High-precision magnetic control and thruster combined attitude control method with optimal energy consumption

Abstract

The invention relates to a high-precision magnetic control and thruster combined attitude control method with optimal energy consumption, which comprises the following steps of: firstly, calculating an expected control moment by using a PD control law according to an attitude control quantity and an attitude angular velocity control quantity, and performing moment compensation by combining gravity gradient moment estimation; meanwhile, according to angular velocity measurement information, in combination with gravity gradient torque estimation and upper-cycle theoretical magnetic control torque estimation, filtering estimation is conducted on the residual magnetic moment of the satellite, then distribution is conducted on satellite three-axis control according to satellite magnetic control axis / thrust control axis setting, jet phase plane protection calculation is conducted on the thrust control direction, and the thrust control direction is calculated; the method comprises the following steps: estimating a residual magnetic moment, outputting air injection pulse width when necessary, performing small disturbance magnetic moment distribution calculation on a magnetic control direction according to geomagnetic field information and expected control moment, performing magnetic moment compensation and three-axis magnetic moment output in combination with a residual magnetic moment estimation result, and finally performing theoretical magnetic control moment estimation according to a three-axis magnetic moment calculation result. The method is used for next-period calculation.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft attitude control, and in particular to a high-precision magnetic control and thruster combined attitude control method with optimal energy consumption, which is applicable to all spacecraft that use magnetic control and thrust control to achieve high-precision three-axis attitude stabilization, such as gravity field measurement satellites and high-quietness scientific exploration satellites. The method can also be applied to all other spacecraft that require magnetic control or specific working modes of spacecraft. Background Art

[0002] The satellite used for precision gravity field measurement missions has a payload configuration including an extremely high-precision electrostatically suspended accelerometer. In order to reduce the impact of satellite disturbances on accelerometer measurements, the attitude control system excludes rotating and shaking components such as momentum wheels and liquid fuel from its design, uses cold air thrusters and continuous voltage magnetic torquers as attitude control actuators, and adopts a control method that optimizes working fluid consumption to ensure that the satellite can work normally and meet the design life requirements. During the development process, based on the technology that met the technical indicators in the early stage, the researchers carried out technical innovations, which greatly improved the control accuracy while further reducing the level of cold air working fluid consumption. Summary of the invention

[0003] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and provide a high-precision magnetic control and thruster combined attitude control method with optimal energy consumption. Different from the previous idea of ​​simply performing feedback control design, through filtering estimation and direct compensation of the satellite residual magnetic moment, combined with the feedback control law, it is a novel design method combining feedback control with output feedforward compensation, which converts the residual magnetic moment into a part of the control magnetic moment, avoids the inefficiency of feedback control and the indirect loss of the torque feedforward method, and effectively improves the satellite attitude control accuracy.

[0004] The technical solution of the present invention is: a high-precision magnetic control and thruster combined attitude control method with optimal energy consumption, which is controlled according to the following steps in each cycle:

[0005] Firstly, the expected control torque is calculated using the PD control law according to the attitude control variable and the attitude angular velocity control variable, and the torque compensation is performed in combination with the gravity gradient torque estimation. At the same time, the satellite remanent magnetic moment is filtered and estimated according to the angular velocity measurement information, the gravity gradient torque estimation and the upper period theoretical magnetic control torque estimation.

[0006] Then, the satellite three-axis control is allocated according to the satellite magnetic control axis / thrust control axis setting, the jet phase plane protection calculation is performed on the thrust control direction, and the jet pulse width is output when the phase plane threshold is triggered. The magnetic moment allocation calculation is performed on the magnetic control direction according to the geomagnetic field information and the expected control torque, and the magnetic moment compensation and limiting are performed in combination with the above-mentioned residual magnetic moment filtering estimation result, and the three-axis magnetic moment output is performed;

[0007] Finally, the theoretical magnetic control torque is estimated based on the output results of the three-axis magnetic moment for the calculation of the next cycle.

