CN110803304A - A satellite attitude control system - Google Patents

Abstract

The present application relates to a satellite attitude control system, which consists of a three-axis magnetometer and a three-axis magnetic torquer. The present application also relates to a method for magnetically measuring the attitude of a magnetron satellite, which includes the following steps: estimating the satellite inertial angular velocity by using the magnetic field strength and its rate of change measured by a magnetometer; using the estimated value of the satellite inertial angular velocity as an input, using The magnetic torquers installed in the roll and yaw directions control the satellite to spin around the pitch axis to obtain the rate of change of the magnetic field strength in the direction of the pitch axis; Orientation of the magnetic torquer to damp the non-spinning axis angular velocity.

Description

A satellite attitude control system

technical field

The present application relates to the field of aerospace technology, in particular to a satellite attitude control system.

Background technique

In the existing satellite attitude control technology, a variety of attitude sensors such as sun sensor, star sensor, gyroscope, magnetometer, etc. are often configured at the same time, which are used for satellite attitude determination as the input of the controller, and thrusters are often used at the same time. Or the flywheel is used as the main actuator, and the magnetic torquer is mostly used as an auxiliary actuator and used for angular momentum unloading. Therefore, the configuration of multiple sensors and multiple actuators tends to lead to high cost of the attitude control system.

After the magnetic torque device is energized, a magnetic moment is generated, which interacts with the earth's magnetic field to generate a torque to control the satellite attitude. The magnetic torque device is fixedly installed, has no vibration, and has high reliability. It is a common actuator for satellite attitude control. The magnetometer measures the strength of the local magnetic field, and can be used in conjunction with the magnetic torquer to form a low-cost, high-reliability satellite attitude control system. At present, there are few studies on the attitude control system of magnetometry and magnetron satellite only using magnetometer and magnetic torquer. Under the condition of only configuring the magnetometer and the magnetic torquer, it is of great practical significance to design the satellite attitude control scheme to realize the satellite attitude control task. However, at present, magnetometers and magnetic torquers are mostly used for bias momentum satellite nutation damping and precession control, as well as momentum wheel angular momentum unloading, and rarely for the minimum mode attitude control composed of only magnetometers and magnetic torquers. systematic research.

Therefore, there is an urgent need in the art to develop a novel, simple and low-cost method for attitude control of an omnidirectional antenna communication satellite that can be applied to a sun-synchronous morning and evening orbit.

SUMMARY OF THE INVENTION

The purpose of this application is to provide a satellite attitude control system.

The purpose of the present application is also to provide a method for magnetically measuring the attitude of a magnetron satellite.

In order to achieve the above purpose, the present application provides the following technical solutions.

In a first aspect, the present application provides a satellite attitude control system consisting of a three-axis magnetometer and a three-axis magnetic torquer.

On the other hand, the application provides a method for magnetically measuring the attitude of a magnetron satellite, characterized in that it includes the following steps:

(1) Estimate the satellite inertial angular velocity using the magnetic field strength and its rate of change measured by the magnetometer;

(2) using the estimated value of the satellite inertial angular velocity as an input, utilizing the magnetic torquers mounted in the roll and yaw directions to control the satellite to spin around the pitch axis; and

(3) Using the rate of change of the magnetic field strength in the pitch direction as an input, the non-spin axis angular velocity is damped by a magnetic torquer installed in the pitch direction.

Compared with the prior art, the beneficial effect of the present application is to provide a novel, simple and low-cost satellite attitude control method.

Description of drawings

FIG. 1 is a schematic diagram of the decomposition of the magnetic field strength B _b of the present application in the satellite body coordinate system.

Detailed ways

The technical solutions of the present application will be clearly and completely described below with reference to the accompanying drawings and the embodiments of the present application.

In one aspect of the present application, it provides a satellite attitude control system, which consists only of a three-axis magnetometer and a three-axis magnetic torquer; the three-axis magnetometer is used to measure the satellite body coordinate system The magnetic field strength of the satellite is obtained by difference, and the rate of change of the magnetic field strength in the satellite body coordinate system is obtained as the input of the magnetron; the three-axis magnetic torquer provides the magnetron magnetic moment, and when the satellite runs in the sun-synchronous twilight orbit, the day orientation.

In another aspect of the present application, it provides a magnetic measurement and magnetron satellite attitude control method, which only uses a magnetometer and a magnetic torquer to design a technical solution, and realizes the control of the satellite attitude. This scheme can realize the sun-directed supplementation of on-board energy when the satellite is running in the sun-synchronous twilight orbit. The system consists of the following steps:

(1) Estimate the satellite inertial angular velocity using the magnetic field strength and its rate of change measured by the magnetometer;

(2) Using the estimated value of the inertial angular velocity of the satellite as an input, use the magnetic torque device installed in the roll and yaw directions to control the satellite to spin around the pitch axis (the pitch axis is the maximum or minimum inertia axis), and obtain the pitch axis direction. Magnetic field strength change rate;

(3) Using the rate of change of the magnetic field strength in the pitch direction as an input, the non-spin axis angular velocity is damped by a magnetic torquer installed in the pitch direction.

