CN109823572B - Actuating mechanism configuration and control method for reciprocating and rapid swinging of agile satellite attitude - Google Patents
Actuating mechanism configuration and control method for reciprocating and rapid swinging of agile satellite attitude Download PDFInfo
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Abstract
The invention discloses an actuating mechanism configuration and a control method for reciprocating and rapid swinging of postures of an agile satellite, wherein the actuating mechanism is set as a composite posture control actuating mechanism and comprises a control moment gyro and two reaction flywheels, a right-angle three-dimensional coordinate system Oxyz is established by taking an agile satellite platform as a center, the two reaction flywheels are respectively arranged on a y axis and a z axis, the positive rotation directions of rotors of the two reaction flywheels respectively face to the y axis and the z axis, the installation directions of frame shafts of the two control moment gyros are parallel to the y axis and form a double-control moment gyro structure, and the output control moment of the double-control moment gyro structure acts in the x axis direction. The control method comprises the steps of constructing a dynamics equation and a kinematics equation, designing a controller and then controlling the attitude maneuver of the satellite. The configuration and control method of the actuating mechanism effectively realizes the large-angle reciprocating rapid swing of the satellite attitude.
Description
Technical Field
The invention relates to the technical field of space control, in particular to an actuating mechanism configuration and a control method for reciprocating and rapid swinging of an agile satellite attitude.
Background
The remote sensing satellite has irreplaceable important functions in the civil and military application fields at present, and the agile remote sensing satellite with high-speed posture and wide-angle maneuvering capability receives more and more attention in order to improve the quick response capability and the imaging capability of the remote sensing satellite, shorten the revisit period of a ground target and consider high resolution and wide imaging. Because the agile satellite needs to complete the quick large-angle maneuvering of the attitude within the specified time, the system presents strong nonlinearity, and simultaneously needs to ensure the high-precision attitude stability, which provides great challenge for the design of a satellite attitude control system.
The typical satellite attitude control executing mechanism is mainly divided into an active control executing mechanism and an angular momentum exchange executing mechanism, wherein the active control executing mechanism comprises a spray pipe and a magnetic torquer, and the angular momentum exchange executing mechanism comprises a reaction flywheel and a control moment gyro. Because the output control torque is higher than that under the same power consumption, the control precision is higher than that of the spray pipe, and the pollution to the optical device is avoided, the actuator is an ideal actuator for agile satellites. At present, in order to meet the requirement of large-angle quick maneuvering, an agile satellite at home and abroad is mainly used as an actuating mechanism of the agile satellite, and the attitude maneuvering capability of the satellite reaches 4.5 degrees/s to the maximum. The Violet satellite of the Cannel university takes 8 control moment gyros as an actuating mechanism, and experiments prove that the maneuvering angular speed of the satellite can reach 10 degrees/s. By means of the large-angle quick maneuvering capability, the agile satellite can realize various working modes such as push-broom splicing imaging, three-dimensional splicing imaging, multipoint target quick imaging, dynamic scanning imaging and the like, and the observation capability of the satellite is greatly improved.
However, as the requirements of the current battlefield reconnaissance, disaster area monitoring and other tasks on imaging data with ultra-large width and high timeliness are higher and higher, the existing several types of imaging working modes cannot meet the increasing task requirements. In order to realize seamless continuous imaging of the ground target with ultra-large width in a limited time, the attitude of the agile satellite can be controlled to perform large-angle reciprocating rapid swing along the through-orbit direction. As shown in fig. 1, by the reciprocating and rapid swinging of the attitude of the agile satellite, a plurality of imaging strips (imaging strip 1, imaging strip 2 and imaging strip 3 …) can be completed, and one-time ultra-large-width overhead shooting can be completed by utilizing multi-strip seamless splicing. When the agile satellite is in the working mode, in order to ensure that the scanning imaging width of the agile satellite is larger in the track passing direction and ensure that the time for completing one side swinging period is less than the time for moving one width of the subsatellite point along the track direction, the agile satellite is required to perform side swinging maneuvering within a range of +/-45 degrees from the subsatellite point, and the maneuvering angular speed is kept at 20 degrees/s for a long time. The above requirements for mobility of agile satellites have greatly exceeded the capabilities of existing agile satellite attitude control systems. How to realize long-time large-angle quick maneuvering under the conditions of limited load and limited output control torque of an actuating mechanism provides new challenges for the design of an agile satellite attitude control system.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the configuration and control method of the actuating mechanism is simple in configuration, strong in maneuverability, good in rapidity and capable of effectively realizing large-angle reciprocating rapid swing of the satellite attitude.