[0008] Preferably, the satellite magnetron axis / thrust control axis is fixed during the entire control process.

[0009] Preferably, the allocating of satellite three-axis control includes:

[0010] Two of the three axes of the satellite are set as magnetic control axes, and the other axis is set as a thrust control axis. The axis with an overall weaker magnetic control effect is set as the thrust control axis. The overall weaker magnetic control effect means that the inertia of the satellite along this axis is larger than that of the other two axes, or the direction of the geomagnetic field is closer to this axis than to the other two axes most of the time during the entire orbital period.

[0011] Preferably, the perturbation magnetic moment distribution is calculated as follows:

[0012] First, determine which two of the three axes of pitch, yaw and roll are magnetically controlled axes, and the third axis is the thrust control axis; calculate two candidate magnetic moment values ​​according to the component values ​​of the desired control torque after torque compensation in the two magnetically controlled axes, and determine the magnetic moment of the thrust control axis according to whether the signs of the two candidate magnetic moment values ​​are the same;

[0013] The magnetic moments of the two magnetic control axes are calculated based on the magnetic moment of the thrust control axis.

[0014] Preferably, two alternative magnetic moment values ​​M are calculated i1 =-T cj / B k , M i2 =T ck / B j , j and k represent two magnetic control axes, and i represents the thrust control axis; where T cj , T ck are the expected control torque after torque compensation on the magnetic control axis j and k, respectively, j , B k are the component values ​​of the geomagnetic induction intensity in the magnetic control axis j and k of the body coordinate system, respectively. If M i1 With M i2 The same sign, then the thrust control axis i magnetic moment M i The calculation formula is: i =sgn(M i1 )min(M i1 ,M i2 ); otherwise, let M i =0;

[0015] Then calculate the magnetic moment M of the magnetostriction axis k k =(T cj +M i Bk ) / B i , magnetic moment M of magnetostriction axis j j =(M i B j -T ck ) / B i , where B i It is the component value of the geomagnetic induction intensity on the thrust control axis in the body coordinate system.

[0016] Preferably, the desired control torque after torque compensation is T excomp =T ex +T g ;

[0017] Among them, T ex is the desired control torque, and the gravity gradient torque estimate T g =3ω0 2 (E×JE), ω0 is the satellite orbital angular velocity value, J is the 3*3 dimensional satellite inertia matrix value, and E is the transformation matrix A of the satellite system relative to the orbital system BO The third column vector of , × represents the vector cross product calculation.

[0018] Preferably, filtering and estimating the satellite remanent magnetic moment comprises:

[0019] Calculate the residual magnetic torque to estimate T rem =(Jω-Jω lst ) / Δt-T g -T c , where ω and ω lst are the current cycle value and the previous cycle value of the three-axis angular velocity of the satellite body relative to the inertial system, Δt is the cycle duration, T c Theoretical magnetron torque estimate calculated for the previous cycle;

[0020] Calculate the magnetic field direction vector B v =B / B|, where B = [B x B y B z ] T is the three-axis component representation of the geomagnetic induction intensity vector in the satellite coordinate system, and |·| represents the vector length calculated for the vector;

[0021] If thrust control was not performed in the previous cycle, calculate the filtered update value of the satellite residual magnetic moment:

[0022] M rem =M remlst +K Mrem ((M remlst ×B v )×B v -(T rem ×B v) / B),

[0023] Among them, M remlst is the calculated value of the satellite remanent magnetic moment in the previous period, K Mrem is the filter coefficient value;

[0024] If thrust control was performed in the previous cycle, the satellite residual magnetic moment maintains the estimated value of the previous cycle:

[0025] M rem =M remlst ;

[0026] Update the residual magnetic moment value M for the next cycle calculation remlst =M rem .

