CN109823572B - Actuator configuration and control method for rapid reciprocating swing of agile satellite attitude - Google Patents

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

Actuator configuration and control method for rapid reciprocating swing of agile satellite attitude

technical field

The invention relates to the technical field of aerospace control, in particular to the configuration and control method of an actuator for reciprocating and fast swinging of an agile satellite attitude.

Background technique

Remote sensing satellites currently play an irreplaceable and important role in both civil and military applications. In order to improve the rapid response and imaging capabilities of remote sensing satellites, shorten the revisit cycle of ground targets, take into account high-resolution and wide-format imaging, and have high-speed attitude and large angles. Maneuverable agile remote sensing satellites have received increasing attention. Since the agile satellite needs to complete the attitude fast and large-angle maneuver within the specified time, the system exhibits strong nonlinearity, and at the same time, it needs to ensure high-precision attitude stability, which poses a great challenge to the design of the satellite attitude control system.

Typical satellite attitude control actuators are mainly divided into active control actuators, including nozzles and magnetic torquers, and angular momentum exchange actuators, including reaction flywheels and control torque gyroscopes. Because the output control torque is higher than that under the same power consumption, the control accuracy is higher than that of the nozzle, and it will not cause pollution to the optical devices, so it is an ideal actuator for agile satellites. At present, in order to meet the needs of large-angle and fast maneuvering, domestic and foreign agile satellites are mainly used as their actuators, and the satellite's attitude maneuverability can reach a maximum of 4.5°/s. The Violet satellite of Cornell University uses eight control moment gyroscopes as the actuators, and experiments have confirmed that the satellite's maneuvering angular velocity can reach 10°/s. Relying on the ability of large-angle and fast maneuvering, Agile Satellite can realize various working modes such as push-broom mosaic imaging, three-dimensional mosaic imaging, multi-point target fast imaging, dynamic scanning imaging, etc., which greatly improves the observation capability of the satellite.

However, as the current tasks such as battlefield reconnaissance and disaster area monitoring have higher and higher requirements for ultra-wide and time-sensitive imaging data, the existing several types of imaging work modes cannot meet the increasing task requirements. In order to realize the seamless continuous imaging of the ground target with super wide width in a limited time, the attitude of the agile satellite can be controlled to reciprocate rapidly at a large angle along the orbital direction. As shown in Figure 1, multiple imaging strips (imaging strip 1, imaging strip 2, imaging strip 3...) can be completed through the reciprocating and rapid swing of the attitude of the agile satellite, which can be completed by seamless splicing of multiple strips A super wide over-the-top shot. When the agile satellite is in this type of working mode, in order to ensure a larger scanning imaging width in the orbital direction, and at the same time to ensure that the time to complete a side-swing cycle is less than the time for the sub-satellite point to move a width in the orbital direction, agility is required. The satellite performs a side-swing maneuver within a range of ±45° away from the sub-satellite point, and the maneuvering angular velocity remains at 20°/s for a long time. The above maneuverability requirements for agile satellites have greatly exceeded the capabilities of existing agile satellite attitude control systems. How to achieve long-term high-angle and fast maneuvering under the condition of limited load and limited output control torque of the actuator poses a new challenge to the design of agile satellite attitude control system.

SUMMARY OF THE INVENTION

The technical problem to be solved by the present invention is to provide an actuator configuration and control method which is simple in configuration, strong in maneuverability, good in rapidity and can effectively realize the reciprocating and fast swinging of the satellite attitude at a large angle.

In order to solve the above-mentioned technical problems, the present invention is achieved through the following technical solutions:

A method for configuring an actuator for agile satellite attitude reciprocating rapid swing, the actuator is configured to carry a composite attitude control actuator installed on an agile satellite platform, including a control torque gyro and a reaction flywheel, the control torque gyro and the reaction flywheel The number of are set to two, and the configuration method is:

A rectangular three-dimensional coordinate system Oxyz is constructed with the agile satellite platform as the center, wherein the two reaction flywheels are respectively installed on the y and z axes, and the positive directions of their rotor rotations are respectively directed towards the negative directions of the y and z axes, and the two control torques The installation direction of the frame axis of the gyro is parallel to the y-axis, and constitutes a dual-control torque gyro structure. The output control torque acts in the x-axis direction, where the x-axis is the roll axis of the agile satellite, and the agile satellite has the ability to quickly maneuver in the x-axis direction. , by controlling the moment gyro to control the agile satellite to reciprocate quickly around the roll axis, and the reaction flywheel to control the attitude stability of the agile satellite in the y and z directions.

