A high-precision satellite attitude control method controlled only by a magnetic torque device

Abstract

A high-precision satellite attitude control method controlled by only a magnetic torque device, first, an expected control torque is calculated according to an attitude angle and an attitude angular velocity, then an expected control magnetic torque is calculated according to geomagnetic field information and the expected control torque, then an inertia angular velocity information is combined with a gravity gradient torque estimation and a previous period theory control torque estimation, a satellite residual magnetic torque is filtered and estimated by using a formula provided by the application, finally, the expected control magnetic torque is deducted from the estimated residual magnetic torque and converted into a magnetic torque device control voltage for output.

Description

Technical Field

[0001] This invention relates to the field of spacecraft attitude control, and in particular to a high-precision satellite attitude control method using only magnetic torque generators. Background Technology

[0002] Magnetic torquers are a commonly used attitude control actuator in low-Earth orbit spacecraft, offering advantages such as high reliability and low cost. They can be used for functions like satellite magnetic attitude acquisition and angular momentum unloading in momentum exchange mechanisms such as momentum wheels. When other actuators, such as momentum wheels or thrusters, fail and satellite attitude control is impossible, or when other actuators are not used or configured due to low power consumption or low cost requirements, using only magnetic torquers for high-precision satellite attitude control can significantly improve satellite operational reliability and extend satellite lifespan. However, due to the working principle of magnetic torquers, the generated working magnetic moment needs to interact with the Earth's magnetic field to form a magnetic control torque, and its direction is perpendicular to the Earth's magnetic field direction. Control along the Earth's magnetic field direction is impossible, therefore, using only magnetic torquers for three-axis satellite attitude control is highly challenging, and current technologies struggle to achieve high-precision attitude control. Summary of the Invention

[0003] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a high-precision satellite attitude control method that uses only magnetic torque generators for control, avoiding the inefficiency of feedback control and the indirect losses of torque feedforward methods, and effectively improving the accuracy of satellite attitude control.

[0004] The technical solution of this invention is: a high-precision satellite attitude control method using only magnetic torque converters, comprising:

[0005] Based on the satellite's current attitude angle Φ and attitude angular velocity Calculate the desired control torque T ex ;

[0006] Based on geomagnetic field information B and the desired control torque T ex Calculate the desired control magnetic moment M ex ;

[0007] Based on the inertial angular velocity information ω and the gravitational gradient torque, T is estimated. g And the theoretical control torque estimate T of the previous cycle c Regarding the satellite's remanent magnetic moment M rem Perform filtering estimation;

[0008] The desired control magnetic moment M ex Subtract the estimated remanent moment M rem And convert it into magnetic torque control voltage V c This output enables the control of the satellite's attitude using only a magnetic torque generator.

[0009] Preferably, the remanent magnetic moment M of the satellite rem Performing filter estimation includes:

[0010] Calculate the gravitational gradient torque T g ;

[0011] Based on the calculated gravitational gradient torque T g Theoretical control torque estimate T calculated in the previous cycle c And the three-axis angular velocities ω and ω1 of the satellite body relative to the inertial frame in this period and the previous period. lst Calculate the estimated remanent torque T rem ;

[0012] Calculate the filtered update value for the satellite's remanent magnetic moment:

[0013] M rem =M remlst +K Mrem ((M remlst ×B v )×B v -(T rem ×B v ) / |B|),

[0014] Where M remlst This is the calculated value of the satellite's remanent magnetic moment over the previous period, K. Mrem These are the filter coefficient values; B v B is the direction vector of the magnetic field, and B is the three-axis component representation of the geomagnetic field vector in the satellite body coordinate system.

[0015] Preferably, the gravity gradient torque T g =3ω0 2 (E×JE)

[0016] Where ω0 is the satellite's orbital angular velocity, J is the 3*3 dimensional satellite inertia matrix, and E is the transformation matrix A of the satellite's own system relative to its orbital system. BO The third column vector.

[0017] Preferably, the remanent torque estimation T rem =(Jω-Jω) lst ) / Δt-T g -T c

[0018] Where Δt is the period duration and J is the 3*3 dimensional satellite inertia matrix value.

[0019] Preferably, the filter coefficient value K Mrem The principle for determining this value is to strike a balance between reducing filtering noise and improving filtering tracking speed, and it is recommended that the value be between 0.001 and 0.01.

[0020] Preferably, the magnetic torque control voltage V c The calculation is as follows:

[0021] Calculation of theoretical control magnetic moment M c =M ex -M rem ;

[0022] Converted to magnetic torque control voltage V c =K VM M c K VM is the magnetic moment-to-voltage conversion coefficient.