[0027] Preferably, the magnetic moment compensation and limiting are as follows:

[0028] The three-axis magnetic moment calculated by the small perturbation magnetic moment distribution is written as a vector form M c =[M x M y M z ] T , calculate the three-axis magnetic moment M after magnetic moment compensation out =M c -M rem ;M rem is the residual magnetic moment filtering estimation result;

[0029] Then the compensated three-axis magnetic moment M out After limiting, output M outlim That is, if the absolute value of the magnetic moment of an axis does not exceed the maximum magnetic moment value allowed by the axis, it will be output according to the magnetic moment value; otherwise, it will be output according to the maximum magnetic moment value allowed by the axis while retaining the positive and negative signs of the magnetic moment.

[0030] Preferably, the theoretical magnetron torque is estimated as follows:

[0031] T c =M outlim ×B;

[0032] M outlim is the value after the three-axis magnetic moment is limited, and B is the representation of the geomagnetic induction intensity vector in the satellite body coordinate system.

[0033] A computer-readable storage medium stores a computer program, wherein the computer program, when executed by a processor, implements the steps of any one of the methods of claims 1 to 8.

[0034] A high-precision magnetic control and thruster combined attitude control method with optimal energy consumption is applied to gravity field measurement satellites and high-quietness scientific exploration satellites, all spacecraft requiring magnetic control or spacecraft-specific working modes.

[0035] The advantages of the present invention compared with the prior art are:

[0036] (1) The present invention converts the residual magnetic moment into a part of the control magnetic moment by filtering, estimating and directly compensating the satellite residual magnetic moment, thereby avoiding the inefficiency of feedback control and the indirect loss of the torque feedforward method, effectively improving the satellite attitude control accuracy and reducing the thrust working fluid consumption;

[0037] (2) The present invention ensures the continuity of magnetic control by fixing the magnetic control axis / thrust control axis, avoids the disturbance of satellite attitude caused by magnetic control / jet switching, further reduces the consumption of thrust working fluid and also further improves the satellite attitude control accuracy;

[0038] (3) The present invention performs low-fluid consumption and high-precision satellite attitude control based on magnetic control and thrusters. When other actuators such as momentum wheels or CMGs fail and 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, thereby greatly improving the reliability of satellite operation and extending the service life of the satellite. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 This is a flow chart of a high-precision magnetic control and thruster combined attitude control method with optimal energy consumption according to the present invention; DETAILED DESCRIPTION

[0040] A high-precision magnetic control and thruster combined attitude control method with optimal energy consumption comprises the following steps:

[0041] (1) The desired control torque is calculated using the PD control law according to the attitude control variable and the attitude angular velocity control variable, and torque compensation is performed in combination with the gravity gradient torque estimation;

[0042] Calculate the expected control torque Where K p and K d are 3*3 dimensional control coefficients, Φ and They are the three-axis Euler attitude angle and the three-axis Euler attitude angular velocity respectively.

[0043] Calculate the gravity gradient moment estimate T g =3ω0 2 (E×JE), where ω0 is the satellite orbital angular velocity, J is the 3*3 dimensional satellite inertia matrix, and E is the transformation matrix A of the satellite system relative to the orbital system. BOThe third column vector of , × represents vector cross product calculation;

[0044] Calculate the desired control torque T after torque compensation excomp =T ex +T g .

[0045] (2) Based on the angular velocity measurement information combined with the gravity gradient torque estimation and the upper period theoretical magnetic control torque estimation, the satellite remanent magnetic moment is filtered and estimated;

[0046] Calculate the residual magnetic torque to estimate T rem =(Jω-Jω lst ) / Δt-T g -T c , where ω and ω lst are the current cycle value and the previous cycle value of the three-axis angular velocity of the satellite body relative to the inertial system (ω lst The initial value of is the initial value of ω), Δt is the cycle length of the algorithm call (this cycle refers to the satellite time when the algorithm of the present invention is called this time - the satellite time when the algorithm was called last time. For satellites, this cycle is usually called the control cycle, which is a constant value), T c Theoretical magnetron torque estimate calculated for the previous cycle;

[0047] Calculate the magnetic field direction vector B v =B / |B|, where B = [B x B y B z ] T is the three-axis component representation of the geomagnetic induction intensity vector in the satellite coordinate system, and |·| represents the vector length calculated for the vector;