Preferably, the structural parameters of the two control torque gyroscopes are the same.

Preferably, the frame axes of the two control torque gyro structures are parallel to each other, and the rotational speeds are equal but opposite.

Preferably, the frame angles and frame angular velocities of the two control torque gyro structures are equivalently reversed.

The control method for the reciprocating rapid swing of the agile satellite attitude, comprising:

Constructing a satellite attitude dynamics model based on the compound attitude control actuator, the specific steps are to use the modeling method of vector mechanics to first establish the angular momentum equation of the whole satellite including the control moment gyroscope and the reaction flywheel about its system center of mass, Then, the relationship between the system angular momentum change and the moment is obtained by using the angular momentum theorem, and finally the above equation is projected to the satellite system, and then the satellite attitude dynamics model is established;

Constructing the satellite attitude kinematics model based on the compound attitude control actuator, the specific steps are to use the quaternion attitude parameter representation method, first establish the attitude kinematic equation of the satellite under this system, and then according to the expected attitude quaternion and current The relationship between attitude quaternions, establish satellite attitude kinematic error equation;

Using the constructed agile satellite attitude dynamics and kinematics model, an attitude controller based on a compound attitude control actuator that can control the agile satellite to maneuver according to the desired attitude is designed. The specific steps are to first construct the Lyapunov function of the system for the attitude error system and derive it, and then use the Lyapunov theorem and the LaSalle invariant set principle to design the attitude controller, output the command signal through the attitude controller, and input the command signal into the composite The attitude control actuator, and then the composite attitude control actuator controls the satellite to maneuver according to the desired attitude.

Compared with the prior art, the advantages of the present invention are: the configuration and control method of the actuator for the agile satellite attitude reciprocating fast swing, the configuration of the actuator is simple, the control method is concise and clear, and has the following advantages:

1. The attitude of the agile satellite is highly maneuverable and fast. By controlling the attitude of the agile satellite, it swings back and forth at a large angle along the orbiting direction within a specified time, which greatly improves the acquisition of target information in a short period of time. Current requirements for ultra-wide and time-sensitive imaging tasks such as battlefield situational reconnaissance and target monitoring in disaster-stricken areas;

2. The configuration of its composite attitude control actuator is simple, which can realize the agility of light weight and small volume, and improve the flexibility of use.

Description of drawings

Below in conjunction with accompanying drawing, the present invention is further described:

Figure 1 is a schematic diagram of the existing agile satellite attitude reciprocating and fast swinging at a large angle to achieve ultra-large-width seamless continuous imaging;

Fig. 2 is the composite attitude control executive mechanism configuration schematic diagram of the agile satellite in the present invention;

3 is a schematic diagram of the configuration and configuration of the dual control moment gyro in the present invention;

Fig. 4 is the agile satellite attitude angle time response curve generated in the simulation experiment process of the present invention;

Fig. 5 is the agile satellite attitude angular velocity time response curve generated in the simulation experiment process of the present invention;

FIG. 6 is a control torque gyro frame angle time response curve generated during the simulation experiment of the present invention.