[0023] Preferably, the magnetic moment M is controlled theoretically using calculations. c Update the theoretical control torque estimate T used for the next cycle calculation. c =M c ×B, where B is the three-axis component representation of the geomagnetic field vector in the satellite body coordinate system.

[0024] Preferably, the desired control torque is calculated as follows: Where K p and K d All are 3*3 dimension control coefficients, Φ and These are the three-axis Euler attitude angle and the three-axis Euler attitude angular velocity, respectively, where the control coefficient K... p and K d The stability of closed-loop control should be ensured.

[0025] Preferably, the calculation of the desired control magnetic moment includes:

[0026] Calculate the magnetic field direction vector B v =B / |B|, where B = [B x B y B z ] T The three-axis component representation of the geomagnetic field vector in the satellite body coordinate system, where |·| represents the vector length calculated for the vector;

[0027] Calculate the desired control magnetic moment M ex =T ex ×B v , where × represents the vector cross product calculation.

[0028] The advantages of this invention compared to the prior art are:

[0029] This invention differs from previous approaches that simply focus on feedback control design. By filtering and estimating the satellite's residual magnetic moment and directly compensating for it, combined with a feedback control law, it presents a novel design method that integrates feedback control and output feedforward compensation. This transforms the residual magnetic moment into a part of the control magnetic moment, avoiding the inefficiency of feedback control and the indirect losses of torque feedforward methods, and effectively improving the satellite attitude control accuracy.

[0030] This invention uses only a magnetic torque generator for high-precision satellite attitude control. Even when other actuators such as momentum wheels or thrusters fail or cannot be used, or when other actuators are not used or configured due to low power consumption or low cost requirements, high-precision satellite attitude control can still be achieved, greatly improving the reliability of satellite operation and extending the service life of the satellite. Attached Figure Description

[0031] Figure 1 This is a flowchart of a high-precision satellite attitude control method using only magnetic torque generators according to the present invention.

[0032] Figure 2 The three-axis Euler attitude angle curves of the satellite;

[0033] Figure 3 The output voltage curve of the satellite magnetic torque device. Detailed Implementation

[0034] This invention proposes a high-precision satellite attitude control method using only magnetic torque converters, such as... Figure 1 As shown, the method of the present invention includes the following specific implementation process:

[0035] (1) Based on attitude angle Φ and attitude angular velocity Calculate the desired control torque T ex Specifically:

[0036] Calculate the desired control torque Where K p and K d All are 3*3 dimension control coefficients, Φ and These are the three-axis Euler attitude angle and the three-axis Euler attitude angular velocity, respectively.

[0037] (2) Based on the geomagnetic field information B and the desired control torque T ex Calculate the desired control magnetic moment M ex Specifically:

[0038] Calculate the magnetic field direction vector B v =B / |B|, where B = [B x B y B z ] TThe three-axis component representation of the geomagnetic field vector in the satellite body coordinate system, where |·| represents the vector length calculated for the vector;

[0039] Calculate the desired control magnetic moment M ex =T ex ×B v , where × represents the vector cross product calculation.

[0040] (3) Estimate T based on the inertial angular velocity information ω and the gravitational gradient torque. g And the theoretical control torque estimate T of the previous cycle c Using the formula provided in this patent, the satellite's remanent magnetic moment M rem Perform filtering estimation. Specifically:

[0041] Calculate the gravitational gradient torque T g =3ω0 2 (E×JE), where ω0 is the satellite's orbital angular velocity, J is the 3*3 satellite inertia matrix, and E is the transformation matrix A of the satellite's own system relative to its orbital system. BO The 3rd column vector;

[0042] Calculate the estimated remanent torque T rem =(Jω-Jω) lst ) / Δt-T g -T c , where ω and ω lst These represent the current and previous period values ​​of the satellite's three-axis angular velocities relative to the inertial frame, respectively, where Δt is the period duration called by this algorithm, and T is the current period value. c This is an estimate of the theoretical control torque calculated in the previous cycle;

[0043] Calculate the filtered update value for the satellite's remanent magnetic moment:

[0044] M rem =M remlst +K Mrem ((M remlst ×B v )×B v -(T rem ×B v ) / B),

[0045] Where M remlst This is the calculated value of the satellite's remanent magnetic moment over the previous period, K. Mrem These are the filter coefficient values;

[0046] Update the remanent moment value M for the next cycle calculation. remlst =M rem .

[0047] (4) The desired control magnetic moment M ex Subtract the estimated remanent moment Mrem And convert it into magnetic torque control voltage V c Output the results. Specifically:

[0048] Calculation of theoretical control magnetic moment M c =M ex -M rem ;

[0049] Converted to magnetic torque control voltage V c =K VM M c K VM It is the magnetic moment-to-voltage conversion coefficient;

[0050] Update the theoretical control torque estimate T for the next cycle calculation. c =M c ×B.