[0048] If thrust control was not performed in the previous cycle, calculate the filtered update value of the satellite residual magnetic moment:

[0049] M rem =M remlst +K Mrem ((M remlst ×B v )×B v -(T rem ×B v ) / |B|),

[0050] Among them, M remlst is the calculated value of the satellite remanent magnetic moment in the previous period, K Mrem is the filter coefficient value;

[0051] If thrust control was performed in the previous cycle, the satellite residual magnetic moment maintains the estimated value of the previous cycle:

[0052] M rem =M remlst ;

[0053] Update the residual magnetic moment value M for the next cycle calculation remlst =M rem .

[0054] (3) Allocate the satellite three-axis control according to the satellite magnetic control axis / thrust control axis setting, calculate the thrust control direction according to the following step (4), and calculate the magnetic control direction according to the following step (5);

[0055] In step (3), the satellite magnetic control axis / thrust control axis fixed setting and three-axis control allocation are as follows:

[0056] Two of the three axes of the satellite are set as magnetic control axes, and magnetic control calculation is performed according to step (5); the other axis is set as the thrust control axis, and thrust control calculation is performed according to step (4); when setting the magnetic control axis / thrust control axis, the axis with the weaker overall magnetic control effect should generally be set as the thrust control axis, for example, the inertia of the satellite along this axis is relatively large, or the direction of the geomagnetic field is closer to this axis most of the time during the entire orbital cycle; the set magnetic control axis / thrust control axis is fixed during the entire control process to ensure the continuity of magnetic control.

[0057] (4) Perform jet phase plane protection calculation for the thrust control direction and output the jet pulse width when the phase plane threshold is triggered; the algorithm can be found in Satellite Attitude Dynamics and Control edited by Academician Tu Shancheng (Aerospace Press, 2001), page 442.

[0058] (5) performing small disturbance magnetic moment distribution calculation for the direction of magnetic control according to the geomagnetic field information and the desired control torque, and performing magnetic moment compensation and limiting in combination with the residual magnetic moment estimation result of the above step (2), and outputting the three-axis magnetic moment;

[0059] The small disturbance magnetic moment distribution calculation in step (5) is as follows:

[0060] a. If the pitch axis and yaw axis are magnetically controlled axes, then:

[0061] First, calculate two alternative magnetic moment values ​​M x1 =-T cy / B z , M x2 =T cz / B y , where T cy , T cz are the desired control torque components on the pitch axis and yaw axis after torque compensation calculated in step (1), respectively, and B y , B z are the component values ​​of the geomagnetic induction intensity in the pitch axis and yaw axis of the body coordinate system, respectively. If M x1 With M x2The same sign, then the rolling axis magnetic moment M x The calculation formula is: x =sgn(M x1 )min(M x1 ,M x2 ); otherwise, let M x =0;

[0062] Then calculate the yaw axis magnetic moment M z =(T cy +M x B z ) / B x , pitch axis magnetic moment M y =(M x B y -T cz ) / B x , where B x It is the component value of the geomagnetic induction intensity about the rolling axis of the body coordinate system.

[0063] b. If the roll axis and pitch axis are magnetically controlled axes, then:

[0064] First, calculate two alternative magnetic moment values ​​M z1 =-T cx / B y , M z2 =T cy / B x , where T cx , T cy are the expected control torque components on the roll axis and pitch axis after torque compensation calculated in step (1), respectively, x , B y are the component values ​​of the geomagnetic induction intensity in the rolling axis and the pitch axis of the body coordinate system respectively. If M z1 With M z2 The same sign, then the yaw axis magnetic moment M z The calculation formula is M z =sgn(M z1 )min(M z1 ,M z2 ); otherwise, let M z =0;

[0065] Then calculate the pitch axis magnetic moment M y =(T cx +M z B y ) / B z , rolling axis magnetic moment M x =(M z B x -T cy ) / B z , where B zIt is the component value of the geomagnetic induction intensity on the yaw axis of the body coordinate system.