In order to solve the technical problems, the invention is realized by the following technical scheme:
an actuating mechanism configuration method for reciprocating and rapid swinging of agile satellite attitude is characterized in that the actuating mechanism is set as a composite attitude control actuating mechanism mounted on an agile satellite platform, and comprises a control moment gyro and reaction flywheels, the number of the control moment gyro and the number of the reaction flywheels are both set to be two, and the configuration method comprises the following steps:
the method is characterized in that a rectangular three-dimensional coordinate system Oxyz is constructed by taking an agile satellite platform as a center, wherein two reaction flywheels are respectively installed on a y axis and a z axis, the positive rotation direction of a rotor of the reaction flywheels respectively faces to the y axis and the z axis, the frame axis installation direction of a control moment gyro is parallel to the y axis, a dual-control moment gyro structure is formed, the control moment output by the dual-control moment gyro structure acts on the x axis, the x axis is the rolling axis of an agile satellite, the agile satellite has rapid maneuvering capacity in the x axis direction, the control moment gyro controls the agile satellite to reciprocate rapidly along the rolling axis, and the reaction flywheels control the attitude of the agile satellite to be stable in the y and z axis directions.
Preferably, the two control moment gyros have the same structural parameters.
Preferably, the frame axes of the two control moment gyro structures are parallel to each other, the rotating speeds are equal and opposite, one frame axis rotates along the positive direction of the y axis, and the other frame axis rotates along the negative direction of the y axis.
Preferably, the frame angles and the frame angular velocities of the two control moment gyro structures are equally opposite.
The method for controlling the reciprocating rapid swinging of the attitude of the agile satellite comprises the following steps
Constructing a satellite attitude dynamics model based on the composite attitude control actuating mechanism, and specifically, utilizing a vector mechanics modeling method, firstly establishing an angular momentum equation of the whole satellite including a control moment gyro and a reaction flywheel about the system centroid, then utilizing an angular momentum theorem to obtain the relation between the system angular momentum change and the moment, finally projecting the equation to a satellite body system, and then establishing the satellite attitude dynamics model;
constructing a satellite attitude kinematics model based on the composite attitude control actuating mechanism, specifically, utilizing a quaternion attitude parameter representation method, firstly establishing an attitude kinematics equation of a satellite under a system, and then establishing a satellite attitude kinematics error equation according to the relation between an expected attitude quaternion and a current attitude quaternion;
and designing an attitude controller which is based on a composite attitude control executing mechanism and can control the agile satellite to maneuver according to the expected attitude by utilizing the constructed agile satellite attitude dynamics and kinematics model. The method comprises the following specific steps of firstly constructing a Lyapunov function of a system aiming at an attitude error system and deriving the Lyapunov function, then designing an attitude controller by utilizing the Lyapunov theorem and the LaSalle invariant set principle, outputting an instruction signal through the attitude controller, inputting the instruction signal into the composite attitude control executing mechanism, and enabling the composite attitude control executing mechanism to control a satellite to maneuver according to an expected attitude.
Compared with the prior art, the invention has the advantages that: the configuration and control method of the actuating mechanism for the reciprocating and rapid swinging of the attitude of the agile satellite has the advantages that the configuration of the actuating mechanism is simple, the control method is simple and clear, and the method has the following advantages:
the attitude maneuverability of the agile satellite is strong, the rapidity is good, the target information acquisition amount is greatly improved in a short time by controlling the attitude of the agile satellite to perform large-angle reciprocating rapid swing along the through-orbit direction within a specified time, and the requirements of ultra-large-breadth and high-timeliness imaging tasks such as current battlefield situation reconnaissance, disaster area target monitoring and the like can be met;
and the configuration of the composite attitude control actuating mechanism is simple, the agile micro-search satellite with light weight and small volume can be realized, and the use flexibility is improved.
