CN111123961A - Constraint analysis-based double-vector included angle limit range determining method and system - Google Patents
Constraint analysis-based double-vector included angle limit range determining method and system Download PDFInfo
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Abstract
The invention provides a method and a system for determining a double-vector included angle limiting range based on constraint analysis. Judging whether the attitude singularity phenomenon of attitude turnover can occur or not according to the three-axis attitude change rule determined by the double vectors; determining errors by analyzing attitude references according to the three-axis attitude determined by the double vectors, and judging whether load pointing errors of an antenna, a camera and the like meet on-orbit requirements or not; and judging whether the angular momentum and the moment meet the requirements or not by analyzing the maneuvering capacity of the attitude according to the three-axis attitude determined by the double vectors. And (4) integrating the attitude singular constraint determined by the double vectors, the attitude reference determination error constraint and the attitude mobility constraint, and setting the double-vector included angle limiting range. The method solves the problem of determining the double-vector included angle limit range based on constraint analysis, and is simple and easy to implement in engineering.
Description
Technical Field
The invention relates to the field of spacecraft attitude control, in particular to a constraint analysis-based method and system for determining a double-vector included angle limit range.
Background
The attitude control mode of the spacecraft has various modes, and a three-axis stable attitude control mode is commonly used. The method is used for limiting the pointing direction of two shafts, determining the three-shaft attitude of the spacecraft by utilizing a right-hand rule and ensuring the three-shaft attitude of the spacecraft to be stable.
Any double vectors in the space can be used as a reference for determining the three-axis attitude of the spacecraft, but are influenced by factors such as energy, measurement and control and the like, and the geocentric vector, the solar vector and the orbital plane negative normal vector are one of the commonly used double vectors. If the geocentric vector and the orbital plane negative normal vector are used as double vectors, the included angle of the double vectors is fixed to 90 degrees, and the analysis of the spacecraft attitude singularity condition is simple; if the earth center vector and the sun vector are used as double vectors, the included angle of the double vectors is not fixed but changes from 0 to 360 degrees due to the revolution of the earth and the orbital motion of a spacecraft, and the attitude singularity phenomenon is likely to occur.
On one hand, the dual-vector attitude reference determination influences the calculation precision of the load orientation of an antenna, a camera and the like. The accuracy of the attitude reference is determined to have a direct relationship with the included angle of the dual vectors. When the included angle of the double vectors is smaller, the influence of the double-vector determination precision on the attitude reference determination precision is larger.
On the other hand, along with the orbital motion of the spacecraft, a certain vector changes quickly, so that the determined attitude reference changes quickly, and the spacecraft is required to be capable of real-time attitude maneuver tracking the changed expected attitude. However, due to the influences of rocket launching capability, spacecraft size and the like, a spacecraft actuating mechanism is not infinitely configured but has an attitude mobility envelope, and if the attitude change exceeds the range of the attitude mobility envelope of the spacecraft, the spacecraft can track the three-axis attitude direction which cannot be determined by double vectors.
The method has the advantages that the problem that the attitude singularity phenomenon of the spacecraft, the influence of the double-vector included angle on the attitude reference determination error and the on-orbit maneuvering capability of the spacecraft have to be faced, in order to ensure the on-orbit safe operation and task completion of the spacecraft, the constraints are analyzed, the requirement for the double-vector included angle is provided, and therefore relevant measures are made according to the requirement, and the on-orbit safe and reliable performance of the spacecraft is ensured.
The invention patent 'big ellipse orbit attitude reference determination method' (publication number: CN106767811A) proposes to adopt geocentric vector and sun vector as double vectors, but does not relate to the double vector included angle range based on constraint analysis.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a method and a system for determining a double-vector included angle limit range based on constraint analysis.