[0066] c. If the roll axis and yaw axis are magnetically controlled axes, then:

[0067] First, calculate two alternative magnetic moment values ​​M y1 =T cx / B z , M y2 =-T cz / B x , where T cx , T cz are the expected control torque components on the roll axis and yaw axis after torque compensation calculated in step (1), respectively, x , B z are the component values ​​of the geomagnetic induction intensity in the roll axis and yaw axis of the body coordinate system, respectively. If M y1 With M y2 The same sign, then the pitch axis magnetic moment M y The calculation formula is M y =sgn(M y1 )min(M y1 ,M y2 ); otherwise, let M y =0;

[0068] Then calculate the rolling axis magnetic moment M x =(T cz +M y B x ) / B y , yaw axis magnetic moment M z =(M y B z -T cx ) / B y , where B y It is the component value of the geomagnetic induction intensity on the pitch axis of the body coordinate system.

[0069] The magnetic moment compensation and amplitude limiting output in step (5) are as follows:

[0070] The three-axis magnetic moment calculated in step (4) is written as a vector form M c =[M x M y M z ] T , calculate the three-axis magnetic moment M after magnetic moment compensation out =M c -M rem ;

[0071] Then the compensated three-axis magnetic moment M out After limiting, output M outlimThat is, if the absolute value of the magnetic moment of an axis does not exceed the maximum magnetic moment value allowed by the axis, it will be output according to the magnetic moment value; otherwise, it will be output according to the maximum magnetic moment value allowed by the axis while retaining the positive and negative signs of the magnetic moment.

[0072] (6) According to the result of the three-axis magnetic moment calculation in step (5), the theoretical magnetic control torque is estimated for the calculation of the next cycle. The theoretical magnetic control torque is estimated as follows:

[0073] T c =M outlim ×B.

[0074] Embodiment 1:

[0075] Take a low-orbit satellite with an orbital inclination of about 90 degrees as an example. Figure 1 As shown, the specific steps of the present invention are as follows:

[0076] (1) The desired control torque is calculated using the PD control law according to the attitude control variable and the attitude angular velocity control variable, and torque compensation is performed in combination with the gravity gradient torque estimation; First, the desired control torque is calculated using the PD control law Where K p and K d are 3*3 dimensional control coefficients, which can be set according to the general PD control design method. are the three-axis Euler attitude angle and the three-axis Euler attitude angular velocity, respectively, which are provided by the satellite attitude determination algorithm; then the gravity gradient moment is calculated to estimate T g =3ω0 2 (E×JE), where ω0 is the satellite orbital angular velocity value, which is given by the satellite orbit calculation, J is the 3*3 dimensional satellite inertia matrix value, and E is the transformation matrix A of the satellite system relative to the orbital system BO The third column vector is provided by the satellite attitude algorithm, and × represents the vector cross product calculation; then the desired control torque T after torque compensation is calculated. excomp =T ex +T g .

[0077] (2) Based on the angular velocity measurement information combined with the gravity gradient torque estimate and the upper period theoretical magnetic control torque estimate, the satellite remanent magnetic moment is filtered and estimated. First, the remanent magnetic moment estimate T is calculated. rem =(Jω-Jω lst ) / Δt-T g -T c , where ω and ω lst are the current cycle value and the previous cycle value of the three-axis angular velocity of the satellite body relative to the inertial system, which are provided by the satellite attitude determination algorithm. Δt is the cycle duration called by this algorithm. T c is the theoretical magnetic control torque estimate calculated in the previous cycle; then calculate the magnetic field direction vector Bv =B / |B|, where B = [B x B y B z ] T is the three-axis component representation of the geomagnetic induction intensity vector in the satellite coordinate system, which is measured by the satellite magnetometer or calculated by the geomagnetic field. |·| represents the length of the vector calculated for the vector. Then, if thrust control was not performed in the previous cycle, the filtered update value of the satellite residual magnetic moment is calculated:

[0078] M rem =M remlst +K Mrem ((M remlst ×B v )×B v -(T rem ×B v ) / |B|),

[0079] Among them, M remlst is the calculated value of the satellite remanent magnetic moment in the previous period, K Mrem is the filter coefficient value, which is generally not more than 0.01 to ensure the filter stability;

[0080] If thrust control was performed in the previous cycle, the satellite residual magnetic moment maintains the estimated value M of the previous cycle rem =M remlst ; Finally update the residual magnetic moment value M used for the next cycle calculation remlst =M rem .