Drawings
The invention is further described below with reference to the accompanying drawings:
FIG. 1 is a schematic diagram of a principle of realizing seamless continuous imaging with an ultra-large width by large-angle reciprocating rapid swing of an existing agile satellite;
FIG. 2 is a schematic diagram of a composite attitude control actuator configuration for an agile satellite according to the present invention;
FIG. 3 is a schematic structural configuration diagram of a dual control moment gyroscope according to the present invention;
FIG. 4 is an agile satellite attitude angle time response curve generated during a simulation experiment of the present invention;
FIG. 5 is an agile satellite attitude angular velocity time response curve generated during a simulation experiment of the present invention;
FIG. 6 is a control moment gyro frame angular time response curve generated during a simulation experiment of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all embodiments. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention:
as shown in fig. 2 to fig. 3, the actuator is a composite attitude control actuator mounted on an agile satellite platform, and includes a control moment gyro and a reaction flywheel, where the number of the control moment gyro and the number of the reaction flywheel are both two, and the method includes: the method comprises the steps of constructing a right-angle three-dimensional coordinate system Oxyz by taking an agile satellite platform as a center, wherein two reaction flywheels are respectively arranged on y and z axes, the positive rotation directions of the rotors of the reaction flywheels respectively face the negative directions of the y and z axes, and v1，v2The rotation direction unit vectors of the two rotors are respectively, the mounting direction of frame shafts of the two control moment gyroscopes is parallel to the y axis, and a double-control moment gyroscope structure is formed, the output control moment acts in the x axis direction, wherein the x axis is the rolling axis of the agile satellite, the agile satellite has rapid maneuvering capability in the x axis direction, the agile satellite is controlled to reciprocate rapidly around the rolling axis by the control moment gyroscopes, the postures of the agile satellite in the y and z axes directions are controlled to be stable by the accurate continuous moment with small amplitude provided by the reaction flywheel, in practical application, in order to meet the requirement of the agile satellite on the ultra-large-width seamless continuous imaging of a ground target, the agile satellite needs to carry out large-angle reciprocating rapid swinging in the through orbit direction, therefore,the agile satellite only needs to have the rapid maneuvering capability on the rolling axis and other two axes keep stable postures, the execution structure of the control moment gyroscope and the reaction flywheel is selected, so that the rotational inertia of the whole satellite is effectively reduced, the gyro can ensure the fast maneuvering ability around the rolling shaft by the larger output control moment of the control moment gyro, and ensure the stable postures of other two shafts by the accurate and continuous moment with smaller amplitude provided by the reaction flywheel, and further, because the control moment gyro acting on the roll axis direction can also generate a coupling effect on the yaw axis direction, considering that the torque output capacities of the reaction flywheel and the control moment gyro are greatly different, the reaction flywheel acting in the direction of the yaw axis cannot effectively compensate for such disturbing moments, therefore, the dual-control moment gyroscope structure can effectively eliminate the influence of the reaction flywheel on the yaw axis direction.
Further, for improving control accuracy and overall stability, satisfy the control demand of rapidity, two control moment gyro structure parameter the same, and two control moment gyro structure's frame axle be parallel to each other, the rotational speed equals but is opposite, a frame axle rotation angular velocity vector direction is along y axle positive direction, another frame axle rotation angular velocity vector direction is along y axle negative direction, two control moment gyro structure's frame angle and frame angular velocity reverse, in addition, further, when control moment gyro closes on the equivalence, or the angular momentum of reaction flywheel reaches when saturating, utilizes the active characteristic of magnetic torquer, unloads control moment gyro and the angular momentum of reaction flywheel.
The method for controlling the reciprocating rapid swinging of the attitude of the agile satellite by applying the actuating mechanism comprises the following steps
Firstly, constructing a satellite attitude dynamics model based on the composite attitude control actuating mechanism;
hereinafter, for an arbitrary matrix AaMatrix of
Or
Representation matrix AaTransposing; for arbitrary vectors aaVector of
Representing a vector aaDerivative, vector of
Representing a vector aaTransposing; if any vector aaCan be represented as aa＝[aa1,aa2,aa3]TThen [ a ] isa ×]Is represented by vector aaGenerated antisymmetric array, expressed as
[aa]diagIs represented by vector aaThe diagonal matrix is generated and expressed as
Specifically, firstly, the mass center of the whole satellite is set as O, the mass of the satellite platform except the control moment gyro and the reaction flywheel of the whole satellite is set as m, and the mass center of the satellite platform is set as mBC, a satellite body coordinate system b, and the mass center of the control moment gyro are
Hereinafter, the same definition is given for subscript i, the mass and center of mass of the reaction flywheel being