Detailed ways

The technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by those of ordinary skill in the art without making creative work all belong to the protection scope of the present invention:

As shown in FIG. 2 to FIG. 3 , a method for configuring an actuator for agile satellite attitude reciprocating and rapid swing, the actuator is configured to carry a compound attitude control actuator installed on the agile satellite platform, including a control moment gyroscope and a reaction flywheel , the number of the control moment gyroscope and the reaction flywheel is set to two, and the configuration method is: constructing a rectangular three-dimensional coordinate system Oxyz with the agile satellite platform as the center, wherein the two reaction flywheels are respectively installed on the y and z axes. and the positive directions of its rotor rotation are respectively toward the negative directions of the y and z axes, where v ₁ and v ₂ are the unit vectors of the rotation directions of the two rotors, respectively, and the frame axes of the two control torque gyroscopes are installed in the direction parallel to the y axis. , and constitutes a dual control torque gyroscope structure, the output control torque acts in the x-axis direction, where the x-axis is the roll axis of the agile satellite, the agile satellite has the ability to quickly maneuver in the x-axis direction, and the agile satellite is controlled by the control torque gyroscope. The rotating shaft reciprocates rapidly, and the precise and continuous torque provided by the reaction flywheel is used to control the attitude stability of the agile satellite in the y and z directions. To meet the requirements of wide and seamless continuous imaging, the agile satellite needs to swing back and forth at a large angle in the orbital direction. Therefore, the agile satellite only needs to have fast maneuvering ability in the roll axis, and the other two axes can keep the attitude stable, and the above control torque gyroscope is used. And the execution structure of the reaction flywheel can effectively reduce the rotational inertia of the whole star, and rely on the large output control torque of the control torque gyro to ensure the rapid maneuverability around the roll axis, relying on the reaction flywheel. A small torque is used to ensure the attitude stability of the other two axes. Furthermore, since the control torque gyroscope acting in the direction of the roll axis will also have a coupling effect on the yaw axis direction, considering that the torque output capability of the reaction flywheel and the control torque gyroscope is very different , so the reaction flywheel acting in the direction of the yaw axis cannot effectively compensate for this interference torque. Therefore, the above-mentioned dual-control torque gyro structure can effectively eliminate the influence of the reaction flywheel on the direction of the yaw axis.

Further, in order to improve the control accuracy and overall stability and meet the rapid control requirements, the two control torque gyroscopes have the same structural parameters, and the frame axes of the two control torque gyroscope structures are parallel to each other, and the rotational speeds are equal but opposite. , the direction of the rotational angular velocity vector of one frame axis is along the positive direction of the y-axis, and the direction of the rotational angular velocity vector of the other frame axis is along the negative direction of the y-axis. Further, when the control torque gyroscope is approaching the singularity, or the angular momentum of the reaction flywheel reaches saturation, the angular momentum of the control torque gyroscope and the reaction flywheel is unloaded by using the active characteristics of the magnetic torquer.

The control method of the agile satellite attitude reciprocating fast swing using the above-mentioned actuator, including

First, construct a satellite attitude dynamics model based on the composite attitude control actuator;

或

表示矩阵A_a的转置；对于任意矢量a_a，矢量

表示矢量a_a的导数，矢量

表示矢量a_a的转置；若任意矢量a_a可表示为a_a＝[a_a1,a_a2,a_a3]^T，则[a_a ^×]表示由矢量a_a生成的反对称阵，表示为In the following, for any matrix A _a , the matrix

or

represents the transpose of matrix A _a ; for any vector a _a , the vector

represents the derivative of the vector a _a , the vector

Represents the transpose of the vector a _a ; if any vector a _a can be represented as a _a =[a _a1 ,a _a2 ,a _a3 ] ^T , then [a _a ^× ] represents the antisymmetric matrix generated by the vector a _a , which is expressed as

[a _a ] ^diag represents the diagonal matrix generated by the vector a _a , denoted as

下文中关于下标i皆有相同的定义，反作用飞轮的质量和质心为

Q_i，控制力矩陀螺框架转轴方向单位矢量g_i、自转轴方向单位矢量s_i、输出力矩反方向单位矢量t_i，控制力矩陀螺框架固连坐标系c_i，反作用飞轮转轴方向单位矢量v_i、另两轴方向单位矢量为p_i、l_i，卫星平台、控制力矩陀螺及反作用飞轮的质心相对于系统质心的位置矢量分别为r_B、