[0081] (3) Allocate the satellite three-axis control according to the satellite magnetic control axis / thrust control axis setting, calculate the thrust control direction according to the following step (4), and calculate the magnetic control direction according to the following step (5); set two of the three satellite axes as magnetic control axes, and perform magnetic control calculation according to step (5); and the other axis as thrust control axis, and perform thrust control calculation according to step (4); when setting the magnetic control axis / thrust control axis, the axis with the weaker magnetic control effect should generally be set as the thrust control axis, for example, the satellite has a relatively large inertia along this axis, or the direction of the geomagnetic field is closer to this axis most of the time during the entire orbital cycle, etc. Since the satellite orbit inclination is near 90 degrees, the direction of the geomagnetic field is mainly located near the plane opened by the roll axis and the yaw axis, the satellite configuration is long and narrow, and the roll axis inertia is significantly smaller than the inertia of the other two axes. Therefore, the yaw axis is set as the thrust control axis, and the roll axis and the pitch axis are set as the magnetic control axis; the set magnetic control axis / thrust control axis is fixed throughout the control process to ensure the continuity of magnetic control.

[0082] (4) Perform jet phase plane protection calculation for the thrust control direction and output the jet pulse width when necessary. For the specific algorithm, refer to page 442 of Satellite Attitude Dynamics and Control edited by Academician Tu Shancheng (Aerospace Press, 2001).

[0083] (5) The small disturbance magnetic moment distribution calculation is performed on the direction of magnetic control according to the geomagnetic field information and the expected control torque, and the magnetic moment compensation and limiting are performed in combination with the residual magnetic moment estimation result of the above step (2), and the three-axis magnetic moment output is performed; since the satellite sets the roll axis and the pitch axis as the magnetic control axis, first calculate two alternative magnetic moment values ​​M z1 =-T cx / B y , M z2 =T cy / B x , where T cx , T cy are the expected control torque components on the roll axis and pitch axis after torque compensation calculated in step (1), respectively, x , B y are the component values ​​of the geomagnetic induction intensity in the rolling axis and the pitch axis of the body coordinate system respectively. If M z1 With M z2 The same sign, then the yaw axis magnetic moment M z The calculation formula is M z =sgn(M z1 )min(M z1 ,M z2 ), otherwise let M z =0; then calculate the pitch axis magnetic moment M y =(T cx +M z B y ) / B z , rolling axis magnetic moment M x =(M z B x -T cy ) / B z , where B z is the component value of the geomagnetic induction intensity in the yaw axis of the body coordinate system; then the three-axis magnetic moment M after magnetic moment compensation is calculated out =M c -M rem , where M c =[M x M y M z ] T is the vector form of the three-axis magnetic moment calculated in step (4); finally, the compensated three-axis magnetic moment M out After limiting, output M outlimThat is, if the absolute value of the magnetic moment of an axis does not exceed the maximum magnetic moment value allowed by the axis, it will be output according to the magnetic moment value; otherwise, it will be output according to the maximum magnetic moment value allowed by the axis while retaining the positive and negative signs of the magnetic moment.

[0084] (6) According to the result of the three-axis magnetic moment calculation in step (5), the theoretical magnetic control torque is estimated for the calculation of the next cycle; the specific formula is T c =M outlim ×B;

[0085] (7) Repeat the process throughout the satellite operation. Figure 1 The energy-optimized high-precision magnetic control and thruster combined attitude control process shown in the figure achieves fluid-saving and high-precision attitude control, ensuring that the satellite completes its scientific exploration mission safely and effectively.