The invention provides a method for determining a double-vector included angle limit range based on constraint analysis, which comprises the following steps:
attitude singular constraint analysis: with a chosen spatial dual vector r1(t) and r2(t) as an attitude reference, determining a three-axis attitude criterion X according to the set double vectors, and analyzing the change rule of the three-axis attitude of the spacecraft along with the time t
Determining whether an attitude singularity phenomenon that the attitude turns over at plus or minus 180 degrees occurs;
and (3) attitude reference determination error constraint analysis: determining a triaxial attitude criterion X according to the double vectors, analyzing the determined error of the double vectors, analyzing the relation between the error magnification factor and the included angle of the double vectors, and determining the constraint gamma of the determined error of the attitude reference to the included angle of the double vectors1；
And (3) attitude mobility constraint analysis: determining a constraint Γ of angular momentum to dual-vector angle20Determining the constraint gamma of the moment to the dual vector angle21Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
And (3) setting a double-vector included angle limiting range: determining constraint gamma of error to double-vector included angle according to attitude singular constraint and attitude reference1Constraint gamma of dual vector angle with attitude maneuver capability2And setting a double-vector included angle limiting range gamma.
Preferably, in the attitude reference determination error constraint analysis step, the constraint Γ for the dual vector included angle of the attitude reference determination error is determined1And meanwhile, the requirement of spacecraft load pointing error including an antenna and a camera is considered.
Preferably, the attitude mobility constraint analyzing step includes:
estimating angular momentum H (t) required by a three-axis expected attitude determined by double vectors on spacecraft tracking by using the rigid moment theorem, determining constraint gamma of the angular momentum to the included angle of the double vectors according to whether the configuration of a spacecraft attitude maneuvering executing mechanism meets the requirement of the angular momentum or not20；
Estimating the moment M (t) required by the triaxial expected attitude determined by double vectors on spacecraft tracking according to the law of moment of rigid momentum, determining the constraint gamma of the moment to the double-vector included angle according to whether the configuration of a spacecraft attitude maneuvering executing mechanism meets the requirement of the moment or not21；
Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
Preferably, the dual vectors include a geocentric vector and a sun vector.
Preferably, the double-vector determination of the three-axis attitude criterion χ comprises: the X-axis is determined by the right hand rule according to the Y, Z axis, and when the centroid vector and the sun vector are parallel, the yaw attitude is flipped by 180 °.
The invention provides a constraint analysis-based dual-vector included angle limit range determining system, which comprises:
the attitude singular constraint analysis module: with a chosen spatial dual vector r1(t) andr2(t) as an attitude reference, determining a three-axis attitude criterion χ according to the set double vectors, and analyzing the change rule of the three-axis attitude of the spacecraft along with the time t
Determining whether an attitude singularity phenomenon that the attitude turns over at plus or minus 180 degrees occurs;
an attitude reference determination error constraint analysis module: determining a triaxial attitude criterion chi according to the double vectors, analyzing the determined error of the double vectors, analyzing the relation between the error magnification factor and the included angle of the double vectors, and determining the constraint gamma of the determined error of the attitude reference to the included angle of the double vectors1；
An attitude mobility constraint analysis module: determining a constraint Γ of angular momentum to dual-vector angle20Determining the constraint gamma of the moment to the dual vector angle21Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
The double-vector included angle limiting range setting module: determining constraint gamma of error to double-vector included angle according to attitude singular constraint and attitude reference1Constraint gamma of dual vector angle with attitude maneuver capability2And setting a double-vector included angle limiting range gamma.
Preferably, in the attitude reference determination error constraint analysis module, the constraint Γ for determining the dual vector angle of the attitude reference determination error is determined1And meanwhile, the requirement of spacecraft load pointing error including an antenna and a camera is considered.
Preferably, the attitude mobility constraint analysis module includes:
estimating angular momentum H (t) required by a three-axis expected attitude determined by double vectors on spacecraft tracking by using the rigid moment theorem, determining constraint gamma of the angular momentum to the included angle of the double vectors according to whether the configuration of a spacecraft attitude maneuvering executing mechanism meets the requirement of the angular momentum or not20；
According to the theorem of moment of rigid body momentum, estimating the moment M (t) required by the three-axis expected attitude determined by double vectors on the spacecraft tracking, wherein the maneuvering executing mechanism is configured according to the spacecraft attitudeIf the torque requirement is not met, determining the constraint gamma of the torque to the dual-vector included angle21；
Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
Preferably, the dual vectors include a geocentric vector and a sun vector.
