CN101214860A - A method of autonomously selecting the attitude determination method in the process of orbit control - Google Patents

Abstract

The method of independently selecting the attitude determination method in the orbit control process includes: (1) estimating the inertial attitude of the satellite based on the gyro measurement data; (2) judging whether it is necessary to introduce star sensor attitude correction: according to the measurement information of the gyroscope, determine the attitude angular velocity of the satellite Whether it exceeds the measurement range of the gyroscope, if not, continue to use the gyroscope to estimate the inertial attitude of the satellite, go to step (1), and estimate the inertial attitude of the satellite in the next cycle, if it exceeds the measurement range set by the gyroscope threshold value, then set the gyro overrun flag, and then judge whether the satellite’s three-axis attitude angular velocity meets the working condition requirements of the star sensor, and if it meets the working condition requirements of the star sensor, then go to step (3) , otherwise turn to (1); (3) introduce star sensor for attitude correction. The method of the invention reduces the attitude error in the track control process, improves the track control precision, and effectively ensures that the track control is completed in time and accurately.

Description

A method for autonomously selecting the attitude determination method in the course of orbit control

technical field

The invention relates to a method for determining the attitude of a spacecraft during orbit control. Especially when the disturbance torque of the orbit control engine is large, resulting in a large angular velocity of the satellite attitude, which exceeds the range of gyro measurement, the satellite can independently introduce the method of measuring information from the star sensor to accurately correct the satellite attitude.

Background technique

The gyro attitude estimation method is commonly used in the process of satellite attitude control, but due to the influence of gyro drift, it can only be used in a short period of time to ensure the accuracy of attitude determination. Moreover, the speed measurement range of a general high-precision gyro is relatively limited. When the angular velocity of the satellite attitude varies greatly beyond the gyro measurement range, only relying on the gyro data to estimate the satellite attitude will make the estimated attitude deviate from the real attitude of the satellite and affect the satellite control accuracy. The method of gyro prediction in the process of satellite orbit change avoids the more complex filtering algorithm, simplifies the implementation on the satellite, and is conducive to improving the reliability of orbit control, but the attitude deviation caused by the limited speed measurement range will make the satellite's The ignition attitude deviates from the required nominal ignition attitude, which directly affects the orbit control effect of the satellite. For satellites with high requirements for orbit control, such as lunar exploration and deep space exploration, orbit control errors may lead to the failure of the entire mission.

The method of combining gyroscope and star sensor to determine the attitude is also commonly used in the three-axis stabilization control of high-precision satellites. But it is mainly used in the steady-state control process of the satellite. When the satellite has a large attitude angular velocity, it will have an adverse effect on the star map recognition ability of the star sensor, and the star sensor cannot output data normally, so that the estimated attitude of the satellite cannot be corrected in time, and the estimated attitude of the satellite will still deviate from the actual situation. attitude. If it is used in the track control process, it will also affect the track control effect, and in severe cases, it will cause the track control task to fail.

During the orbit change control of most geostationary orbit or medium and low orbit satellites, the attitude determination only uses the measurement information of the gyroscope. When the orbit-changing engine disturbance torque is large, it may cause the satellite to have a large attitude angular velocity. When the angular velocity exceeds the measurement range of the gyroscope, only relying on the gyroscope for attitude estimation will cause the satellite's determined attitude to be inconsistent with the satellite's actual attitude. The actual attitude continuously deviates from the ignition attitude, which affects the accuracy of orbit control, and may lead to orbit control failure in severe cases.

Contents of the invention

The technology of the present invention solves the problem: overcomes the deficiencies of the prior art, and aims at the situation that the attitude angular velocity during orbit control is large and exceeds the measurement range of the gyro, and proposes a method of using the The gyro data estimates the satellite attitude. When a large attitude angular velocity exceeds the gyro measurement range, the attitude correction method of the star sensor is introduced independently, that is, a method of autonomously selecting the attitude determination method according to the measurement value of the satellite attitude angular velocity during orbit control is proposed. Reduce the attitude error in the orbit control process, improve the accuracy of orbit control, and effectively ensure that the orbit control is completed in a timely and accurate manner.

The method of autonomously selecting the attitude determination method in the orbit control process means that when the angular velocity of the star is small during the orbit control process, only the gyro prediction is selected for attitude determination; when the satellite attitude angular velocity is large and exceeds the gyro measurement range, the gyro prediction is selected Attitude determination is performed with the star sensor attitude correction method.