卫星平台绝对角速度在本体系中投影为ω，控制力矩陀螺和反作用飞轮转子的转速矢量分别为

控制力矩陀螺框架角及框架角速度矢量为δ_i、

卫星平台、控制力矩陀螺、反作用飞轮的转动惯量为

控制力矩陀螺框架和飞轮部分的转动惯量分别为

Specifically, first, let the center of mass of the whole star be O, the mass and center of mass of the satellite platform outside the control moment gyro and the reaction flywheel are m _B , C, the satellite body coordinate system b, the mass and center of mass of the control moment gyro are

The following subscript i has the same definition, the mass and center of mass of the reaction flywheel are

Q _i , the unit vector g _i in the direction of the rotation axis of the control moment gyro frame, the unit vector s _i in the direction of the rotation axis, the unit vector t _i in the reverse direction of the output torque, the fixed coordinate system c _i of the control moment gyro frame, the unit vector v _i in the direction of the rotation axis of the reaction flywheel , the other two axis direction unit vectors are p _i , li _, the position vectors of the center of mass of the satellite platform, the control moment gyro and the reaction flywheel relative to the center of mass of the system are r _B ,

The absolute angular velocity of the satellite platform is projected as ω in this system, and the rotational speed vectors of the control moment gyro and the reaction flywheel rotor are respectively

The frame angle and frame angular velocity vector of the control moment gyro are δ _i ,

The moment of inertia of the satellite platform, control moment gyroscope, and reaction flywheel is

The moment of inertia of the control moment gyro frame and the flywheel part are respectively

First, according to the calculation formula of the angular momentum of the whole star about the center of mass of the system:

为平台关于系统质心的角动量，计算公式为：

其中，

为卫星平台的转动惯量，ω为卫星平台绝对角速度在本体系中投影，m_B为整星除去控制力矩陀螺及反作用飞轮外卫星平台质量，r_B为卫星平台相对于系统质心的位置矢量；In formula (1),

is the angular momentum of the platform about the center of mass of the system, the calculation formula is:

in,

is the moment of inertia of the satellite platform, ω is the projection of the absolute angular velocity of the satellite platform in this system, m _B is the mass of the satellite platform outside the whole satellite except the control moment gyro and the reaction flywheel, r _B is the position vector of the satellite platform relative to the center of mass of the system;

为控制力矩陀螺关于质心的角动量，计算公式为：

其中，

为控制力矩陀螺的转动惯量，具体地

其中，

分别为控制力矩陀螺在g_i、s_i、t_i方向的转动惯量，ω为卫星平台绝对角速度在本体系中投影，

为控制力矩陀螺框架角速度矢量，

为控制力矩陀螺飞轮部分的转动惯量，具体地，

其中，

分别为控制力矩陀螺飞轮部分在g_i、s_i、t_i方向的转动惯量，

为控制力矩陀螺的转速矢量，

为控制力矩陀螺的质量，

为控制力矩陀螺的质心相对于系统质心的位置矢量；In formula (1),

In order to control the angular momentum of the moment gyro about the center of mass, the calculation formula is:

in,

In order to control the moment of inertia of the moment gyro, specifically

in,

are the moment of inertia of the control moment gyro in the directions of g _i , s _i and t _i respectively, ω is the projection of the absolute angular velocity of the satellite platform in this system,

To control the angular velocity vector of the moment gyro frame,

In order to control the moment of inertia of the flywheel part of the moment gyro, specifically,

in,

are the moment of inertia of the flywheel part of the control moment gyro in the directions of g _i , s _i , and t _i , respectively,

To control the rotational speed vector of the moment gyro,

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;

为反作用飞轮关于质心的角动量，其计算公式为：

其中，

为反作用飞轮的转动惯量，具体地，

其中，

分别为反作用飞轮在p_i、v_i、l_i方向的转动惯量，ω为卫星平台绝对角速度在本体系中投影，

为反作用飞轮转子的转速矢量，

为反作用飞轮的质心，

为反作用飞轮的质心相对于系统质心的位置矢量；In formula (1),

is the angular momentum of the reaction flywheel about the center of mass, and its calculation formula is:

in,

is the moment of inertia of the reaction flywheel, specifically,

in,

are the moment of inertia of the reaction flywheel in the directions of p _i , vi _, and li respectively _, ω is the projection of the absolute angular velocity of the satellite platform in this system,

is the rotational speed vector of the reaction flywheel rotor,

is the center of mass of the reaction 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 part of the moment of inertia of the whole star is expressed as the following formula:

其中，E为单位矩阵，

分别为r_B、

的转置。

where E is the identity matrix,

are r _B ,

transposition of .