Preferably, the double-vector determination of the three-axis attitude criterion χ comprises: the X-axis is determined by the right hand rule according to the Y, Z axis, and when the centroid vector and the sun vector are parallel, the yaw attitude is flipped by 180 °.
Compared with the prior art, the invention has the following beneficial effects:
the method solves the problem of determining the double-vector included angle limiting range based on constraint analysis, supports and solves the problem of limiting the double-vector included angle by the orbit and configuration of the spacecraft in the design of the spacecraft in-orbit attitude control method, and is simple and easy to realize in engineering.
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Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a flow chart of the present invention;
FIG. 2 is a diagram of the change law of the yaw attitude of the spacecraft with the geocentric vector and the solar vector as double vectors in a certain orbit;
FIG. 3 is a diagram showing the relationship between the attitude error magnification factor and the included angle between the geocentric vector and the solar vector.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
An embodiment of a method for determining a dual vector angle limit range based on constraint analysis according to the present invention is described below with reference to fig. 1, fig. 2, and fig. 3. The content includes four parts: analyzing attitude singular constraint, analyzing attitude benchmark to determine error constraint, analyzing attitude mobility constraint and setting dual-vector included angle limiting range.
As shown in fig. 1, the geocentric vector and the sun vector are taken as attitude references:
1. and analyzing the three-axis attitude singularity determined by the geocentric vector and the sun vector.
By geocentric vector and sun vector r1(t) and r2(t) as an attitude reference, determining a three-axis attitude criterion chi according to the set geocentric vector and sun vector, and analyzing the three-axis attitude change rule of the spacecraft
And determining whether the gesture singularity phenomenon that the gesture is turned by plus or minus 180 degrees occurs.
As shown in fig. 2, the geocentric vector and the sun vector determine the three-axis attitude criterion χ: the Z axis is the geocentric vector direction, the Y axis is determined by the vector product of the sun vector and the geocentric vector, the X axis is determined by the right-hand rule according to the Y, Z axis, and when the geocentric vector is parallel to the sun vector, the yaw attitude is turned over by 180 degrees.
2. The geocentric vector and the sun vector are analyzed to determine an error constraint.
Step 2-1: and analyzing the geocentric vector and the sun vector to determine errors. And analyzing the determination errors of the geocentric vector and the sun vector by considering factors such as orbit calculation, orbit determination errors and the like according to the triaxial attitude criterion chi determined by the geocentric vector and the sun vector. For the orbit calculation and orbit determination error A, under the geocentric inertial system, the geocentric vector and the sun vector determination errors are A/R1 and A/R2, R1 is the spacecraft geocentric distance, and R2 is the sun geocentric distance.
Step 2-2: and analyzing the relation between the error magnification times and the included angles of the geocentric vector and the sun vector. According to the cross multiplication operation appearing in the three-axis attitude criterion chi, the smaller the included angle between the geocentric vector and the sun vector is, the larger the influence of the determined errors of the geocentric vector and the sun vector on the attitude direction is, and the amplification effect is presented.
The sun vector is basically unchanged in one-orbit time, the center-earth distance of the sun is far, and the sun vector r can be ignored2(t) error, geocentric vector r1(t) due to factors such as orbit determination, there is an error Δ r1Determining a three-axis attitude criterion χ according to the geocentric vector and the sun vector, the three-axis attitude pointing to
[(r2(t)×(r1(t)+Δr1))×(r1(t)+Δr1) r2(t)×(r1(t)+Δr1) r1(t)+Δr1]
According to the cross-multiplication algorithm, | c | ═ a × b | ═ a | · | b | sin<a,b>The smaller the included angle between the geocentric vector and the sun vector is, the smaller the sine function is, and the error delta r is1The greater the effect of the resulting angular deviation. As shown in fig. 3, the smaller the included angle between the geocentric vector and the solar vector is, the greater the influence of the geocentric vector and the solar vector determination accuracy on the attitude reference determination accuracy is, and the magnification effect is 3.09 times at 20 °.