Specifically, step (1) uses a magnetometer to measure the magnetic field strength and its rate of change, and uses the magnetic field strength and rate of change to estimate the satellite inertial angular velocity. The decomposition of the magnetic field strength B _b in the satellite body coordinate system is shown in Figure 1.

(i) Angular velocity estimation

When the satellite spins around the pitch axis Y _b , ignoring the angular velocity around the roll axis X _b and the yaw axis Z _b , the angular velocity can be estimated

分别是B_x，B_y，B_z的变化率；α为磁场强度B_b在卫星本体系X_bOZ_b平面的投影与Z_b轴的夹角；

为α的变化率。上标T表示转置。Among them, ω _bi is the angular velocity vector relative to the inertial system of the satellite system; ω _bix , ω _biy , ω _biz are the components of ω _bi in the X _b , Y _b , Z _b directions of the satellite system, respectively; Β _b is the satellite system The magnetic field strength vector under , B _x , By , and B _z are the components of Β _b in the X _b , Y _b , and _{Z b} directions under the satellite body system respectively;

are the rate of change of B _x , By , and B _z respectively; α is the angle between the projection of the magnetic field intensity B _b on the X _b OZ _b plane of the satellite body and the _{Z b} axis;

is the rate of change of α. The superscript T means transpose.

(ii) Spin control law around the pitch axis

T _cy =k*(ω _bic -ω _biy )

Among them, ω _bic is the desired spin angular velocity, T _cy is the desired control torque of the pitch axis of the satellite; M _cx is the desired magnetic moment generated by the magnetic torque device installed in the roll direction; M _cz is the magnetic torque generated by the magnetic torque device installed in the yaw direction. the desired magnetic moment. * means multiplication. The superscript ² means squaring. k is the angular velocity gain coefficient, and k>0.

(iii) Non-spin axis angular velocity damping control law

Among them, M _cy is the expected magnetic moment generated by the magnetic moment device installed in the pitch direction of the satellite, sign(·) represents the sign of the variable, when (·) is positive, sign(·)=1, when (·) is When negative, sign(·)=-1.

The above description of the embodiments is for the convenience of those of ordinary skill in the art to understand and apply the present application. It will be apparent to those skilled in the art that various modifications to these embodiments can be readily made, and the generic principles described herein can be applied to other embodiments without inventive step. Therefore, the present application is not limited to the embodiments herein, and improvements and modifications made by those skilled in the art based on the contents disclosed in the present application without departing from the scope and spirit of the present application are all within the scope of the present application.

Claims (4)

1. a satellite attitude control system, is characterized in that, described system is made up of a three-axis magnetometer and a three-axis magnetic torque device, it is characterized in that, described three-axis magnetometer measures under the satellite body coordinate system The magnetic field strength, and the rate of change of the magnetic field strength in the satellite body coordinate system is obtained by difference, which is used as the input of the magnetron and the three-axis magnetic torquer provides the magnetron magnetic moment. Orientation, where: (1) Estimate the satellite inertial angular velocity using the magnetic field strength and its rate of change measured by the magnetometer; (2) using the estimated value of the satellite inertial angular velocity as input, and utilizing the magnetic torquers installed in the roll and yaw directions to control the satellite to spin around the pitch axis; and (3) Using the rate of change of the magnetic field strength in the pitch direction as an input, the non-spin axis angular velocity is damped by a magnetic torquer installed in the pitch direction. 2. The system of claim 1, wherein the calculation formula of the estimated value of the satellite inertial angular velocity is as follows:

Among them, ω _bi is the angular velocity vector relative to the inertial system of the satellite system; ω _bix , ω _biy , ω _biz are the components of ω _bi in the X _b , Y _b , Z _b directions of the satellite system, respectively; Β _b is the satellite system The magnetic field strength vector under , B _x , By , and B _z are the components of Β _b in the X _b , Y _b , and _{Z b} directions under the satellite body system respectively;

are the rate of change of B _x , By , and B _z respectively; α is the angle between the projection of the magnetic field intensity B _b on the X _b OZ _b plane of the satellite body and the _{Z b} axis;

is the rate of change of α. The superscript T means transpose. 3. system as claimed in claim 1, is characterized in that, the described calculation formula that controls satellite spin around pitch axis by magnetic torque device is as follows: T _cy =k*(ω _bic -ω _biy )

Among them, ω _bic is the desired spin angular velocity, T _cy is the desired control torque of the pitch axis of the satellite; M _cx is the desired magnetic moment generated by the magnetic torque device installed in the roll direction; M _cz is the magnetic torque generated by the magnetic torque device installed in the yaw direction. the desired magnetic moment. * means multiplication. The superscript ² means squaring. k is the angular velocity gain coefficient, and k>0. 4. The system of claim 1, wherein the calculation formula of the non-spin angular velocity is as follows:

Among them, M _cy is the expected magnetic moment generated by the magnetic moment device installed in the pitch direction of the satellite, sign(·) represents the sign of the variable, when (·) is positive, sign(·)=1, when (·) is When negative, sign(·)=-1.

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