QiUnit vector g of control moment gyro frame rotation shaft directioniUnit vector s of direction of rotation axisiUnit vector t in the opposite direction of output torqueiControlling moment gyro frame fixed connection coordinate system ciUnit vector v of direction of rotation axis of reaction flywheeliThe other two axis direction unit vector ispi、liThe position vectors of the mass centers of the satellite platform, the control moment gyro and the reaction flywheel relative to the mass center of the system are rB、
The absolute angular velocity of the satellite platform is projected to be omega in the system, and the rotating speed vectors of the control moment gyro and the reaction flywheel rotor are respectively
Controlling moment gyro frame angle and frame angular velocity vector asi、
The rotational inertia of the satellite platform, the control moment gyroscope and the reaction flywheel is
The moment of inertia of the control moment gyro frame and the flywheel part are respectively
Firstly, according to the angular momentum calculation formula of the whole star about the system centroid:
in the formula (1), the reaction mixture is,
for the angular momentum of the platform with respect to the system centroid, the calculation formula is:
wherein,
is the rotational inertia of the satellite platform, omega is the projection of the absolute angular velocity of the satellite platform in the system, mBRemoving the control moment gyro and the mass r of the satellite platform outside the reaction flywheel for the whole satelliteBIs the position vector of the satellite platform relative to the system centroid;
in the formula (1), the reaction mixture is,
for controlling the angular momentum of the moment gyro about the centroid, the calculation formula is as follows:
wherein,
for controlling moment of inertia of moment gyros, in particular
Wherein,
are respectively a control moment gyro at gi、si、tiThe rotational inertia of the direction, omega, is the projection of the absolute angular velocity of the satellite platform in the system,
to control the moment gyro frame angular velocity vector,
to control the moment of inertia of the flywheel portion of the moment gyro, specifically,
wherein,
the flywheel parts of the gyros are respectively in gi、si、tiThe moment of inertia of the direction of rotation,
in order to control the rotation speed vector of the moment gyro,
in order to control the mass of the moment gyro,
is the position vector of the center of mass of the control moment gyro relative to the center of mass of the system;
in the formula (1), the reaction mixture is,
to counteract the angular momentum of the flywheel about the center of mass, the formula is calculated as:
wherein,
to react the moment of inertia of the flywheel, specifically,
wherein,
respectively, reaction flywheel at pi、vi、liThe rotational inertia of the direction, omega, is the projection of the absolute angular velocity of the satellite platform in the system,
in order to react to the rotational speed vector of the flywheel rotor,
in order to react to the center of mass of the flywheel,
is the position vector of the center of mass of the reaction flywheel relative to the center of mass of the system;
the invariant portion of the total star moment of inertia is expressed as:
Due to the fact that
Fastening a coordinate system c to the frameiIf the middle is also constant, then the external moment is combined according to the theorem of moment of momentum
In the above formula
The derivatives of the absolute angular velocity of the satellite platform, the angular velocity of the control moment gyro frame and the angular velocity of the flywheel rotor are respectively.
The above formula is projected in the main system b, because the control moment gyro frame is fixedly connected with the system ciCoordinate transformation array to body system b
Can be expressed as:
in the above formula, giUnit vector s of gyro frame rotation shaft direction for controlling momentiIs unit vector of rotation axis direction, tiFor output torque reactionThe unit vector of the direction is,ifor controlling the angle of the moment gyro frame, g0i、s0i、t0iAre respectively gi、si、tiIs started.
According to the relation of the projected rotational inertia in different coordinate systems,
project in the main system b
Can be expressed as:
in the above formula, matrix Ap、Av、AlAre respectively Ap＝[p1,p2]、Av＝[v1,v2]、Al＝[l1,l2]The matrix Ip、Iv、IlAre respectively as
Project in the main system b
Can be expressed as:
in the above formula, matrix Ag、As、AtAre respectively Ag＝[g1,g2]、As＝[s1,s2]、At＝[t1,t2]The matrix Ig、Is、ItAre respectively as
Then, the order is executed again,
due to the fact that
(
Are respectively as
A projection in the corresponding direction),
the resultant external moment is:
order to
The above formula can be represented by the following formula,
wherein,
C1＝C+R(ω)+ω×AgIg,
DΩ＝AvIv,
the equation (2) is the constructed agile satellite attitude dynamics model.
Then, constructing an attitude kinematics equation based on the composite attitude control actuating mechanism;
specifically, let the agile satellite attitude four elements be
The agile satellite attitude kinematics equation expressed by the attitude quaternion is
Setting the expected attitude four elements and the expected attitude angular velocity of the agile satellite as
Error four elements and error angular velocity of
Then there is
Wherein, the coordinate transformation matrix is fixedly connected to the body coordinate system from the expected attitude of the satellite
Is composed of
Wherein E is33 × 3 unit matrix, and the error four-element and the error angular velocity satisfy the following equation:
and finally, designing an algorithm model which is based on a conforming attitude control executing mechanism and can enable the satellite to maneuver according to the expected attitude according to the constructed agile satellite attitude dynamics and kinematics equation, namely the attitude controller.