Based on the steps 2-1 and 2-2, considering the requirements of the load pointing errors of the spacecrafts such as the antenna and the camera, determining the constraint gamma of the attitude reference determination error to the included angle between the geocentric vector and the sun vector1。
3. And analyzing the attitude mobility constraint.
Step 3-1: angular momentum constraints are analyzed. Using rigid body moment theorem H ═ I ω, pair
Difference estimation triaxial angular velocity time-varying rule [ omega ] is solvedx(t) ωy(t) ωz(t)]From the known three-axis moment of inertia [ I ] of spacecraftxIyIz]Estimating angular momentum H (t) required by a double-vector determined three-axis expected attitude on spacecraft tracking, determining the constraint gamma of the angular momentum to the double-vector included angle according to whether the configuration of a spacecraft attitude maneuver actuator meets the angular momentum requirement or not, wherein the lower the orbit height, the higher the inertial angular velocity of the spacecraft and the higher the angular momentum20。
Step 3-2: analysis of torqueAnd (4) restraining. According to the theorem of moment of momentum of rigid body
To [ omega ]x(t) ωy(t) ωz(t)]Calculating the time-varying law of differential estimation of triaxial angular acceleration
Estimating the moment M (t) required by the three-axis expected attitude determined by the ground center vector and the sun vector tracked by the spacecraft, determining the constraint gamma of the moment to the included angle of the double vectors according to whether the configuration of the spacecraft attitude maneuvering executing mechanism meets the requirement of the moment21。
Based on the steps 3-1 and 3-2, synthesizing a constraint gamma of the attitude maneuver capability to the dual-vector included angle2。
On the basis of the method for determining the double-vector included angle limit range based on constraint analysis, the invention also provides a system for determining the double-vector included angle limit range based on constraint analysis, which comprises the following steps:
the attitude singular constraint analysis module: with a chosen spatial dual vector r1(t) and r2(t) as an attitude reference, determining a three-axis attitude criterion χ according to the set double vectors, and analyzing the change rule of the three-axis attitude of the spacecraft along with the time t
Determining whether an attitude singularity phenomenon that the attitude turns over at plus or minus 180 degrees occurs;
an attitude reference determination error constraint analysis module: determining a triaxial attitude criterion chi according to the double vectors, analyzing the determined error of the double vectors, analyzing the relation between the error magnification factor and the included angle of the double vectors, and determining the constraint gamma of the determined error of the attitude reference to the included angle of the double vectors1；
An attitude mobility constraint analysis module: determining a constraint Γ of angular momentum to dual-vector angle20Determining the constraint gamma of the moment to the dual vector angle21Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
The double-vector included angle limiting range setting module: determining constraint gamma of error to double-vector included angle according to attitude singular constraint and attitude reference1Constraint gamma of dual vector angle with attitude maneuver capability2And setting a double-vector included angle limiting range gamma.
Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, units provided by the present invention as pure computer readable program code, the system and its various devices, modules, units provided by the present invention can be fully implemented by logically programming method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system and various devices, modules and units thereof provided by the invention can be regarded as a hardware component, and the devices, modules and units included in the system for realizing various functions can also be regarded as structures in the hardware component; means, modules, units for performing the various functions may also be regarded as structures within both software modules and hardware components for performing the method.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A method for determining a double-vector included angle limit range based on constraint analysis is characterized by comprising the following steps:
attitude singular constraint analysis: to selectDouble vectors r in definite space1(t) and r2(t) as an attitude reference, determining a three-axis attitude criterion χ according to the set double vectors, and analyzing the change rule of the three-axis attitude of the spacecraft along with the time t
Determining whether an attitude singularity phenomenon that the attitude turns over at plus or minus 180 degrees occurs;
and (3) attitude reference determination error constraint analysis: determining a triaxial attitude criterion chi according to the double vectors, analyzing the determined error of the double vectors, analyzing the relation between the error magnification factor and the included angle of the double vectors, and determining the constraint gamma of the determined error of the attitude reference to the included angle of the double vectors1；
And (3) attitude mobility constraint analysis: determining a constraint Γ of angular momentum to dual-vector angle20Determining the constraint gamma of the moment to the dual vector angle21Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
And (3) setting a double-vector included angle limiting range: determining constraint gamma of error to double-vector included angle according to attitude singular constraint and attitude reference1Constraint gamma of dual vector angle with attitude maneuver capability2And setting a double-vector included angle limiting range gamma.