在框架固连坐标系c_i中也为常量，则根据动量矩定理，合外力矩because

is also constant in the frame-fixed coordinate system c _i , then according to the moment of momentum theorem, the combined external moment

分别为卫星平台绝对角速度、控制力矩陀螺框架角速度及飞轮转子角速度的导数。in the above formula

are the absolute angular velocity of the satellite platform, the angular velocity of the control moment gyro frame and the derivative of the angular velocity of the flywheel rotor.

可表示为：The above formula is projected in the system b, because the coordinate transformation matrix from the control moment gyro frame fixed connection c _i to the system b

can be expressed as:

In the above formula, _gi is the unit vector of the direction of the rotation axis of the control torque gyro frame, s _i is the unit vector of the direction of the rotation axis, t _i is the unit vector of the reverse direction of the output torque, δ _i is the frame angle of the control torque gyro, g _0i , s _0i , t _0i are the initial values of g _i , s _i , and t _i respectively.

在本体系b中投影

可表示为：According to the relational expression of the moment of inertia projected in different coordinate systems,

Projection in this system b

can be expressed as:

在本体系b中投影

可表示为：In the above formula, the matrices A _p , A _v , and A _l are respectively A _p =[p ₁ ,p ₂ ], A _v =[v ₁ ,v ₂ ], A _l =[l ₁ ,l ₂ ], the matrix I _p , I _v , I _l are respectively

Projection in this system b

can be expressed as:

In the above formula, the matrices A _g , A _s and At are respectively A _g =[g ₁ ,g ₂ ], A _s =[s ₁ ,s ₂ ],A _t =[ _{t 1} ,t ₂ ], the matrix I _g , _Is , and _It are respectively

order again,

(

分别为

在相应方向的投影)，because

(

respectively

projection in the corresponding direction),

Then the resultant external moment is:

make

The above formula can be transformed into the following formula,

in,

C _δ1 =C _δ +R(ω)+ω×A _g I _g ,

D _Ω =A _v I _v ,

The above equation (2) is the constructed agile satellite attitude dynamics model.

Then, construct the attitude kinematics equation based on the compound attitude control actuator;

用姿态四元数表示的敏捷卫星姿态运动学方程为Specifically, let the four elements of agile satellite attitude be

The attitude kinematics equation of agile satellite expressed by attitude quaternion is:

误差四元素和误差角速度为

则有Let the four elements of the desired attitude of the agile satellite and the angular velocity of the desired attitude be

The error four elements and the error angular velocity are

then there are

为

其中，E₃为3×3单位矩阵。同时，误差四元素和误差角速度也满足下式：Among them, the coordinate transformation matrix that is fixedly connected to the body coordinate system by the desired attitude of the satellite

for

where E ₃ is a 3×3 identity matrix. At the same time, the error four elements and the error angular velocity also satisfy the following formula:

Finally, according to the above constructed agile satellite attitude dynamics and kinematic equations, an algorithm model based on compliance with the attitude control actuator and enabling the satellite to maneuver according to the desired attitude, that is, the attitude controller is designed.

Specifically, according to the constructed agile satellite attitude dynamics and kinematic equations, using Lyapunov theorem and LaSalle invariant set principle, the attitude controller can be designed to make the attitude error system asymptotically stable, that is, the satellite can maneuver according to the desired attitude.