Technical solution of the present invention: the method for independently selecting the attitude-fixing mode in the track control process, is characterized in that comprising:

(1) Estimate satellite inertial attitude based on gyro measurement data:

First calculate the attitude angular velocity of the satellite according to the measurement information of the gyroscope, then calculate the three-axis absolute angular velocity, and finally estimate the inertial attitude of the satellite according to the absolute angular velocity;

(2) Determine whether it is necessary to introduce star sensor attitude correction:

According to the measurement information of the gyroscope, it is judged whether the attitude angular velocity of the satellite exceeds the measurement range of the gyroscope. If it does not exceed the measurement range, continue to use the gyroscope to estimate the inertial attitude of the satellite, and go to step (1) to perform the satellite inertial attitude of the next cycle. It is estimated that if the threshold value set according to the gyro measurement range is exceeded, the gyro overrun flag will be set, and then it will be judged whether the attitude angular velocity of the satellite's three-axis meets the working condition requirements of the star sensor. When the conditions require, then go to step (3), otherwise continue to use the gyro measurement data to estimate the inertial attitude of the satellite, go to (1).

(3) Introduce star sensor for attitude correction.

The advantage of the present invention compared with prior art is:

(1) The present invention can independently select different attitude-fixing methods according to the actual situation during the track control process, does not easily stop the track control, and can ensure the smooth progress of the track control.

(2) Through the introduction of star sensor information, the present invention effectively ensures the accuracy of attitude determination during the orbit change process of the high-thrust orbital control engine. Therefore, in the process of orbit control, the attitude control can be carried out accurately and in time.

(3) This method is suitable for high-thrust orbit change control of spacecraft. It is especially suitable for application in the follow-up series of deep space exploration satellites. It has the characteristics of good inheritance and good portability.

Description of drawings

Fig. 1 is the flowchart of the inventive method;

Fig. 2 is a flow chart of attitude determination mode conversion in the orbit control process;

Fig. 3 is the simulation curve of the inventive method;

Fig. 4 is the simulation curve of conventional method;

Among them, in the simulation curves of Fig. 3 and Fig. 4, curve 1 represents qw (0), curve 2 represents qw (1), and curve 3 represents qw (2), which are respectively the vector part of the satellite's actual attitude quaternion; curve 4 represents q(0), curve 5 represents q(1), and curve 6 represents q(2), which are the on-board estimated values of the satellite attitude quaternion vector part respectively.

Detailed ways

Velocity gyros mounted on satellites have a limited measurement range. In the process of satellite orbit change, due to the interference caused by the start-up of the orbit control engine and the interference caused by the sloshing of liquid fuel, etc., a relatively large satellite attitude angular velocity may be caused. When this angular velocity exceeds the gyro measurement range, the gyro will not be able to provide Correct angular velocity information, so the satellite cannot obtain accurate attitude information, and the attitude control accuracy cannot be guaranteed, thus affecting the orbit control accuracy.

The present invention adopts the measurement data of the gyroscope in the orbit control process to firstly use the measurement data of the gyroscope to estimate the attitude of the satellite, and at the same time judge the measurement output of the gyroscope in real time. After the measurement requirements of the gyroscope), the method of correcting the estimated attitude of the gyro by introducing the data of the star sensor is used to reduce the estimation error caused by the saturation of the gyroscope.

As shown in Figure 1, the specific implementation steps of the present invention are as follows:

(1) Gyro estimated inertial attitude: Estimate satellite inertial attitude based on gyroscope measurement data.

a. Gyro data conversion rate integral

The three-axis attitude angular velocity information of the satellite can be obtained by using the measurement information of the three gyroscopes. Taking the rate-integrating gyroscope as an example, according to the working principle of the rate-integrating gyroscope, the angle increment of the gyroscope in this sampling period can be obtained for each sampling, and the attitude angular velocity of the satellite's three-axis can be calculated after proper conversion.

According to the installation matrix B of the three gyroscopes (number i, j, k) participating in the attitude determination in the satellite body coordinate system, calculate the gyroscope output conversion matrix A=B ^-1 , and combine the measurement outputs of the three gyroscopes △g _i , △ g _j , △g _k , the satellite three-axis attitude angle incremental information △g _x , △g _y , △g _z can be obtained. Calculated by the following formula:

$\left[\begin{array}{c}\mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; x</span>}}\\ \mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; the\; y</span>}}\\ \mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; z</span>}}\end{array}\right]<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; A</span>}\left[\begin{array}{c}\mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; i</span>}}\\ \mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; j</span>}}\\ \mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; k</span>}}\end{array}\right]$

b. Three-axis absolute angular velocity calculation

According to the calculation result of the gyro data conversion, calculate the average angular velocity in a sampling period: $\widehat{\mathrm{\&omega;}}={\left[{\widehat{\mathrm{\&omega;}}}_x{\widehat{\mathrm{\&omega;}}}_{\mathrm{the\; y}}{\widehat{\mathrm{\&omega;}}}_z\right]}^T.$ In the calculation, pay attention to subtract the calibrated gyro constant value drift =[ _x  _y  _z ] ^T , unit: radian/hour. The calibration of the gyro constant value drift can be obtained by pre-calibration on the ground or real-time calibration on the star through the information of other sensors such as star sensors.

