CN115817856B - Method and device for controlling stable attitude of satellite to solar spin based on pure magnetic control mode - Google Patents

Abstract

The invention proposes a method and device for controlling a satellite's spin-stabilized attitude towards the sun based on a pure magnetic control method, which belongs to the technical field of spacecraft attitude control. Wherein, the method includes: calculating the estimated value of the three-axis angular rate of the satellite, and calculating the target output torque of each control stage of the satellite's rotation to the sun according to the estimated value of the satellite's three-axis angular rate, wherein, in the acquisition phase of the sun, In the spin-up phase and the sun-spin stabilization phase, when the sun is within the measurement field of view of the sun sensor, the sun vector measured by the sun sensor is considered in the target output torque; according to the target output torque and the measured value of the geomagnetic vector, the calculation The target output magnetic moment of the three-axis magnetic torque device is converted into control commands to realize satellite attitude control. The invention fully considers the effects of factors such as the shadow area of the earth, the offset installation of the sun sensor and the solar sail, the field of view of the sun sensor, and the environmental disturbance moment, and can control the satellite from any initial state to finally achieve spin stability towards the sun.

Description

Method and device for satellite-sun spin-stabilized attitude control based on pure magnetron control

technical field

The invention belongs to the technical field of spacecraft attitude control, and in particular relates to a method and device for controlling a satellite-sun spin-stabilized attitude based on a pure magnetic control method.

Background technique

Generally speaking, in order to ensure the energy supply of the on-board system, the satellite is oriented toward the sun most of the time in orbit. In the state of orientation to the sun, the satellite relies on the attitude control system to continuously adjust the attitude of the satellite, so that the solar panels are always aligned with the sun vector. The state of orientation toward the sun is crucial to the energy supply on the planet. Whether a stable and reliable attitude toward the sun can be established directly determines the success or failure of the satellite mission. Therefore, it is of great significance to study how to use the smallest system to achieve directional attitude control toward the sun, especially for satellites with limited configuration and device failures.

In 2007, Luo Junhong and other researchers designed a spin-stabilized attitude control method for solar sail spacecraft, discussed the way to obtain the spin velocity, and carried out a simulation analysis of the attitude control characteristics of the solar sail spacecraft. The simulation results show that Solar sail spacecraft can achieve single-axis pointing by spinning. In 2019, Xia Xiwang and other researchers chose the sun sensor and magnetometer as the attitude sensor, and the magnetic torque device as the attitude actuator, and proposed a pure magnetron spin alignment method and its correction method, realizing the satellite alignment. Spin stabilized attitude control. In 2020, Liu Shanwu and other researchers proposed a sun capture attitude control method based on a minimalist system configuration for satellites operating in morning and evening sun-synchronous orbits. This method makes full use of the characteristics of satellite orbits and uses magnetometers and magnetic torque devices As the attitude sensor and the attitude actuator respectively, the pitch axis spin-up method is used to realize the spin-stabilized attitude control of the satellite to the sun.

However, in the existing sun-oriented attitude control methods, the influence of factors such as the shadow area of the satellite, the offset installation of the sun sensor and the solar panel, the field of view of the sun sensor, and the environmental disturbance moment during the actual on-orbit operation of the satellite are not fully considered. , the scope of application is limited, and it is difficult to apply to engineering practice. For most of the satellites, they will inevitably enter the earth shadow area during their in-orbit operation. At this time, the output of the sun sensor is invalid. This requires the design of a new attitude control method to meet the needs of satellites in different orbits. Directional attitude control requirements.

Contents of the invention

The purpose of the present invention is to overcome the deficiencies of the prior art, and propose a method and device for controlling the satellite's spin-stabilized attitude towards the sun based on a pure magnetic control method. The invention is based on a pure magnetic control method, and can control the satellite from any initial state under the condition of fully considering the shadow area, the offset installation of the sun sensor and the solar sail, the field of view of the sun sensor, and the environmental disturbance moment. Finally, the spin stability to the sun is achieved, which can effectively ensure the smooth implementation of satellite equipment debugging, energy acquisition and other flight tasks. It has the advantages of low cost, low power consumption, simple process, high reliability, and good stability. It is suitable for most Earth orbit satellite.

The embodiment of the first aspect of the present invention proposes a method for controlling the satellite's spin-stabilized attitude towards the sun based on a pure magnetron method, including:

Compute estimates of the satellite's triaxial angular rate;

According to the estimated value of the three-axis angular rate of the satellite, the target output torques of the satellite's rotation to the sun control stages are calculated respectively. phase and the sun-spin stabilization phase; wherein, during the sun-gathering phase, the spin-up phase, and the sun-spin stabilization phase, when the sun is within the measurement field of view of the sun sensor of the satellite , the sun vector measured by the sun sensor is considered in the target output torque;

Calculate the target output magnetic moment of the three-axis magnetotorque of the satellite according to the target output moment and the geomagnetic vector measurement value measured by the three-axis magnetometer of the satellite;

The target output magnetic moment is converted into a control command of the three-axis magnetic torque device, so as to drive the three-axis magnetic torque device to perform attitude control of the satellite.

In a specific embodiment of the present invention, the estimated value calculation method of the satellite three-axis angular velocity is as follows:

(1) Build the state vector

为卫星本体坐标系中表示的卫星三轴角速率；ω_x、ω_y、ω_z分别为/>

在卫星本体坐标系O_bX_bY_bZ_b的O_bX_b、O_bY_b、O_bZ_b轴向上的分量；in,

is the three-axis angular velocity of the satellite expressed in the satellite body coordinate system; ω _x , ω _y , ω _z are respectively />

Components in the O _{b X b} , O _b Y _{b ,} O _{b Z b axial} directions of the satellite body coordinate system O _b X b Y b Z b;

(2) Construct state equation and state transition matrix;

The satellite attitude dynamic equation is constructed as follows:

为卫星的转动惯量矩阵，当卫星本体坐标系O_bX_bY_bZ_b为质心惯性主轴坐标系时，矩阵J为对角矩阵，矩阵J的对角元素J_x、J_y、J_z分别代表卫星绕O_bX_b、O_bY_b、O_bZ_b轴方向旋转的转动惯量；/>

为三轴磁力矩器输出的控制力矩；in,

is the moment of inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b is the center of mass inertial axis coordinate system, the matrix J is a diagonal matrix, and the diagonal elements J _x , J _y , J _z of the matrix J are respectively Represents the moment of inertia of the satellite rotating around the O _b X _b , O _b Y _b , O _b Z _b axes; />

is the control torque output by the three-axis magnetic torque device;

The attitude dynamic equation is the state equation, and the attitude dynamic equation is linearized to obtain the state transition matrix

Among them, ΔT is the running cycle;

(3) Construct measurement equation and measurement matrix;

其中，

为当前时刻的地磁矢量测量值，/>

分别为当前时刻地磁矢量测量值在卫星本体坐标系O_bX_b、O_bY_b、O_bZ_b轴向上的分量；/>

为前一时刻地磁矢量测量值，/>

分别为前一时刻地磁矢量测量值在卫星本体坐标系O_bX_b、O_bY_b、O_bZ_b轴向上的分量；The measured values of the geomagnetic vector obtained by two consecutive measurements by the three-axis magnetometer are denoted as B ^b and

in,

is the measured value of the geomagnetic vector at the current moment, />

Respectively, the geomagnetic vector measured values at the current moment are in the satellite body coordinate system O _b X _b , O _b Y _b , O _b Z _b axial components; />

is the measured value of the geomagnetic vector at the previous moment, />

are the components of the geomagnetic vector measurement value at the previous moment in the satellite body coordinate system O _b X _b , O _b Y _b , O _b Z _b axis;