2. The constraint analysis-based bivector included angle limitation range determination method as claimed in claim 1, wherein in the attitude reference determination error constraint analysis step, the constraint Γ of the attitude reference determination error on the bivector included angle is determined1And meanwhile, the requirement of spacecraft load pointing error including an antenna and a camera is considered.
3. The constraint analysis-based dual-vector included angle limitation range determination method according to claim 1, wherein the attitude mobility constraint analysis step comprises:
estimating the angle required by the three-axis expected attitude determined by double vectors on the spacecraft tracking by using the rigid body moment theoremMomentum H (t), determining the constraint gamma of the angular momentum to the double-vector included angle according to whether the configuration of the spacecraft attitude maneuver actuator meets the angular momentum requirement20；
Estimating the moment M (t) required by the triaxial expected attitude determined by double vectors on spacecraft tracking according to the law of moment of rigid momentum, determining the constraint gamma of the moment to the double-vector included angle according to whether the configuration of a spacecraft attitude maneuvering executing mechanism meets the requirement of the moment or not21；
Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
4. The constraint analysis-based dual-vector included angle limitation range determination method according to claim 1, wherein the dual vectors include a geocentric vector and a solar vector.
5. The constraint analysis-based bivector included angle limitation range determination method of claim 4, wherein the bivector determination of the tri-axial attitude criterion χ comprises: the X-axis is determined by the right hand rule according to the Y, Z axis, and when the centroid vector and the sun vector are parallel, the yaw attitude is flipped by 180 °.
6. A dual vector angle limit range determination system based on constraint analysis, comprising:
the attitude singular constraint analysis module: with a chosen spatial dual vector r1(t) and r2(t) as an attitude reference, determining a three-axis attitude criterion χ according to the set double vectors, and analyzing the change rule of the three-axis attitude of the spacecraft along with the time t
Determining whether an attitude singularity phenomenon that the attitude turns over at plus or minus 180 degrees occurs;
an attitude reference determination error constraint analysis module: determining a three-axis attitude criterion χ according to the double vectors, analyzing the determined error of the double vectors, and analyzing the included angle between the error magnification factor and the double vectorsDetermining the constraint gamma of the attitude reference determination error to the dual vector angle1；
An attitude mobility constraint analysis module: determining a constraint Γ of angular momentum to dual-vector angle20Determining the constraint gamma of the moment to the dual vector angle21Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
The double-vector included angle limiting range setting module: determining constraint gamma of error to double-vector included angle according to attitude singular constraint and attitude reference1Constraint gamma of dual vector angle with attitude maneuver capability2And setting a double-vector included angle limiting range gamma.
7. The constraint analysis-based dual-vector included angle limitation range determination system as claimed in claim 6, wherein in the attitude reference determination error constraint analysis module, the constraint Γ for determining the attitude reference determination error to the dual-vector included angle is1And meanwhile, the requirement of spacecraft load pointing error including an antenna and a camera is considered.
8. The constraint analysis-based dual-vector included angle limitation range determination system of claim 6, wherein the attitude mobility constraint analysis module comprises:
estimating angular momentum H (t) required by a three-axis expected attitude determined by double vectors on spacecraft tracking by using the rigid moment theorem, determining constraint gamma of the angular momentum to the included angle of the double vectors according to whether the configuration of a spacecraft attitude maneuvering executing mechanism meets the requirement of the angular momentum or not20；
Estimating the moment M (t) required by the triaxial expected attitude determined by double vectors on spacecraft tracking according to the law of moment of rigid momentum, determining the constraint gamma of the moment to the double-vector included angle according to whether the configuration of a spacecraft attitude maneuvering executing mechanism meets the requirement of the moment or not21；
Will constrain Γ20And constraint Γ21Constraint gamma of synthetic attitude maneuver capability to double-vector included angle2。
9. The constrained analysis-based dual-vector included angle limit range determination system of claim 6, wherein the dual vectors comprise a geocentric vector and a solar vector.