According to the theorem of the derivatives of vectors with respect to different coordinate systems with respect to time, we have

Then according to equations (2) and (6), we have

Assuming k ₁ , k ₂ >0, where k ₁ , k ₂ are the angular error gain and the angular velocity error gain, respectively, then the Lyapunov function can be obtained:

Deriving from the above formula,

make

且等号仅在ω_e＝0时才取到。要求

则需ω_e≡0。将T_c表达式代入系统动力学方程(2)中可得but

And the equal sign is only taken when ω _e =0. Require

Then ω _e ≡ 0 is required. Substituting the expression of T _c into the system dynamics equation (2), we can get

且q_e＝0。根据LaSalle不变集原理，系统任何轨线都趋向于q_e＝0,ω_e＝0。从而，误差系统渐近稳定。When ω _e ≡ 0, we have

and q _e =0. According to the LaSalle invariant set principle, any trajectory of the system tends to q _e =0,ω _e =0. Thus, the error system is asymptotically stable.

Then according to equation (3), the steering equations for controlling the moment gyro and the reaction flywheel can be obtained:

in

远小于

产生的力矩，ω远小于Ω_c，因此可将操纵方程简化为because

much smaller than

The resulting moment, ω is much smaller than Ω _c , so the steering equation can be simplified to

再依据执行机构的配置方案，可得敏捷卫星的执行机构的的输入如下式：According to the configuration of the dual control torque gyroscope, the frame angles and frame angular velocities of the two control torque gyroscopes are equivalently reversed, so that the

Then according to the configuration scheme of the actuator, the input of the actuator of the agile satellite can be obtained as follows:

u _c is the output command signal of the attitude controller. By inputting the command signal to the compound attitude actuator, the agile satellite can maneuver according to the desired attitude.

Below in conjunction with the Simulink simulation of Matlab, the specific embodiments of the present invention are described in more detail,

Let the width D required to be detected by the agile satellite be 1000km, the relative motion speed of the satellite to the ground V _d = 6.8km/s, the width of the sub-satellite point L × W = 117 × 21km, the corresponding field of view θ _L × θ _W = 13.41 ×2.41, L=117km is the direction along the track. Therefore, the swing angle θ _D = Dθ _W /W = 90° in the orbiting direction needs to be completed when the entire star imaging width is D. To achieve seamless continuous imaging, the satellite is swung back to the initial position, and the longest time limit of one pendulum is t _max =L/V _d =17.2s.

In order to verify the effectiveness of the present invention when the agile satellite frequently performs large-angle reciprocating maneuvering tasks, the attitude maneuvering command is: a roll command executed from -45°→45° from 0s, and then executed when the satellite roll angle reaches 45° 45°→-45° roll command, and repeat the process. Based on this, various simulation parameters can be set as:

1. Satellite parameters

Moment of inertia matrix: J=[7,0,0;0,12,0;0,0,12]kgm ² ,

Initial Euler angles:

Initial angular velocity: ω ₀ =[0,0,0] ^T ,

Expected Euler angles:

Pendulum period: T=16s,

2. Control torque gyro parameters

The moment of inertia in the direction of the rotor axis:

Rotor speed:

Initial frame corners:

Maximum frame angular velocity:

3. Reaction flywheel parameters

Reaction flywheel axial moment of inertia:

Reaction flywheel initial speed:

Reaction flywheel maximum speed:

Maximum angular acceleration of the reaction flywheel:

4. Controller parameters

Angle error gain: k ₁ =100,

Angular velocity error gain: k ₂ =100.

According to the above simulation parameters, the Simulink simulation process is performed, and the simulation time is set to 160s, that is, ten pendulum cycles.

The simulation results are as follows:

As shown in Figure 4, the satellite completed the roll maneuver from -45°→45°, and then from 45°→-45° within 16s, and the satellite three-axis attitude angle tracking accuracy was better than 0.1°.

As shown in Figure 5, the satellite completed the acceleration process from 0°/s→17.7°/s within 4s, and completed the process of first decelerating and then accelerating from 17.7°/s→0→-17.7°/s within 8s .