如下：Build the observation vector

as follows:

为表示时间段ΔT内卫星本体坐标系转动角度的向量，Θ_Δ所对应的方向余弦矩阵/>

近似表示为：make

is a vector representing the rotation angle of the satellite body coordinate system within the time period ΔT, the direction cosine matrix corresponding to Θ _Δ />

Approximately expressed as:

为向量/>

的叉乘矩阵，表达式如下：in,

is the vector />

The cross product matrix of , the expression is as follows:

Then construct the measurement equation as follows:

为误差项；in,

is the error term;

为：According to the measurement equation, the measurement matrix is obtained

for:

(4) According to the state transition matrix obtained in step (2) and the measurement matrix obtained in step (3), use the Kalman filter method to calculate the estimated value of the three-axis angular rate of the satellite

In a specific embodiment of the present invention, the method also includes:

During the initial derotation phase, or when the sun enters the sun sensor measurement field of view of the satellite during the sun acquisition phase, or when the satellite is in the shadow area during the sun acquisition phase , the calculation expression of the target output torque is as follows:

为卫星三轴角速率的估计值，K₁为正的第一控制系数；in,

is the estimated value of the three-axis angular velocity of the satellite, and K ₁ is the positive first control coefficient;

Calculate the target output magnetic moment M of the three-axis magnetic torque device:

Among them, B ^b is the measured value of the geomagnetic vector at the current moment; α ₁ is the angle between the vector B ^b and T _c1 .

In a specific embodiment of the present invention, the method also includes:

In the initial derotation phase, if the angular rate value of the three-axis angular rate of the satellite is always less than or equal to the set first angular rate threshold within the time period greater than or equal to the set first time threshold, the initial derotation phase ends, and the The satellite enters the stage of capturing the sun.

In a specific embodiment of the present invention, the method also includes:

In the phase of capturing the sun and when the sun is within the measurement field of view of the sun sensor of the satellite, the calculation expression of the target output torque is as follows:

为卫星三轴角速率的估计值，K₂、K₃、K₄分别为正的第二、三、四控制系数；

为太阳敏感器光轴方向单位矢量；S^b为当前时刻太阳敏感器测量得到的太阳矢量，/>

为前一时刻太阳敏感器测得的太阳矢量；/>

为卫星的转动惯量矩阵，当卫星本体坐标系O_bX_bY_bZ_b为质心惯性主轴坐标系时，矩阵J为对角矩阵，矩阵J的对角元素J_x、J_y、J_z分别代表卫星绕O_bX_b、O_bY_b、O_bZ_b轴方向旋转的转动惯量；in,

is the estimated value of the three-axis angular velocity of the satellite, and K ₂ , K ₃ , K ₄ are the positive second, third, and fourth control coefficients respectively;

is the unit vector of the optical axis direction of the sun sensor; S ^b is the sun vector measured by the sun sensor at the current moment, />

The sun vector measured by the sun sensor at the previous moment; />

is the moment of inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b is the center of mass inertial axis coordinate system, the matrix J is a diagonal matrix, and the diagonal elements J _x , J _y , J _z of the matrix J are respectively Represents the moment of inertia of the satellite rotating around the O _b X _b , O _b Y _b , O _b Z _b axes;

Calculate the target output magnetic moment M of the three-axis magnetic torque device:

Among them, B ^b is the measured value of the geomagnetic vector at the current moment; α ₂ is the angle between the vector B ^b and T _c2 .

In a specific embodiment of the present invention, the method also includes:

与当前时刻太阳敏感器测量得到的太阳矢量S^b的夹角在大于等于第二时间阈值的时间内始终小于等于设定的第一角度阈值，则对日捕获阶段结束，所述卫星进入起旋阶段。In the sun capture phase, if the sun sensor optical axis direction unit vector

The included angle with the sun vector S ^b measured by the sun sensor at the current moment is always less than or equal to the set first angle threshold within the time greater than or equal to the second time threshold, then the phase of capturing the sun ends, and the satellite enters the spin stage.

In a specific embodiment of the present invention, the method also includes:

In the spin-up phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the calculation expression of the target output torque is as follows:

为卫星三轴角速率的估计值，K₅、K₆、K₇、K₈分别为正的第五、六、七、八控制系数；/>

为设定的需对准太阳矢量的卫星星体面法向单位矢量；ω_spin为设定的自旋角速率值；S^b为当前时刻太阳敏感器测量得到的太阳矢量，/>

为前一时刻太阳敏感器测得的太阳矢量；/>

为卫星的转动惯量矩阵，当卫星本体坐标系O_bX_bY_bZ_b为质心惯性主轴坐标系时，矩阵J为对角矩阵，矩阵J的对角元素J_x、J_y、J_z分别代表卫星绕O_bX_b、O_bY_b、O_bZ_b轴方向旋转的转动惯量；in,

is the estimated value of the three-axis angular rate of the satellite, K ₅ , K ₆ , K ₇ , and K ₈ are the positive fifth, sixth, seventh, and eighth control coefficients respectively; />

is the set normal unit vector of the satellite star surface that needs to be aligned with the sun vector; ω _spin is the set spin angular rate value; S ^b is the sun vector measured by the sun sensor at the current moment, />

The sun vector measured by the sun sensor at the previous moment; />

is the moment of inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b is the center of mass inertial axis coordinate system, the matrix J is a diagonal matrix, and the diagonal elements J _x , J _y , J _z of the matrix J are respectively Represents the moment of inertia of the satellite rotating around the O _b X _b , O _b Y _b , O _b Z _b axes;

Calculate the target output magnetic moment M of the three-axis magnetic torque device:

Among them, B ^b is the measured value of the geomagnetic vector at the current moment; α ₃ is the angle between the vector B ^b and T _c3 .

In a specific embodiment of the present invention, the method also includes:

In the spinning phase, if the satellite flies into the shadow area, the satellite will return to the sun capture phase;

与当前时刻太阳敏感器测量得到的太阳矢量S^b间的夹角在大于等于设定的第三时间阈值的时间内始终小于等于设定的第二角度阈值，且卫星三轴角速率估计值/>

与设定的卫星自旋角速率向量/>

间的偏差在相同的所述时间内始终小于等于设定的第二角速率阈值，则起旋阶段结束，所述卫星进入对日自旋稳定阶段。If the set satellite star surface normal vector that needs to be aligned with the sun vector

The angle between the sun vector S ^b measured by the sun sensor at the current moment is always less than or equal to the set second angle threshold within the time greater than or equal to the set third time threshold, and the estimated value of the satellite three-axis angular rate / >

with the set satellite spin angular rate vector />

If the deviation between them is always less than or equal to the set second angular rate threshold within the same said time, the spin-up phase ends, and the satellite enters the sun-spin stabilization phase.