10. The constraint analysis-based dual-vector included angle limitation range determination system of claim 9, wherein the dual-vector determination of the tri-axial pose criterion χ comprises: the X-axis is determined by the right hand rule according to the Y, Z axis, and when the centroid vector and the sun vector are parallel, the yaw attitude is flipped by 180 °.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112572835A (en) * | 2020-12-15 | 2021-03-30 | 长光卫星技术有限公司 | Satellite in-orbit angular momentum management and control method with attitude switching function |
| CN113485395A (en) * | 2021-06-29 | 2021-10-08 | 北京控制工程研究所 | Fixed time relative attitude and orbit tracking control method under error constraint |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0949143A2 (en) * | 1998-04-10 | 1999-10-13 | Laser Communications International, L.L.C. | Method and apparatus for improved attitude determination of spacecraft |
| CN101858746A (en) * | 2010-03-26 | 2010-10-13 | 航天东方红卫星有限公司 | A Method for Analytical Determination of the Attitude of Satellite-oriented Sun Orientation Targets Effectively Avoiding the Effect of Earth, Atmospheric Light |
| US20140244210A1 (en) * | 2013-02-28 | 2014-08-28 | Philip Thomas Cobbin | System and method for real time determination of 3 axis orientation from reference vectors with vector misalignment estimation |
| CN106647793A (en) * | 2017-02-20 | 2017-05-10 | 上海航天控制技术研究所 | Magnetometer and sun sensor-based sun pointing method |
| CN108657467A (en) * | 2018-05-16 | 2018-10-16 | 清华大学 | A kind of spacecraft yawing maneuvering control method and system using virtual solar vector |
-
2019
- 2019-12-11 CN CN201911277694.2A patent/CN111123961B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0949143A2 (en) * | 1998-04-10 | 1999-10-13 | Laser Communications International, L.L.C. | Method and apparatus for improved attitude determination of spacecraft |
| CN101858746A (en) * | 2010-03-26 | 2010-10-13 | 航天东方红卫星有限公司 | A Method for Analytical Determination of the Attitude of Satellite-oriented Sun Orientation Targets Effectively Avoiding the Effect of Earth, Atmospheric Light |
| US20140244210A1 (en) * | 2013-02-28 | 2014-08-28 | Philip Thomas Cobbin | System and method for real time determination of 3 axis orientation from reference vectors with vector misalignment estimation |
| CN106647793A (en) * | 2017-02-20 | 2017-05-10 | 上海航天控制技术研究所 | Magnetometer and sun sensor-based sun pointing method |
| CN108657467A (en) * | 2018-05-16 | 2018-10-16 | 清华大学 | A kind of spacecraft yawing maneuvering control method and system using virtual solar vector |
Non-Patent Citations (4)
| Title |
|---|
| JUSTIN KRUGER: "Autonomous angles-only multitarget tracking for spacecraft swarms" * |
| 张剑慧: "捷联惯导系统双矢量定姿方法研究" * |
| 江洁: "一种改进的基于双矢量观测的姿态确定算法" * |
| 王献忠: "陀螺与磁强计组合定姿及陀螺漂移估计" * |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112572835A (en) * | 2020-12-15 | 2021-03-30 | 长光卫星技术有限公司 | Satellite in-orbit angular momentum management and control method with attitude switching function |
| CN112572835B (en) * | 2020-12-15 | 2022-07-05 | 长光卫星技术股份有限公司 | A satellite on-orbit angular momentum management and control method with attitude switching |
| CN113485395A (en) * | 2021-06-29 | 2021-10-08 | 北京控制工程研究所 | Fixed time relative attitude and orbit tracking control method under error constraint |
| CN113485395B (en) * | 2021-06-29 | 2024-03-15 | 北京控制工程研究所 | A fixed-time relative attitude-orbit tracking control method under error constraints |
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| CN111123961B (en) | 2023-08-08 |
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