As shown in Figure 6, the initial frame angle is 90°, which gradually increases to 121° and then decreases to 59°. When one pendulum cycle is completed, the frame angle returns to the vicinity of the initial frame angle, avoiding the occurrence of the singularity of the frame angle.

The above simulation results of Fig. 4 to Fig. 6 verify that the attitude control method proposed by the present invention can realize the rapid reciprocating swing of the roll angle and large angle, and can meet the requirements of ultra-large-width seamless continuous maneuvering imaging of ground targets, and has the ability to maneuver. It has the characteristics of strong ability, good speed, and avoiding the singularity of frame corners.

It should be emphasized that the above are only preferred embodiments of the present invention, and are not intended to limit the present invention in any form. Any simple modifications, equivalent changes and modifications made to the above embodiments according to the technical essence of the present invention are Still belong to the scope of the technical solution of the present invention.

Claims (5)

1. an executive mechanism configuration method of agile satellite attitude reciprocating fast swing, it is characterized in that: described executive mechanism is arranged to carry the compound attitude control executive mechanism that is mounted on the agile satellite platform, comprises control moment gyroscope and reaction flywheel, described The number of control moment gyroscopes and reaction flywheels is set to two, and the configuration method is as follows: A rectangular three-dimensional coordinate system Oxyz is constructed with the agile satellite platform as the center, wherein the two reaction flywheels are respectively installed on the y and z axes, and the positive directions of their rotor rotations are respectively directed towards the negative directions of the y and z axes, and the two control torques The installation direction of the frame axis of the gyro is parallel to the y-axis, and constitutes a dual-control torque gyro structure. The output control torque acts in the x-axis direction, where the x-axis is the roll axis of the agile satellite, and the agile satellite has the ability to quickly maneuver in the x-axis direction. , by controlling the moment gyro to control the agile satellite to reciprocate quickly around the roll axis, and the reaction flywheel to control the attitude stability of the agile satellite in the y and z directions. 2 . The method for configuring an actuator for reciprocating and fast swinging of an agile satellite attitude according to claim 1 , wherein the two control torque gyroscopes have the same structural parameters. 3 . 3. The implementing mechanism configuration method of the agile satellite attitude reciprocating fast swing according to claim 1 is characterized in that: the frame axes of the two described control torque gyroscope structures are parallel to each other, and the rotational speeds are equal but opposite, and a frame axis rotational angular velocity The direction of the vector is along the positive direction of the y-axis, and the direction of the rotational angular velocity vector of the other frame axis is along the negative direction of the y-axis. 4 . The method for configuring an actuator for reciprocating and rapidly swinging an agile satellite attitude according to claim 1 , wherein the frame angles and frame angular velocities of the two control torque gyro structures are equivalently reversed. 5 . 5. The method for configuring the actuator for the rapid reciprocating swing of the agile satellite attitude according to any one of claims 1 to 4, characterized in that: comprising: Constructing a satellite attitude dynamics model based on the compound attitude control actuator, the specific steps are to use the modeling method of vector mechanics to first establish the angular momentum equation of the whole satellite including the control moment gyroscope and the reaction flywheel about its system center of mass, Then, the relationship between the system angular momentum change and the moment is obtained by using the angular momentum theorem, and finally the above equation is projected to the satellite system, and then the satellite attitude dynamics model is established; Constructing the satellite attitude kinematics model based on the compound attitude control actuator, the specific steps are to use the quaternion attitude parameter representation method, first establish the attitude kinematic equation of the satellite under this system, and then according to the expected attitude quaternion and current The relationship between attitude quaternions and establishing satellite attitude kinematic error equation; Using the constructed agile satellite attitude dynamics and kinematics model, an attitude controller based on the compound attitude control actuator is designed, which can control the agile satellite to maneuver according to the desired attitude. The specific steps are to first construct the Lyapunov function of the system for the attitude error system. And take its derivation, and then use Lyapunov theorem and LaSalle invariant set principle to design the attitude controller, output the command signal through the attitude controller, input the command signal into the compound attitude control actuator, and then make the compound attitude control actuator control The satellite maneuvers according to the desired attitude.

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