In a specific embodiment of the present invention, the method also includes:

During the spin stabilization phase against the sun, calculate the target output torque;

Wherein, when the satellite is in the sunshine area, the calculation expression of the target output torque is as follows:

为卫星三轴角速率的估计值，K₅、K₆、K₇、K₈分别为正的第五、六、七、八控制系数；/>

为设定的需对准太阳矢量的卫星星体面法向单位矢量；ω_spin为设定的自旋角速率值；S^b为当前时刻太阳敏感器测量得到的太阳矢量，/>

为前一时刻太阳敏感器测得的太阳矢量；/>

为卫星的转动惯量矩阵，当卫星本体坐标系O_bX_bY_bZ_b为质心惯性主轴坐标系时，矩阵J为对角矩阵，矩阵J的对角元素J_x、J_y、J_z分别代表卫星绕O_bX_b、O_bY_b、O_bZ_b轴方向旋转的转动惯量；in,

is the estimated value of the three-axis angular rate of the satellite, K ₅ , K ₆ , K ₇ , and K ₈ are the positive fifth, sixth, seventh, and eighth control coefficients respectively; />

is the set normal unit vector of the satellite star surface that needs to be aligned with the sun vector; ω _spin is the set spin angular rate value; S ^b is the sun vector measured by the sun sensor at the current moment, />

The sun vector measured by the sun sensor at the previous moment; />

is the moment of inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b is the center of mass inertial axis coordinate system, the matrix J is a diagonal matrix, and the diagonal elements J _x , J _y , J _z of the matrix J are respectively Represents the moment of inertia of the satellite rotating around the O _b X _b , O _b Y _b , O _b Z _b axes;

When the satellite is in the shadow area, the calculation expression of the target output torque is as follows:

Wherein, K ₉ is positive ninth control coefficient;

Calculate the target output magnetic moment M of the three-axis magnetic torque device:

Among them, B ^b is the measured value of the geomagnetic vector at the current moment; α ₄ is the angle between the vector B ^b and T _c4 .

The embodiment of the second aspect of the present invention proposes a satellite-to-sun spin stabilization attitude control device based on a pure magnetic control method, including:

The triaxial angular rate estimation module is used to calculate the estimated value of the satellite triaxial angular rate;

The target output torque calculation module is used to calculate the target output torque of each control stage of the satellite's rotation to the sun according to the estimated value of the three-axis angular rate of the satellite, and the control stages of the rotation to the sun include: the initial derotation stage , the sun capture phase, the spinning phase and the sun spin stabilization phase; wherein, in the sun capture phase, the spin phase and the sun spin stabilization phase, when the sun is located at the sun of the satellite When within the measurement field of view of the sensor, the sun vector measured by the sun sensor is considered in the target output torque;

The target output magnetic moment calculation module calculates the target output magnetic moment of the three-axis magnetic torque device of the satellite according to the target output torque and the geomagnetic vector measurement value measured by the three-axis magnetometer of the satellite;

The attitude control command calculation module is used to convert the target output magnetic moment into a control command of the three-axis magnetic torque device, so as to drive the three-axis magnetic torque device to perform attitude control of the satellite.

The embodiment of the third aspect of the present invention proposes an electronic device, including:

at least one processor; and, a memory communicatively coupled to the at least one processor;

Wherein, the memory stores instructions that can be executed by the at least one processor, and the instructions are configured to execute the above-mentioned method for controlling the satellite's spin-stabilized attitude toward the sun based on a purely magnetic control method.

The embodiment of the fourth aspect of the present invention proposes a computer-readable storage medium, the computer-readable storage medium stores computer instructions, and the computer instructions are used to enable the computer to execute the above-mentioned pure magnetron-based satellite alignment Spin stabilized attitude control method

Features and beneficial effects of the present invention are:

The present invention is based on a pure magnetic control method, and under the condition of fully considering the influence of factors such as the shadow area of the earth, the offset installation of the sun sensor and the solar sail, the field of view of the sun sensor, and the environmental disturbance torque, the three-axis magnetometer and The sun sensor is used as an attitude sensitive device, and the three-axis magnetic torquer is used as an attitude actuator, which can control the satellite from any initial state to finally achieve spin stabilization towards the sun.

The present invention uses the measured value of the three-axis magnetometer as an input solution to obtain the three-axis angular rate of the satellite, and designs a control stage including initial derotation, capture to the sun, spin-up, and spin stabilization to the sun in combination with the measured value of the sun sensor. Attitude control scheme, and use the three-axis magnetic torque device to realize the control torque output. The invention can effectively ensure the smooth implementation of flight tasks such as satellite equipment debugging and energy acquisition, has the advantages of low cost, low power consumption, simple process, high reliability, good stability, etc., and can meet the sun-oriented attitude control of satellites in different orbits need.

Description of drawings

FIG. 1 is an overall flow chart of a method for controlling a satellite's sun-spin-stabilized attitude based on pure magnetron control according to an embodiment of the present invention.

Detailed ways

The present invention proposes a satellite spin-stabilized attitude control method and device based on a pure magnetic control method, which will be further described in detail below in conjunction with the accompanying drawings and specific embodiments.

The embodiment of the first aspect of the present invention proposes a method for controlling the satellite's spin-stabilized attitude towards the sun based on a pure magnetron method, including:

Compute estimates of the satellite's triaxial angular rate;

According to the estimated value of the three-axis angular rate of the satellite, the target output torques of the satellite's rotation to the sun control stages are calculated respectively. phase and the sun-spin stabilization phase; wherein, during the sun-gathering phase, the spin-up phase, and the sun-spin stabilization phase, when the sun is within the measurement field of view of the sun sensor of the satellite , the sun vector measured by the sun sensor is considered in the target output torque;

Calculate the target output magnetic moment of the three-axis magnetic torque device according to the target output torque and the geomagnetic vector measurement value obtained by the three-axis magnetometer measurement of the satellite;

The target output magnetic moment is converted into a control command of the three-axis magnetic torque device, so as to drive the three-axis magnetic torque device to perform attitude control of the satellite.

In a specific embodiment of the present invention, the satellite is set to operate in a sun-synchronous orbit with an orbital height of 545 km and a descending node local time of 10:30. The three-axis moment of inertia of the satellite is [1.37 1.69 2.05]kg·m ² , the remanent magnetism of the satellite is 0.012A·m ² , and the reflection coefficient of the satellite surface is 0.6. Set the measurement error of the three-axis magnetometer to 3×10 ^-4 μT; set the included angles between the optical axis of the sun sensor and the O _b X _b , O _b Y _b , and O _b Z _b axes of the satellite to be 69.2952° , 150°, 110.7048°, the field of view of the sun sensor is 90°×90°, and the two-axis accuracy is 0.1°; the maximum output magnetic moment of the three-axis magnetic torque device is set to [2.523.21 2.52]A·m ² . Set the initial three-axis angular rate of the satellite to [-1.5 -1.5 -1.5]°/s, and the three-axis Euler angles of the initial attitude to [160 20 60]°.

In this embodiment, the overall flow of the satellite-to-sun spin-stabilized attitude control method based on a pure magnetic control method is shown in Figure 1, and includes the following steps:

(1) Calculate the estimated value of the satellite's three-axis angular velocity;

In this embodiment, an Extended Kalman Filter method (ExtendedKalmanFilter, EKF) is used, and the three-axis magnetometer measurement value is used as an input to estimate the three-axis angular rate of the satellite.

(1-1) Build a state vector;

定义如下：In this embodiment, the state vector

It is defined as follows:

为卫星本体坐标系中表示的卫星三轴角速率；ω_x、ω_y、ω_z分别为/>

在O_bX_bY_bZ_b的O_bX_b、O_bY_b、O_bZ_b轴向上的分量。in,

is the three-axis angular velocity of the satellite expressed in the satellite body coordinate system; ω _x , ω _y , ω _z are respectively />

Components in the O _{b X b} , O _b Y _{b ,} O _b Z _b axial directions _of O _{b X b} Y b Z b.