Specifically, according to the constructed agile satellite attitude dynamics and kinematic equations, by utilizing the Lyapunov theorem and the LaSalle invariant set principle, the attitude controller can be designed to enable an attitude error system to be gradually stable, namely the satellite maneuvers according to the expected attitude.
According to the theorem of the derivative relationship of the vector with respect to different coordinate systems with respect to time, there are
According to equations (2) and (6), there are
Let k1,k2> 0, where k1,k2The angle error gain and the angular velocity error gain are obtained by the Lyapunov function:
the derivation of the above formula is obtained,
order to
Then
And the equal sign is only at omegaeIs taken when the value is 0And (4) obtaining. Require that
Then need omegaeIs equal to 0. Will TcSubstituting the expression into the system dynamics equation (2) can obtain
When ω iseAt ≡ 0, there are
And q ise0. According to the LaSalle invariant set principle, any trajectory of the system tends to qe＝0,ω e0. Thus, the error system asymptotically stabilizes.
Then, according to equation (3), the control moment gyro and the reaction flywheel manipulation equation can be obtained:
wherein
Due to the fact that
Much less than
The generated moment, omega, is much less than omegacThus, the steering equations can be reduced to
According to the structural configuration of the double-control moment gyro, the frame angles and the frame angular velocities of the two control moment gyros are equivalently reversed, thereby enabling the two control moment gyros to have the same value
Then, according to the configuration scheme of the actuator, the input of the actuator of the agile satellite can be obtained as follows:
ucnamely, the command signal is output by the attitude controller, and the agile satellite maneuvers according to the expected attitude by inputting the command signal into the composite attitude executing mechanism.
In the following, a more detailed description of embodiments of the invention will be given in connection with Simulink simulations of Matlab,
setting the detection width D of an agile satellite as 1000km and the relative movement speed V of the satellite to the groundd6.8km/s, the width of the satellite point L × W117 × 21km, corresponding to the field of view thetaL×θW13.41 × 2.41.41, and L117 km is the in-track direction, therefore, the required swing angle θ in the through-track direction is the required swing angle θ for the whole star imaging width DD＝DθWand/W is 90 °. If the seamless continuous imaging is to be realized, the satellite swings back to the initial position to complete the longest time limit t of one pendulummax＝L/Vd＝17.2s。
In order to verify the effectiveness of the invention when the agile satellite frequently carries out a large-angle reciprocating maneuver task, the attitude maneuver instruction is a rolling instruction which is executed from 0s to 45 degrees → 45 degrees, and when the rolling angle of the satellite reaches 45 degrees, the rolling instruction of 45 degrees → 45 degrees is executed again, and the process is repeated. Based on this, various simulation parameters can be set as:
1. satellite parameters
A rotational inertia matrix: j ═ 7,0, 0; 0,12, 0; 0,0,12]kgm2，
Initial Euler angle：
Initial angular velocity: omega0＝[0,0,0]T，
Desired euler angle:
the clock pendulum period: t is 16s,
2. controlling moment gyro parameters
Rotor shaft direction moment of inertia:
rotor speed:
initial frame angle:
maximum frame angular velocity:
3. reaction flywheel parameters
Reaction flywheel axial moment of inertia:
initial speed of reaction flywheel:
maximum speed of reaction flywheel:
maximum angular acceleration of reaction flywheel:
4. controller parameters
Angle error gain: k is a radical of1＝100，
Angular velocity error gain: k is a radical of2＝100。
According to the simulation parameters, the Simulink simulation process is carried out, and the simulation time is set to be 160s, namely ten pendulum cycles.
The simulation result is as follows:
as shown in fig. 4, the satellite completes the roll maneuver of-45 ° → 45 °, further 45 ° → -45 ° within 16s, and the attitude angle tracking accuracy of the satellite three axes is better than 0.1 °.
As shown in fig. 5, the satellite completes the acceleration process from 0 °/s → 17.7 °/s within 4s, and completes the deceleration and then acceleration process from 17.7 °/s → 0 → -17.7 °/s within 8 s.
As shown in fig. 6, the initial frame angle is 90 °, gradually increasing to 121 °, and decreasing to 59 °. After one pendulum period is finished, the frame angle returns to the vicinity of the initial frame angle, and the occurrence of the singular condition of the frame angle is avoided.
