting coefficient and the fundamental frequency of a flexible accessory, and calculating the control parameters of the rapid attitude maneuver of the satellite;
and step 3: calculating a saturated sliding mode surface in the rolling direction of the satellite, and calculating linear sliding mode surfaces in the pitching and yawing directions;
and 4, step 4: calculating to obtain a control moment according to the calculation results of the steps 1-3;
and 5: and 4, obtaining a control instruction of the corresponding executing mechanism by the control moment in the step 4, sending the instruction to the corresponding executing mechanism, and controlling the executing mechanism to quickly maneuver the satellite attitude.
2. The saturated sliding mode variable structure control method for the satellite fast attitude maneuver according to claim 1, wherein the substep of step 1 is:
step 1.1: measuring satellite inertial angular velocity information through the gyroscope combination, and measuring and calculating through the star sensor to obtain an attitude quaternion of the satellite relative to an orbital system;
step 1.2: and (4) calculating the attitude angle error and the attitude angular speed error of the current attitude of the satellite relative to the target attitude according to the information measured in the step 1.1.
3. The method according to claim 2, wherein the control parameters of step 2 comprise: boundary layer thickness and sliding mode face coefficient.
4. The method for controlling the saturated sliding mode variable structure of the satellite fast attitude maneuver according to claim 2, wherein the bandwidth in step 2 is 1/10 of the fundamental frequency of the flexible attachment.
5. The saturated slip mode variable structure control method for the satellite fast attitude maneuver according to claim 4, wherein the flexible attachment is a solar windsurfing board of the satellite.
6. The method for controlling the saturated sliding mode variable structure of the satellite rapid attitude maneuver according to claim 3, wherein the method for calculating the control torque in the step 5 comprises the following steps: and calculating to obtain the control moment by adopting a method of replacing a sign function with a saturation function according to comparison between the attitude angular velocity error of the satellite and a threshold value and comparison between the sliding mode surface size and the boundary layer size.
7. The method as claimed in claim 1 or 6, wherein the actuator in step 5 is a flywheel or a single-frame control moment gyro group.
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| CN104090489B (en) * | 2014-07-02 | 2016-12-07 | 中国科学院长春光学精密机械与物理研究所 | A kind of flexible agile satellite attitude maneuvers rolling optimization control method |
| CN105807778A (en) * | 2014-12-31 | 2016-07-27 | 上海新跃仪表厂 | Sliding mode variable control method for controlling attitude of flexible micro-satellite |
| CN104898683B (en) * | 2015-05-20 | 2017-12-08 | 哈尔滨工业大学 | A kind of flexible satellite neutral net contragradience Sliding Mode Attitude control method |
| CN105180946B (en) * | 2015-09-02 | 2019-01-01 | 上海新跃仪表厂 | Satellite high-precision attitude determination method and system based on wideband measurement |
| CN105438499B (en) * | 2015-11-17 | 2017-06-06 | 上海新跃仪表厂 | Around the drift angle tracking and controlling method of spatial axes |
| CN106275508B (en) * | 2016-08-15 | 2019-03-01 | 上海航天控制技术研究所 | A kind of shortest path attitude maneuver control method of satellite around spatial axes |
| CN106379560B (en) * | 2016-08-30 | 2018-12-11 | 上海航天控制技术研究所 | Gas puff Z-pinch method based on quaternary number information |
| CN106379558B (en) * | 2016-09-09 | 2018-09-11 | 上海航天控制技术研究所 | A kind of sliding moding structure composite control method based on angular acceleration feedforward |
| CN109774977B (en) * | 2019-03-28 | 2021-05-07 | 上海微小卫星工程中心 | Quaternion-based time-optimal satellite attitude rapid maneuvering method |
| 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 |
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Effective date of registration: 20170623 Address after: 200233 Xuhui District, Yishan Road, No. 710, Patentee after: SHANGHAI AEROSPACE CONTROL TECHNOLOGY RESEARCH INSTITUTE Address before: 200233 Xuhui District, Yishan Road, No. 710, Patentee before: Shanghai Xinyue Instrument Factory |
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