(1-2) Construct state equation and state transition matrix;

In this embodiment, since the requirements for the estimation accuracy of the angular rate are not high, factors such as the space interference moment in the satellite attitude dynamic equation are ignored, and the satellite attitude dynamic equation is constructed as follows:

为卫星的转动惯量矩阵，当卫星本体坐标系O_bX_bY_bZ_b为质心惯性主轴坐标系时，矩阵J为对角矩阵，其对角元素J_x、J_y、J_z分别代表卫星绕O_bX_b、O_bY_b、O_bZ_b轴方向旋转的转动惯量；/>

为三轴磁力矩器输出的控制力矩。in,

is the moment of inertia matrix of the satellite. When the satellite body coordinate system O _b X _b Y _b Z _b is the center of mass inertial axis coordinate system, the matrix J is a diagonal matrix, and its diagonal elements J _x , J _y , and J _z respectively represent satellite Moment of inertia of rotation around O _b X _b , O _b Y _b , O _b Z _b axes; />

is the control torque output by the three-axis magnetic torque device.

The above attitude dynamics equation is the state equation, which can be linearized to obtain the state transition matrix

Among them, ΔT is the set operation cycle.

(1-3) Construct measurement equation and measurement matrix;

其中，

为当前时刻的地磁矢量测量值，/>

分别为当前时刻地磁矢量测量值在卫星本体坐标系O_bX_b、O_bY_b、O_bZ_b轴向上的分量；/>

为前一时刻地磁矢量测量值，/>

分别为前一时刻地磁矢量测量值在卫星本体坐标系O_bX_b、O_bY_b、O_bZ_b轴向上的分量。The measured values of the geomagnetic vector obtained by two consecutive measurements by the three-axis magnetometer are denoted as B ^b and

in,

is the measured value of the geomagnetic vector at the current moment, />

Respectively, the geomagnetic vector measured values at the current moment are in the satellite body coordinate system O _b X _b , O _b Y _b , O _b Z _b axial components; />

is the measured value of the geomagnetic vector at the previous moment, />

are the components of the geomagnetic vector measurement value at the previous moment in the satellite body coordinate system O _b X _b , O _b Y _b , and O _b Z _b axes, respectively.

为三轴磁强计相邻两次测量值的差值：Build the observation vector

is the difference between two adjacent measurements of the three-axis magnetometer:

为表示时间段ΔT内卫星本体坐标系转动角度的向量，则在小角度假设下，Θ_Δ所对应的方向余弦矩阵/>

可近似表示为：In the case of relatively small ΔT, the change of the geomagnetic field vector in a short period of time can be ignored, and it is approximately considered that the difference between the two measured values of the magnetometer before and after is only caused by the change of the satellite attitude. Let the three-axis angular velocity of the satellite be

is a vector representing the rotation angle of the satellite body coordinate system within the time period ΔT, then under the assumption of a small angle, the direction cosine matrix corresponding to Θ _Δ />

It can be approximated as:

为向量/>

的叉乘矩阵，表达式如下：in,

is the vector />

The cross product matrix of , the expression is as follows:

Further, construct the measurement equation as follows:

为误差项。in,

is the error term.

为：According to the measurement equation, the measurement matrix can be obtained

for:

(1-4) Kalman filter;

用于姿态控制器各阶段的控制。According to the state transition matrix obtained in step (1-2) and the measurement matrix obtained in step (1-3), according to the general algorithm flow of the extended Kalman filter, the estimated value of the three-axis angular velocity of the satellite is obtained by solving

It is used for the control of each stage of the attitude controller.

(2) Use the satellite three-axis angular velocity estimate obtained in step 1) to perform attitude control on each control stage of the satellite's daily rotation;

Among them, the various control stages of the satellite's spin to the sun include: the initial despin phase, the capture phase to the sun, the spin-up phase, and the spin stabilization phase to the sun; the specific steps are as follows:

(2-1) initial racemization stage;

After the satellite is separated from the vehicle or the attitude control is not performed for a long time, the star may have initial angular momentum, and the initial derotation needs to be carried out first.

设计如下控制律，得到初始消旋阶段三轴磁力矩器的目标输出力矩T_c1：(2-1-1) Estimated value of satellite three-axis angular rate using step (1)

Design the following control law to obtain the target output torque T _c1 of the three-axis magnetic torque device in the initial derotation stage:

Wherein, K ₁ is a positive first control coefficient (the reference value of this embodiment is 0.01).

In the expression of the control torque T _c1 above, the first term on the right side of the equation is used to damp the three-axis angular rate, and the second term is used to eliminate the satellite three-axis angular rate coupling effect.

(2-1-2) Using the measured geomagnetic vector value B ^b at the current moment obtained from the three-axis magnetometer measurement, combined with the target output torque T _c1 obtained in step (2-1-1), calculate the target of the three-axis magnetic torque device Output magnetic moment M:

Among them, α ₁ is the angle between vector B ^b and T _c1 .

(2-1-3) With the three-axis magnetic torque device as the control component, the target output magnetic torque calculated in step (2-1-2) is converted into the control command of the three-axis magnetic torque device to drive the three-axis magnetic torque device for satellite attitude control.

(2-1-4) In the initial derotation stage, if the angular rate of the satellite's three-axis angular rate is greater than or equal to the set first time threshold (the reference value of this embodiment is 100s), the angular rate value is always less than or equal to the set If the first angular velocity threshold is determined (the reference value of this embodiment is 0.2°/s), then the initial derotation phase ends, and the satellite enters the step (2-2) acquisition phase for the sun.

(2-2) The stage of capturing the sun;

After the satellite completes the initial derotation, it enters the phase of capturing the sun, using the sun sensor to capture the sun vector, and through the control, the unit vector in the direction of the sun sensor optical axis coincides with the sun vector.

(2-2-1) Before the sun enters the measurement field of view of the sun sensor, considering that the satellite that only uses the three-axis magnetic torquer as the attitude actuator can only output the control torque perpendicular to the satellite’s local geomagnetic vector, the satellite’s attitude The maneuverability is poor, so the control law of the initial derotation stage is still used to search the sun slowly until the sun enters the field of view of the sun sensor.

(2-2-2) When the sun enters the measurement field of view of the sun sensor, the sun sensor can obtain the sun vector S ^b through measurement. Design the following control law to obtain the target output torque T _c2 of the three-axis magnetotorque when the sun is located in the measurement field of view of the sun sensor of the satellite during the sun capture phase, so that the optical axis of the sun sensor is aligned with the sun vector direction :

为太阳敏感器光轴方向单位矢量，由太阳敏感器在星体上的安装角度决定；/>

为前一时刻太阳敏感器测得的太阳矢量。上述控制力矩T_c2的表达式中，等式右边第一项用于控制/>

与S^b重合，第二项和第三项用于阻尼卫星三轴角速率中与S^b垂直的分量，第四项用于消除卫星三轴角速率耦合效应。本发明中假定太阳敏感器的视场范围小于等于90°。Wherein, K ₂ , K ₃ , and K ₄ are positive second, third, and fourth control coefficients respectively (the reference values in this embodiment are 0.0005, 0.001, and 0.02, respectively);

is the unit vector of the optical axis direction of the sun sensor, determined by the installation angle of the sun sensor on the star; />

is the sun vector measured by the sun sensor at the previous moment. In the expression of the above control torque T _c2 , the first item on the right side of the equation is used to control

Coinciding with S ^b , the second and third items are used to damp the component perpendicular to S ^b in the satellite triaxial angular rate, and the fourth item is used to eliminate the coupling effect of the satellite triaxial angular rate. In the present invention, it is assumed that the field of view of the sun sensor is less than or equal to 90°.