${\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}}_{\mathrm{<span\; class="notranslate">\; x</span>}}<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; x</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; \&Delta;t</span>}<span\; class="notranslate">\; -</span>{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; b</span>}}}_{\mathrm{<span\; class="notranslate">\; x</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; 3600</span>}$

${\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}}_{\mathrm{<span\; class="notranslate">\; the\; y</span>}}<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; the\; y</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; \&Delta;t</span>}<span\; class="notranslate">\; -</span>{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; b</span>}}}_{\mathrm{<span\; class="notranslate">\; the\; y</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; 3600</span>}$

${\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}}_{\mathrm{<span\; class="notranslate">\; z</span>}}<span\; class="notranslate">\; =</span>\mathrm{<span\; class="notranslate">\; \&Delta;</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}_{\mathrm{<span\; class="notranslate">\; z</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; \&Delta;t</span>}<span\; class="notranslate">\; -</span>{\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; b</span>}}}_{\mathrm{<span\; class="notranslate">\; z</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; 3600</span>}$

c. Inertial attitude estimation

According to the calculated three-axis absolute angular velocity, calculate the three-axis angular velocity increment of the satellite in this period, and use the following formula to estimate the inertial attitude of the satellite, in the form of quaternion $\stackrel{\widehat{-}}{q}=\left[\begin{array}{cccc}{\hat{q}}_1& {\hat{q}}_2& {\hat{q}}_3& {\hat{q}}_4\end{array}\right]$ give.

$\mathrm{<span\; class="notranslate">\; \&Delta;</span>}\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}<span\; class="notranslate">\; =</span>\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; \&omega;</span>}}<span\; class="notranslate">\; \&Center\; Dot;</span>\mathrm{<span\; class="notranslate">\; \&Delta;t</span>}$

$\stackrel{\stackrel{<span\; class="notranslate">\; ^</span>}{<span\; class="notranslate">\; -</span>}}{\mathrm{<span\; class="notranslate">\; q</span>}}<span\; class="notranslate">\; =</span>\stackrel{\stackrel{<span\; class="notranslate">\; ^</span>}{<span\; class="notranslate">\; -</span>}}{\mathrm{<span\; class="notranslate">\; q</span>}}<span\; class="notranslate">\; +</span>\frac{\mathrm{<span\; class="notranslate">\; 1</span>}}{\mathrm{<span\; class="notranslate">\; 2</span>}}\mathrm{<span\; class="notranslate">\; Eq</span>}<span\; class="notranslate">\; (</span>\stackrel{\stackrel{<span\; class="notranslate">\; ^</span>}{<span\; class="notranslate">\; -</span>}}{\mathrm{<span\; class="notranslate">\; q</span>}}<span\; class="notranslate">\; )</span>\mathrm{<span\; class="notranslate">\; \&Delta;</span>}\stackrel{<span\; class="notranslate">\; ^</span>}{\mathrm{<span\; class="notranslate">\; g</span>}}$

like ${\hat{q}}_4<0,$ but $\widehat{\stackrel{\&OverBar;}{q}}=-\widehat{\stackrel{\&OverBar;}{q}}$

$\stackrel{\stackrel{<span\; class="notranslate">\; ^</span>}{<span\; class="notranslate">\; -</span>}}{\mathrm{<span\; class="notranslate">\; q</span>}}<span\; class="notranslate">\; =</span>\stackrel{\stackrel{<span\; class="notranslate">\; ^</span>}{<span\; class="notranslate">\; -</span>}}{\mathrm{<span\; class="notranslate">\; q</span>}}<span\; class="notranslate">\; /</span>\mathrm{<span\; class="notranslate">\; Norm</span>}<span\; class="notranslate">\; (</span>\stackrel{\stackrel{<span\; class="notranslate">\; ^</span>}{<span\; class="notranslate">\; -</span>}}{\mathrm{<span\; class="notranslate">\; q</span>}}<span\; class="notranslate">\; )</span>$

In the above formula, Δt is the sampling period, and the function Eq is the calculation formula for estimating the satellite attitude quaternion by using the satellite attitude angular velocity information.