The simulation results of fig. 4 to fig. 6 verify that the attitude control method provided by the invention can realize the large-angle rapid reciprocating swing of the roll angle, can meet the requirement of the ground target on the ultra-large-width seamless continuous maneuvering imaging, and has the characteristics of strong maneuvering capability, good rapidity, avoidance of frame angle singularity and the like.
It is to be emphasized that: the above embodiments are only preferred embodiments of the present invention, and are not intended to limit the present invention in any way, and all simple modifications, equivalent changes and modifications made to the above embodiments according to the technical spirit of the present invention are within the scope of the technical solution of the present invention.
Claims (5)
1. An executing mechanism configuration method for reciprocating and rapid swinging of an agile satellite attitude is characterized by comprising the following steps: the actuating mechanism is a composite attitude control actuating mechanism which is arranged on the agile satellite platform in a loading mode, and comprises a control moment gyro and two reaction flywheels, wherein the number of the control moment gyro and the number of the reaction flywheels are respectively set to be two, and the configuration method comprises the following steps:
the method is characterized in that a rectangular three-dimensional coordinate system Oxyz is constructed by taking an agile satellite platform as a center, wherein two reaction flywheels are respectively installed on a y axis and a z axis, the positive rotation direction of a rotor of the reaction flywheels respectively faces to the y direction and the z axis negative direction, the frame axis installation direction of a control moment gyro is parallel to the y axis, a dual-control moment gyro structure is formed, the control moment output by the dual-control moment gyro structure acts on the x axis direction, the x axis is the rolling axis of an agile satellite, the agile satellite has rapid maneuvering capacity in the x axis direction, the control moment gyro controls the agile satellite to reciprocate rapidly maneuvering around the rolling axis, and the reaction flywheels control the attitude of the agile satellite to be stable in the y and z axis directions.
2. The method for configuring an actuating mechanism for reciprocating and rapidly swinging an attitude of an agile satellite according to claim 1, wherein: the two control moment gyros have the same structural parameters.
3. The method for configuring an actuating mechanism for reciprocating and rapidly swinging an attitude of an agile satellite according to claim 1, wherein: the frame shafts of the two control moment gyroscope structures are parallel to each other, the rotating speeds are equal and opposite, the rotating angular speed vector direction of one frame shaft is along the positive direction of the y shaft, and the rotating angular speed vector direction of the other frame shaft is along the negative direction of the y shaft.
4. The method for configuring an actuating mechanism for reciprocating and rapidly swinging an attitude of an agile satellite according to claim 1, wherein: and the frame angles and the frame angular velocities of the two control moment gyroscope structures are in equal and opposite directions.
5. The method for configuring the actuating mechanism for the agile satellite to perform reciprocating fast swinging on the attitude is characterized by comprising the following steps of: comprises that
Constructing a satellite attitude dynamics model based on the composite attitude control actuating mechanism, and specifically, utilizing a vector mechanics modeling method, firstly establishing an angular momentum equation of the whole satellite including a control moment gyro and a reaction flywheel about the system centroid, then utilizing an angular momentum theorem to obtain the relation between the system angular momentum change and the moment, finally projecting the equation to a satellite body system, and then establishing the satellite attitude dynamics model;
constructing a satellite attitude kinematics model based on the composite attitude control actuating mechanism, specifically, utilizing a quaternion attitude parameter representation method, firstly establishing an attitude kinematics equation of a satellite under a system, and then establishing a satellite attitude kinematics error equation according to the relation between an expected attitude quaternion and a current attitude quaternion;
the method comprises the steps of firstly constructing a Lyapunov function of a system aiming at an attitude error system and deriving the Lyapunov function, then designing an attitude controller by utilizing the Lyapunov theorem and the LaSalle invariant set principle, outputting an instruction signal through the attitude controller, inputting the instruction signal into the composite attitude control executing mechanism, and then enabling the composite attitude control executing mechanism to control the satellite to maneuver according to the expected attitude.
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| CN117719702B (en) * | 2024-02-07 | 2024-04-26 | 长光卫星技术股份有限公司 | Composite attitude control method, equipment and medium for high-precision staring of satellite |
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| WO2012009198A2 (en) * | 2010-07-14 | 2012-01-19 | University Of Florida Research Foundation, Inc. | System and method for assessing the performance of an attitude control system for small satellites |
| CN102627151B (en) * | 2012-05-09 | 2014-07-02 | 哈尔滨工业大学 | Moment distribution method for rapid maneuvering satellite based on mixed actuating mechanism |
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