和向量S^b的夹角大小范围，基于自定义的正相关函数、阶跃函数、限幅函数等对上式中的第一项进行适应性修改。It should be noted that if the field of view of the sun sensor is greater than 90°, the vector

The size range of the included angle with the vector S ^b , based on the self-defined positive correlation function, step function, clipping function, etc., the first item in the above formula is adaptively modified.

In this embodiment, the characteristic of the control torque T _c2 is that the control amount for damping the component perpendicular to S ^b in the satellite triaxial angular rate and the control amount for eliminating the coupling effect of the satellite triaxial angular rate are designed, so that the satellite can In the case of underdrive, align the optical axis of the sun sensor with the sun vector direction smoothly and accurately.

(2-2-3) Using the measured geomagnetic vector value B ^b at the current moment obtained from the three-axis magnetometer measurement, combined with the target output torque T _c2 obtained in step (2-2-2), calculate the target of the three-axis magnetotorque Output magnetic moment M:

Among them, α ₂ is the angle between vector B ^b and T _c2 .

(2-2-4) With the three-axis magnetic torque device as the control component, the target output magnetic moment calculated in step (2-2-3) is converted into the control command of the three-axis magnetic torque device to drive the three-axis magnetic torque Torquer for satellite attitude control.

与其测得的太阳矢量S^b的夹角在大于等于第二时间阈值(本实施例参考取值为25s)的时间内始终小于等于设定的第一角度阈值(本实施例参考取值为10°)，则对日捕获阶段结束，卫星进入步骤(2-3)起旋阶段；否则，按照步骤(2-2-2)-(2-2-4)继续进行对日捕获阶段的姿态控制。(2-2-5) In the phase of capturing the sun, if the unit vector in the direction of the sun sensor optical axis

The included angle of the sun vector S ^b measured with it is always less than or equal to the set first angle threshold (the reference value of this embodiment is 10 s) within the time greater than or equal to the second time threshold (the reference value of this embodiment is 25s). °), the phase of capturing the sun is over, and the satellite enters the spin-up phase of step (2-3); otherwise, continue the attitude control of the phase of capturing the sun according to steps (2-2-2)-(2-2-4) .

Wherein, during the attitude control process of the sun acquisition phase, if the satellite flies into the shadow area of the earth, return to step (2-2-1).

(2-3) Spinning stage;

After the satellite captures the sun, it enters the spinning stage. On the basis of the sun sensor optical axis being roughly aligned with the sun vector, the designated satellite star surface normal vector is further coincided with the sun vector, and the satellite is controlled to orbit the star surface in a normal direction. Swirl toward the vector.

(2-3-1) In the spin-up phase and when the sun is located in the measurement field of view of the sun sensor of the satellite, the following control law is designed to obtain the target output torque T _c3 of the three-axis magnetic torque device in the spin-up phase, Make the specified satellite's face normal vector coincide with the sun's vector and orbit around the star's face normal vector:

为指定的需要对准太阳矢量的卫星星体面法向单位矢量；ω_spin为设定的自旋角速率值(本实施例参考取值为1°/s)。上述控制力矩T_c3的表达式中，等式右边第一项用于控制/>

与S^b重合，第二项和第三项用于阻尼卫星三轴角速率中与S^b垂直的分量，第四项用于起旋，第五项用于消除卫星三轴角速率耦合效应。Among them, K ₅ , K ₆ , K ₇ , and K ₈ are positive fifth, sixth, seventh, and eighth control coefficients respectively (the reference values in this embodiment are 0.0004, 0.0008, 0.01, and 0.004, respectively);

ω _spin is the set spin angular rate value (the reference value of this embodiment is 1°/s). In the above expression of control torque T _c3 , the first item on the right side of the equation is used to control

Coinciding with S ^b , the second and third items are used to damp the component perpendicular to S ^b in the satellite triaxial angular rate, the fourth item is used for spin-up, and the fifth item is used to eliminate the coupling effect of the satellite triaxial angular rate.

It should be noted that the advantage of this embodiment is that the control scheme is designed to make the satellite sun sensor stably align with the sun vector first, and then make the designated satellite star surface normal vector align with the sun vector, which can be viewed from the sun sensor. In the case of limited field and under-actuated control, the reliability and stability of the satellite using the sun sensor to measure the sun vector to realize the daily spin stabilization control are significantly improved.

(2-3-2) Calculate the target of the three-axis magnetotorque by using the geomagnetic vector measurement value B ^b at the current moment measured by the three-axis magnetometer, combined with the target output torque _Tc3 obtained in step (2-3-1) Output magnetic moment M:

Among them, α ₃ is the angle between vector B ^b and T _c3 .

(2-3-3) With the three-axis magnetic torque device as the control component, the target output magnetic moment calculated in step (2-3-2) is converted into the control command of the three-axis magnetic torque device to drive the three-axis magnetic torque device Torquer for satellite attitude control.

与太阳敏感器测得的太阳矢量S^b间的夹角在大于等于设定的第三时间阈值(本实施例参考取值为25s)的时间内始终小于等于设定的第二角度阈值(本实施例参考取值为10°)，且卫星三轴角速率估计值/>

与设定的卫星自旋角速率向量/>

间的偏差在相同的所述时间内始终小于等于设定的第二角速率阈值(本实施例参考取值为0.1°/s)，则起旋阶段结束，卫星进入步骤(2-4)对日自旋稳定阶段；否则，按照步骤(2-3-1)-(2-3-3)继续进行起旋阶段的姿态控制。(2-3-4) In the spin-up phase, when the specified satellite body surface normal vector

The angle between the sun vector S ^b measured by the sun sensor is always less than or equal to the second angle threshold set (this embodiment) within the time greater than or equal to the set third time threshold (the reference value of this embodiment is 25s). The reference value of the embodiment is 10°), and the estimated value of the satellite three-axis angular rate/>

with the set satellite spin angular rate vector />

The deviation between them is always less than or equal to the second angular rate threshold set (the reference value of this embodiment is 0.1 °/s) in the same described time, then the spin-up phase ends, and the satellite enters step (2-4) to Daily spin stabilization phase; otherwise, continue the attitude control of the spin phase according to steps (2-3-1)-(2-3-3).

Wherein, in the attitude control process of the spin-up phase, if the satellite flies into the shadow area of the earth, then return to step (2-2-1).

(2-4) Sunward spin stabilization stage;

After entering the phase of spin stabilization toward the sun, since the satellite has a certain axis orientation along the normal vector direction of the designated star plane, it can be guaranteed that when the satellite enters the shadow area, the designated star plane still points to the sun vector and maintains a controlled attitude.

(2-4-1) When the satellite is in the sunny area, design the following control law to obtain the target output torque T _c4 of the three-axis magnetic torque device in the phase of the solar spin stabilization, so that the specified satellite body surface normal vector is in line with the sun The vectors are coincident and spin around the star's normal vector:

为指定的需要对准太阳矢量的卫星星体面法向单位矢量；ω_spin为设定的自旋角速率值(本实施例参考取值为1°/s)。上述控制力矩T_c4的表达式中，等式右边第一项用于控制/>

与S^b重合，第二项和第三项用于阻尼卫星三轴角速率中与S^b垂直的分量，第四项用于起旋，第五项用于消除卫星三轴角速率耦合效应。Among them, K ₅ , K ₆ , K ₇ , and K ₈ are positive fifth, sixth, seventh, and eighth control coefficients respectively (the reference values in this embodiment are 0.0004, 0.0008, 0.01, and 0.004, respectively);

ω _spin is the set spin angular rate value (the reference value of this embodiment is 1°/s). In the expression of the above control torque T _c4 , the first item on the right side of the equation is used to control

Coinciding with S ^b , the second and third items are used to damp the component perpendicular to S ^b in the satellite triaxial angular rate, the fourth item is used for spin-up, and the fifth item is used to eliminate the coupling effect of the satellite triaxial angular rate.