(2) Judging whether it is necessary to introduce star sensor attitude correction

According to the measurement information of the gyroscope, it is judged whether the attitude angular velocity of the satellite exceeds the measurement range of the gyroscope. If it does not exceed the measurement range, continue to use the gyroscope to estimate the inertial attitude of the satellite, and go to step (1) to perform the satellite inertial attitude of the next cycle. It is estimated that if the threshold value set according to the gyro measurement range is exceeded, the gyro overrun flag will be set, and then it will be judged whether the attitude angular velocity of the satellite's three-axis meets the working condition requirements of the star sensor. When the conditions require, then go to step (3), otherwise continue to use the gyro measurement data to estimate the inertial attitude of the satellite, go to (1).

(3) Star sensor attitude correction method

After the star sensor correction is introduced, the real-time attitude correction is performed using the conventional star sensor filtering algorithm. The filtering algorithm can be designed according to the classic Kalman filtering principle.

The specific application of this method in the lunar exploration satellite is given below:

As shown in Figure 2, the lunar probe uses six three-floating gyroscopes, any three of which can determine the angular velocity of the satellite's inertial attitude. The measurement range of each gyroscope is -0.9°/s～0.9°/s. Considering a certain design margin in engineering applications, the threshold value of the attitude angular velocity exceeding the gyroscope measurement range is selected as 0.8°/s.

Three medium-precision star sensors are used in the lunar probe to determine the three-axis attitude of the satellite. The star sensor can normally identify the star map and give the attitude data when the star has an angular velocity of 0.5°/s. Considering a certain design margin in engineering applications, the threshold value of introducing star sensor data for attitude correction is selected as 0.35°/s.

According to the above implementation steps, a simulation experiment was carried out for a lunar exploration satellite, and the simulation curve is shown in Figure 3. The simulation process is as follows: the orbit control starts, the satellite attitude angular velocity is large (1.5°/s), the gyro output is saturated, and the satellite attitude estimation value gradually deviates from the satellite's actual attitude, which causes the satellite attitude to deviate from the target ignition attitude. After 20s of simulation time, the star sensitivity correction is introduced. It can be seen from the curve that the satellite attitude estimation quickly converges to the real attitude, and the attitude error is less than 0.001 (rad) after 100s.

In order to compare the effect achieved by the method of the present invention, Fig. 4 shows the simulation curve without introducing star sensor correction in the orbit control process. The conditions of the simulation process are consistent with the above process, and the attitude of the satellite is only estimated from the gyro data. From the simulation curve, since the star sensor filter correction is not introduced, the satellite attitude estimation always deviates from the real attitude of the satellite, and the attitude error is about 0.08 (rad) after 100s, which will affect the orbit control accuracy.

The system described above is only one example of the present invention. Those skilled in the art can make various additions, improvements and replacements according to different requirements and design parameters without departing from the present invention. Therefore, the present invention is extensive.

Claims (4)

1. The method for independently selecting the attitude determination mode in the orbit control process is characterized in that it comprises: (1) Estimate satellite inertial attitude based on gyro measurement data: First calculate the attitude angular velocity of the satellite according to the measurement information of the gyroscope, then calculate the three-axis absolute angular velocity, and finally estimate the inertial attitude of the satellite according to the absolute angular velocity; (2) Determine whether it is necessary to introduce star sensor attitude correction: According to the measurement information of the gyroscope, it is judged whether the attitude angular velocity of the satellite exceeds the measurement range of the gyroscope. If it does not exceed the measurement range, continue to use the gyroscope to estimate the inertial attitude of the satellite, and go to step (1) to perform the satellite inertial attitude of the next cycle. It is estimated that if the threshold value set according to the gyro measurement range is exceeded, the gyro overrun flag will be set, and then it will be judged whether the attitude angular velocity of the satellite's three-axis meets the working condition requirements of the star sensor. If required, go to step (3); otherwise, continue to use the gyroscope measurement data to estimate the inertial attitude of the satellite, and go to step (1). (3) Introduce star sensor for attitude correction. 2. The method for independently selecting the attitude determination mode in the orbit control process according to claim 1, characterized in that: the gyroscope in the step (1) selects a rate-integrating gyroscope. 3. The method for autonomously selecting an attitude-fixing mode in the orbit control process according to claim 1, characterized in that: the inertial attitude in the step (1) is given in the form of a quaternion. 4. according to claim 1 in the track control process, select the method for attitude determination mode independently, it is characterized in that: the star sensor in the described step (2) is a CCD optical star sensor, and its working condition requires attitude angular velocity Less than or equal to 0.35 degrees/second.

Images