[0051] Example 1:

[0052] A certain polar-orbiting satellite with an altitude of 500km has principal inertia of 100kgm on each of its three axes. 2 500kgm 2 550kgm 2 The inertia product is zero in all terms, and a 30Am beam is configured along each of the X, Y, and Z axes of the satellite. 2 Magnetic torque converter, with a magnetic moment-to-voltage conversion coefficient of K. VM =1V / Am 2 The satellite's three-axis remanent moments are 1 Am. 2 1Am 2 -1Am 2 The high-precision satellite attitude control method of the present invention, which uses only magnetic torque generators, is used with the following control parameter values:

[0053]

[0054] Satellite attitude control simulation results are as follows Figures 2-3 As shown. Among them Figure 2 It is the satellite's three-axis Euler attitude angle curve, as shown in the figure. The satellite's three-axis attitude control accuracy is better than 1 degree. Figure 3 The figure shows the output voltage curves of the satellite magnetic torque generator. The stable values ​​of the three-axis magnetic torque voltages are around -1V, -1V, and 1V, respectively, which correspond to stable values ​​of -1Am for the three-axis output magnetic torques. 2 -1Am 2 1Am 2 Nearby, accurate direct compensation was performed on the satellite's three-axis residual magnetic moment.

[0055] The above description is only the best specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the protection scope of the present invention.

[0056] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A high-precision satellite attitude control method using only magnetic torque generators, characterized in that... include: According to the satellite's current attitude angle and attitude angular velocity Calculate the desired control torque ; According to geomagnetic field information and desired control torque Calculate the desired control magnetic moment ; Based on inertial angular velocity information Combined with gravity gradient torque estimation And the theoretical control torque estimation of the previous cycle Regarding the satellite's remanent magnetic moment Perform filtering estimation; Desired control magnetic moment Subtract the estimated remanent moment And convert it into magnetic torque control voltage. This outputs data, enabling control of the satellite's attitude using only a magnetic torque converter. Regarding the satellite's remanent magnetic moment Performing filter estimation includes: Calculate the gravitational gradient torque ; Based on the calculated gravitational gradient torque Theoretical control torque estimation calculated in the previous cycle And the three-axis angular velocity values ​​of the satellite body relative to the inertial frame in this period and the previous period. , Calculation of remanent torque estimation ; Calculate the filtered update value of the satellite's remanent magnetic moment: ， in It is the calculated value of the satellite's remanent magnetic moment for the previous period. These are the filter coefficient values; The direction vector of the magnetic field. This represents the three-axis components of the geomagnetic field vector in the satellite's body coordinate system.

2. The high-precision satellite attitude control method using only a magnetic torque converter as described in claim 1, characterized in that: The gravity gradient torque in, This represents the satellite's orbital angular velocity. 3 3D satellite inertia matrix values The transformation matrix of the satellite's own system relative to its orbital system. The third column vector.

3. The high-precision satellite attitude control method using only a magnetic torque converter as described in claim 1, characterized in that: Remanent torque estimation ; in, For the duration of the cycle, 3 3D satellite inertia matrix values.

4. A high-precision satellite attitude control method using only a magnetic torque converter as described in claim 1, characterized in that: Filter coefficient values The principle for determining this value is to strike a balance between reducing filtering noise and improving filtering tracking speed, and it is recommended that the value be between 0.001 and 0.

01.

5. A high-precision satellite attitude control method using only a magnetic torque generator as described in claim 1, characterized in that: The magnetic torque control voltage The calculation is as follows: Calculation theory control magnetic moment ; Converted to magnetic torque control voltage ,in is the magnetic moment-to-voltage conversion coefficient.

6. A high-precision satellite attitude control method using only a magnetic torque converter as described in claim 5, characterized in that: Theoretical control of magnetic moment using calculations Update the theoretical control torque estimate for the next cycle calculation. ,in, This represents the three-axis components of the geomagnetic field vector in the satellite's body coordinate system.

7. A high-precision satellite attitude control method using only a magnetic torque converter as described in claim 1, characterized in that: The desired control torque is calculated as follows: ,in and All are 3 3D control coefficients and These are the three-axis Euler attitude angle and the three-axis Euler attitude angular velocity, respectively.

8. A high-precision satellite attitude control method using only a magnetic torque converter as described in claim 1, characterized in that: The calculation of the desired control magnetic moment includes: Calculate the magnetic field direction vector ,in This represents the three-axis components of the geomagnetic field vector in the satellite's body coordinate system. This indicates that the length of a vector is calculated. Calculate the desired control magnetic moment ,in This indicates the calculation of the vector cross product.

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