When the satellite is in the shadow area, the following control law is designed to obtain the target output torque T _c4 of the three-axis magnetic torque device in the spin stabilization phase against the sun, so as to control the satellite to maintain the spin state:

Wherein, K ₉ is a positive ninth control coefficient (the reference value of this embodiment is 0.01). In the expression of the control torque _Tc4 , the first term on the right side of the equation is used to control the satellite to maintain the spin angular rate, and the second term is used to eliminate the coupling effect of the three-axis angular rate of the satellite.

(2-4-2) Using the measured geomagnetic vector value B ^b at the current moment obtained from the three-axis magnetometer measurement, combined with the target output torque T _c4 obtained in step (2-4-1), calculate the target of the three-axis magnetotorque Output magnetic moment M:

Among them, α ₄ is the angle between vector B ^b and T _c4 .

(2-4-3) With the three-axis magnetic torque device as the control component, the target output magnetic torque calculated in step (2-4-2) is converted into the control command of the three-axis magnetic torque device to drive the three-axis magnetic torque device for satellite attitude control.

In order to realize the above-mentioned embodiments, the embodiment of the second aspect of the present invention proposes a satellite-to-sun spin stabilization attitude control device based on a pure magnetic control method, including:

The triaxial angular rate estimation module is used to calculate the estimated value of the satellite triaxial angular rate;

The target output torque calculation module is used to calculate the target output torque of each control stage of the satellite's rotation to the sun according to the estimated value of the three-axis angular rate of the satellite, and the control stages of the rotation to the sun include: the initial derotation stage , the sun capture phase, the spinning phase and the sun spin stabilization phase; wherein, in the sun capture phase, the spin phase and the sun spin stabilization phase, when the sun is located at the sun of the satellite When within the measurement field of view of the sensor, the sun vector measured by the sun sensor is considered in the target output torque;

The target output magnetic moment calculation module calculates the target output magnetic moment of the three-axis magnetic torque device of the satellite according to the target output torque and the geomagnetic vector measurement value measured by the three-axis magnetometer of the satellite;

The attitude control command calculation module is used to convert the target output magnetic moment into a control command of the three-axis magnetic torque device, so as to drive the three-axis magnetic torque device to perform attitude control of the satellite.

It should be noted that, the foregoing explanations for an embodiment of a satellite-sun spin stabilization attitude control method based on a pure magnetron method are also applicable to a satellite-sun spin stabilization based on a pure magnetron method in this embodiment The attitude control device will not be described in detail here. According to an embodiment of the present invention, based on a purely magnetically controlled satellite spin-stabilized attitude control device for the sun, by calculating the estimated value of the three-axis angular rate of the satellite; according to the estimated value of the three-axis angular rate of the satellite, respectively calculate The target output torque of each control phase of the satellite's spin to the sun, the control phases of the spin to the sun include: the initial derotation phase, the acquisition phase of the sun, the spin-up phase and the spin stabilization phase of the sun; wherein, in the During the acquisition phase for the sun, the spin-up phase, and the spin stabilization phase for the sun, when the sun is located within the measurement field of view of the sun sensor of the satellite, the target output torque is taken into account by the sun sensor The measured solar vector; according to the target output torque and the geomagnetic vector measurement value obtained by the three-axis magnetometer measurement of the satellite, calculate the target output magnetic moment of the three-axis magnetic torque device; the target output magnetic moment The torque is converted into a control instruction of the three-axis magnetic torque device, so as to drive the three-axis magnetic torque device to control the attitude of the satellite. Therefore, based on the pure magnetic control method, the satellite can be controlled from any position under the condition of fully considering the influence of factors such as the shadow area of the earth, the offset installation of the sun sensor and the solar panel, the field of view of the sun sensor, and the environmental disturbance moment. The initial state finally achieves spin stability to the sun, which can effectively ensure the smooth implementation of satellite equipment debugging, energy acquisition and other flight tasks. It has the advantages of low cost, low power consumption, simple process, high reliability, and good stability. It is suitable for absolute Most Earth-orbiting satellites.

In order to realize the above embodiments, the embodiment of the third aspect of the present invention proposes an electronic device, including:

at least one processor; and, a memory communicatively coupled to the at least one processor;

Wherein, the memory stores instructions that can be executed by the at least one processor, and the instructions are configured to execute the above-mentioned method for controlling the satellite's spin-stabilized attitude toward the sun based on a purely magnetic control method.

In order to realize the above-mentioned embodiments, the embodiment of the fourth aspect of the present invention provides a computer-readable storage medium, the computer-readable storage medium stores computer instructions, and the computer instructions are used to make the computer execute the above-mentioned pure magnetic-based The attitude control method of the satellite's spin-stabilized attitude towards the sun in the control mode.

It should be noted that the computer-readable medium mentioned above in the present disclosure may be a computer-readable signal medium or a computer-readable storage medium or any combination of the two. A computer readable storage medium may be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device, or device, or any combination thereof. More specific examples of computer-readable storage media may include, but are not limited to, electrical connections with one or more wires, portable computer diskettes, hard disks, random access memory (RAM), read-only memory (ROM), erasable Programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination of the above. In the present disclosure, a computer-readable storage medium may be any tangible medium that contains or stores a program that can be used by or in conjunction with an instruction execution system, apparatus, or device. In the present disclosure, however, a computer-readable signal medium may include a data signal propagated in baseband or as part of a carrier wave carrying computer-readable program code therein. Such propagated data signals may take many forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination of the foregoing. A computer-readable signal medium may also be any computer-readable medium other than a computer-readable storage medium, which can transmit, propagate, or transmit a program for use by or in conjunction with an instruction execution system, apparatus, or device . Program code embodied on a computer readable medium may be transmitted by any appropriate medium, including but not limited to wires, optical cables, RF (radio frequency), etc., or any suitable combination of the above.

The above-mentioned computer-readable medium may be included in the above-mentioned electronic device, or may exist independently without being incorporated into the electronic device. The above-mentioned computer-readable medium carries one or more programs, and when the above-mentioned one or more programs are executed by the electronic device, the electronic device is made to perform a satellite-sun spin stabilization based on a pure magnetron method in the above embodiment attitude control method.

Computer program code for carrying out the operations of the present disclosure can be written in one or more programming languages, or combinations thereof, including object-oriented programming languages—such as Java, Smalltalk, C++, and conventional Procedural Programming Language - such as "C" or a similar programming language. The program code may execute entirely on the user's computer, partly on the user's computer, as a stand-alone software package, partly on the user's computer and partly on a remote computer or entirely on the remote computer or server. In cases involving a remote computer, the remote computer can be connected to the user computer through any kind of network, including a local area network (LAN) or a wide area network (WAN), or it can be connected to an external computer (such as through an Internet service provider). Internet connection).

In the description of this specification, descriptions referring to the terms "one embodiment", "some embodiments", "example", "specific examples", or "some examples" mean that specific features described in connection with the embodiment or example , structure, material or characteristic is included in at least one embodiment or example of the present application. In this specification, the schematic representations of the above terms are not necessarily directed to the same embodiment or example. Furthermore, the described specific features, structures, materials or characteristics may be combined in any suitable manner in any one or more embodiments or examples. In addition, those skilled in the art can combine and combine different embodiments or examples and features of different embodiments or examples described in this specification without conflicting with each other.

In addition, the terms "first" and "second" are used for descriptive purposes only, and cannot be interpreted as indicating or implying relative importance or implicitly specifying the quantity of indicated technical features. Thus, the features defined as "first" and "second" may explicitly or implicitly include at least one of these features. In the description of the present application, "plurality" means at least two, such as two, three, etc., unless otherwise specifically defined.

Any process or method descriptions in flowcharts or otherwise described herein may be understood to represent modules, segments or portions of code comprising one or more executable instructions for implementing specific logical functions or steps of the process , and the scope of preferred embodiments of the present application includes additional implementations in which functions may be performed out of the order shown or discussed, including in substantially simultaneous fashion or in reverse order depending on the functions involved, which shall It should be understood by those skilled in the art to which the embodiments of the present application belong.

The logic and/or steps represented in the flowcharts or otherwise described herein, for example, can be considered as a sequenced listing of executable instructions for implementing logical functions, can be embodied in any computer-readable medium, For use with instruction execution systems, devices, or devices (such as computer-based systems, systems including processors, or other systems that can fetch instructions from instruction execution systems, devices, or devices and execute instructions), or in conjunction with these instruction execution systems, devices or equipment used. For the purposes of this specification, a "computer-readable medium" may be any device that can contain, store, communicate, propagate or transmit a program for use in or in conjunction with an instruction execution system, device or device. More specific examples (non-exhaustive list) of computer-readable media include the following: electrical connection with one or more wires (electronic device), portable computer disk case (magnetic device), random access memory (RAM), Read Only Memory (ROM), Erasable and Editable Read Only Memory (EPROM or Flash Memory), Fiber Optic Devices, and Portable Compact Disc Read Only Memory (CDROM). In addition, the computer-readable medium may even be paper or other suitable medium on which the program may be printed, as it may be possible, for example, by optically scanning the paper or other medium, followed by editing, interpretation or other suitable means if necessary. Processing to obtain programs electronically and store them in computer memory.

It should be understood that each part of the present application may be realized by hardware, software, firmware or a combination thereof. In the embodiments described above, various steps or methods may be implemented by software or firmware stored in memory and executed by a suitable instruction execution system. For example, if implemented in hardware, as in another embodiment, it can be implemented by any one or combination of the following techniques known in the art: Discrete logic circuits, ASICs with suitable combinational logic gates, programmable gate arrays (PGAs), field programmable gate arrays (FPGAs), etc.

Those of ordinary skill in the art can understand that all or part of the steps carried by the methods of the above embodiments can be completed by instructing related hardware through a program, and the program can be stored in a computer-readable storage medium. During execution, one or a combination of the steps of the method embodiments is included.

In addition, each functional unit in each embodiment of the present application may be integrated into one processing module, each unit may exist separately physically, or two or more units may be integrated into one module. The above-mentioned integrated modules can be implemented in the form of hardware or in the form of software function modules. If the integrated modules are realized in the form of software function modules and sold or used as independent products, they can also be stored in a computer-readable storage medium.

The storage medium mentioned above may be a read-only memory, a magnetic disk or an optical disk, and the like. Although the embodiments of the present application have been shown and described above, it can be understood that the above embodiments are exemplary and should not be construed as limitations on the present application, and those skilled in the art can make the above-mentioned The embodiments are subject to changes, modifications, substitutions and variations.

Claims (5)

1. A satellite-to-solar spin stable attitude control method based on a pure magnetic control mode is characterized by comprising the following steps:

calculating an estimated value of the satellite three-axis angular rate;

according to the estimated value of the satellite three-axis angular rate, respectively calculating the target output moment of the satellite in each control stage of the solar spin, wherein each control stage of the solar spin comprises the following steps: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;

calculating the target output magnetic moment of the triaxial magnetometers of the satellites according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellites;

Converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite;

the method for calculating the estimated value of the satellite three-axis angular rate comprises the following steps:

(1) Constructing a state vector

wherein ,

the three-axis angular rate of the satellite represented in the satellite body coordinate system; omega _x 、ω _y 、ω _z Respectively->

In the satellite body coordinate system O _b X _b Y _b Z _b O of (2) _b X _b 、O _b Y _b 、O _b Z _b An axial component;

(2) Constructing a state equation and a state transition matrix;

the satellite attitude dynamics equation is constructed as follows:

wherein ,

is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction; />

The control moment is output by the triaxial magnetic torquer;

the attitude dynamics equation is a state equation, and is linearized to obtain a state transition matrix

Wherein, deltaT is the operation period;

(3) Constructing a measurement equation and a measurement matrix;

geomagnetic vector measured values obtained by two continuous measurements of the triaxial magnetometer are respectively recorded as B ^b And

wherein ,

for the geomagnetic vector measurement value at the current moment, +. >

Respectively, the geomagnetic vector measured value at the current moment is in a satellite body coordinate system O _b X _b 、O _b Y _b 、O _b Z _b An axial component; />

For the geomagnetic vector measurement value at the previous moment, +.>

Respectively at the previous momentMagnetic vector measurement value is in satellite body coordinate system O _b X _b 、O _b Y _b 、O _b Z _b An axial component;

construction of an observation vector

The following are provided:

order the

To represent the vector of the rotation angle of the satellite body coordinate system in the time period delta T, theta _Δ Corresponding direction cosine matrix->

The approximation is expressed as:

wherein ,

for vector->

Is expressed as follows:

then a measurement equation is constructed as follows:

wherein ,

is an error term;

obtaining a measurement matrix according to the measurement equation

The method comprises the following steps:

(4) According to the state transition matrix obtained in the step (2) and the measurement matrix obtained in the step (3), calculating to obtain an estimated value of the satellite triaxial angular rate by using a Kalman filtering method

The target output torque calculation expression is as follows, either during the initial racemization phase, or during the pair-day acquisition phase, before the sun enters the sun sensor measurement field of view of the satellite, or during the pair-day acquisition phase, when the satellite is located in the ground shadow region:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₁ A positive first control coefficient;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₁ Is vector B ^b and T_c1 An included angle between the two;

during the pair-day capture phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₂ 、K ₃ 、K ₄ Positive second, third and fourth control coefficients respectively; />

The unit vector is the optical axis direction of the sun sensor; s is S ^b Sun vector measured for sun sensor at present moment,/->

The sun vector measured by the sun sensor at the previous moment; />

Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₂ Is vector B ^b and T_c2 An included angle between the two;

during the spinning phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₅ 、K ₆ 、K ₇ 、K ₈ Positive fifth, sixth, seventh and eighth control coefficients, respectively; />

The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega _spin For a set spin angular rate value; s is S ^b Sun vector measured for sun sensor at present moment,/->

The sun vector measured by the sun sensor at the previous moment; />

Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₃ Is vector B ^b and T_c3 An included angle between the two;

in the phase of stabilizing the solar spin, when the satellite is in the sun-irradiated region, the target output torque calculation expression is as follows:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₅ 、K ₆ 、K ₇ 、K ₈ Positive fifth, sixth, seventh and eighth control coefficients, respectively; />

The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega _spin For a set spin angular rate value; s is S ^b Sun vector measured for sun sensor at present moment,/->

The sun vector measured by the sun sensor at the previous moment; / >

Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b In the case of a centroid inertial principal axis coordinate system,matrix J is a diagonal matrix, diagonal elements J of matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction;

in the phase of the para-solar spin stabilization, when the satellite is in the earth shadow region, the target output torque calculation expression is as follows:

wherein ,K₉ A positive ninth control coefficient;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₄ Is vector B ^b and T_c4 An included angle between the two.

2. The method according to claim 1, wherein the method further comprises:

in the initial racemization stage, if the angular rate value of the satellite triaxial angular rate in the time greater than or equal to the set first time threshold value is always smaller than or equal to the set first angular rate threshold value, ending the initial racemization stage, and enabling the satellite to enter a daily capturing stage.

3. The method according to claim 1, wherein the method further comprises:

in the sun-to-sun capturing stage, if the unit vector of the optical axis direction of the sun sensor

Solar vector S measured by sun sensor at current moment ^b And (3) the included angle is always smaller than or equal to a set first angle threshold value within the time of being larger than or equal to a second time threshold value, ending the daily capturing phase, and enabling the satellite to enter the spinning-up phase.

4. The method according to claim 1, wherein the method further comprises:

in the spinning stage, if the satellite flies into an earth shadow area, returning the satellite to the sun-to-sun capturing stage;

if the normal vector of the satellite star surface which is required to be aligned with the sun vector is set

Solar vector S measured by sun sensor at current moment ^b The included angle is always smaller than or equal to the set second angle threshold value within the time larger than or equal to the set third time threshold value, and the satellite triaxial angular rate estimated value is +.>

And the set satellite spin angular rate vector +.>

And if the deviation is always smaller than or equal to the set second angular rate threshold value within the same time, ending the spinning stage, and enabling the satellite to enter a phase of stabilizing the spinning.

5. The utility model provides a satellite is to day spin stable gesture controlling means based on pure magnetic control mode which characterized in that includes:

the triaxial angular rate estimation module is used for calculating an estimated value of the triaxial angular rate of the satellite;

the target output torque calculation module is used for calculating target output torque of the satellite in each control stage of the solar spin according to the estimated value of the satellite three-axis angular rate, and the each control stage of the solar spin comprises: an initial racemization stage, a pair-day capturing stage, a spinning-up stage and a pair-day spin stabilization stage; wherein, during the sun-to-sun capture phase, the spin-up phase, and the sun-to-sun spin stabilization phase, the target output torque takes into account a solar vector measured by a sun sensor of the satellite when the sun is within a measurement field of view of the sun sensor;

The target output magnetic moment calculation module calculates the target output magnetic moment of the triaxial magnetometers of the satellite according to the target output moment and geomagnetic vector measurement values obtained by measurement of the triaxial magnetometers of the satellite;

the attitude control instruction calculation module is used for converting the target output magnetic moment into a control instruction of the triaxial magnetic torquer so as to drive the triaxial magnetic torquer to control the attitude of the satellite;

the method for calculating the estimated value of the satellite three-axis angular rate comprises the following steps:

(1) Constructing a state vector

wherein ,

the three-axis angular rate of the satellite represented in the satellite body coordinate system; omega _x 、ω _y 、ω _z Respectively->

In the satellite body coordinate system O _b X _b Y _b Z _b O of (2) _b X _b 、O _b Y _b 、O _b Z _b An axial component;

(2) Constructing a state equation and a state transition matrix;

the satellite attitude dynamics equation is constructed as follows:

wherein ,

is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction; />

The control moment is output by the triaxial magnetic torquer;

the attitude dynamics equation is a state equation, and is linearized to obtain a state transition matrix

Wherein, deltaT is the operation period;

(3) Constructing a measurement equation and a measurement matrix;

geomagnetic vector measured values obtained by two continuous measurements of the triaxial magnetometer are respectively recorded as B ^b And

wherein ,

for the geomagnetic vector measurement value at the current moment, +.>

Respectively, the geomagnetic vector measured value at the current moment is in a satellite body coordinate system O _b X _b 、O _b Y _b 、O _b Z _b An axial component; />

For the geomagnetic vector measurement value at the previous moment, +.>

Respectively, the geomagnetic vector measured value at the previous moment is in a satellite body coordinate system O _b X _b 、O _b Y _b 、O _b Z _b An axial component;

construction of an observation vector

The following are provided:

order the

To represent the vector of the rotation angle of the satellite body coordinate system in the time period delta T, theta _Δ Corresponding direction cosine matrix->

The approximation is expressed as:

wherein ,

for vector->

Is expressed as follows:

then a measurement equation is constructed as follows:

wherein ,

is an error term;

obtaining a measurement matrix according to the measurement equation

The method comprises the following steps:

(4) According to the state transition matrix obtained in the step (2) and the measurement matrix obtained in the step (3), calculating to obtain an estimated value of the satellite triaxial angular rate by using a Kalman filtering method

The target output torque calculation expression is as follows, either during the initial racemization phase, or during the pair-day acquisition phase, before the sun enters the sun sensor measurement field of view of the satellite, or during the pair-day acquisition phase, when the satellite is located in the ground shadow region:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₁ A positive first control coefficient;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₁ Is vector B ^b and T_c1 An included angle between the two;

during the pair-day capture phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₂ 、K ₃ 、K ₄ Positive second, third and fourth control coefficients respectively; />

The unit vector is the optical axis direction of the sun sensor; s is S ^b Sun vector measured for sun sensor at present moment,/->

The sun vector measured by the sun sensor at the previous moment; />

Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₂ Is vector B ^b and T_c2 An included angle between the two;

during the spinning phase and when the sun is within the measurement field of view of the sun sensor of the satellite, the target output torque calculation expression is as follows:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₅ 、K ₆ 、K ₇ 、K ₈ Positive fifth, sixth, seventh and eighth control coefficients, respectively; />

The normal unit vector of the satellite star surface which is set and needs to be aligned with the sun vector; omega _spin For a set spin angular rate value; s is S ^b Sun vector measured for sun sensor at present moment,/->

The sun vector measured by the sun sensor at the previous moment; />

Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₃ Is vector B ^b and T_c3 An included angle between the two;

in the phase of stabilizing the solar spin, when the satellite is in the sun-irradiated region, the target output torque calculation expression is as follows:

wherein ,

k is the estimated value of the three-axis angular rate of the satellite ₅ 、K ₆ 、K ₇ 、K ₈ Positive fifth, sixth, seventh and eighth control coefficients, respectively; />

For setting the satellite star surface to be aligned with the sun vectorA normal unit vector; omega _spin For a set spin angular rate value; s is S ^b Sun vector measured for sun sensor at present moment,/- >

The sun vector measured by the sun sensor at the previous moment; />

Is the rotational inertia matrix of the satellite, when the satellite body coordinate system O _b X _b Y _b Z _b When the matrix J is a diagonal matrix in the mass center inertial principal axis coordinate system, the diagonal elements J of the matrix J _x 、J _y 、J _z Respectively represent satellite windings O _b X _b 、O _b Y _b 、O _b Z _b Moment of inertia of the rotation in the axial direction;

in the phase of the para-solar spin stabilization, when the satellite is in the earth shadow region, the target output torque calculation expression is as follows:

wherein ,K₉ A positive ninth control coefficient;

calculating a target output magnetic moment M of the triaxial magnetic torquer:

wherein ,B^b The geomagnetic vector measurement value at the current moment; alpha ₄ Is vector B ^b and T_c4 An included angle between the two.

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