[0086] The above description is only the best specific implementation mode of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by any technician familiar with the technical field within the technical scope disclosed by the present invention should be covered within the protection scope of the present invention.

[0087] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A high-precision magnetic control and thruster combined attitude control method with optimal energy consumption, characterized in that In each cycle, the control is performed as follows: Firstly, the expected control torque is calculated using the PD control law according to the attitude control variable and the attitude angular velocity control variable, and the torque compensation is performed in combination with the gravity gradient torque estimation. At the same time, the satellite remanent magnetic moment is filtered and estimated according to the angular velocity measurement information, the gravity gradient torque estimation and the upper period theoretical magnetic control torque estimation. Then, the satellite three-axis control is allocated according to the satellite magnetic control axis / thrust control axis setting, the jet phase plane protection calculation is performed on the thrust control direction, and the jet pulse width is output when the phase plane threshold is triggered. The magnetic moment allocation calculation is performed on the magnetic control direction according to the geomagnetic field information and the expected control torque, and the magnetic moment compensation and limiting are performed in combination with the above-mentioned residual magnetic moment filtering estimation result, and the three-axis magnetic moment output is performed; Finally, the theoretical magnetic control torque is estimated based on the output results of the three-axis magnetic moment for the calculation of the next cycle.

2. The joint posture control method according to claim 1, characterized in that: The satellite magnetostriction axis / thrust control axis is fixedly set during the entire control process.

3. The joint posture control method according to claim 2, characterized in that: The distribution of satellite three-axis control includes: Two of the three axes of the satellite are set as magnetic control axes, and the other axis is set as a thrust control axis. The axis with an overall weaker magnetic control effect is set as the thrust control axis. The overall weaker magnetic control effect means that the inertia of the satellite along this axis is larger than that of the other two axes, or the direction of the geomagnetic field is closer to this axis than to the other two axes most of the time during the entire orbital period.

4. The joint posture control method according to claim 2, characterized in that: The magnetic moment distribution is calculated as follows: First, determine which two of the three axes of pitch, yaw and roll are magnetically controlled axes, and the third axis is the thrust control axis; calculate two candidate magnetic moment values ​​according to the component values ​​of the desired control torque after torque compensation in the two magnetically controlled axes, and determine the magnetic moment of the thrust control axis according to whether the signs of the two candidate magnetic moment values ​​are the same; The magnetic moments of the two magnetic control axes are calculated based on the magnetic moment of the thrust control axis.

5. The joint posture control method according to claim 4, characterized in that: Calculate two alternative magnetic moment values ​​M i1 =-T cj / B k , M i2 =T ck / B j , j and k represent two magnetic control axes, and i represents the thrust control axis; where T cj 、T ck are the expected control torque after torque compensation on the magnetic control axis j and k, respectively, j , B k are the component values ​​of the geomagnetic induction intensity in the magnetic control axis j and k of the body coordinate system, respectively. If M i1 With M i2 The same sign, then the thrust control axis i magnetic moment M i The calculation formula is: i =sgn(M i1 )min(M i1 ,M i2 ); otherwise, let M i =0; Then calculate the magnetic moment M of the magnetostriction axis k k =(T cj +M i B k ) / B i , magnetic moment M of magnetostriction axis j j =(M i B j -T ck ) / B i , where B i It is the component value of the geomagnetic induction intensity on the thrust control axis in the body coordinate system.

6. The joint posture control method according to claim 1, characterized in that: Expected control torque after torque compensation T excomp =T ex +T g ; Among them, T ex is the desired control torque, and the gravity gradient torque estimate T g =3ω0 2 (E×JE), ω0 is the satellite orbital angular velocity value, J is the 3*3 dimensional satellite inertia matrix value, and E is the transformation matrix A of the satellite system relative to the orbital system BO The third column vector of , × represents the vector cross product calculation.

7. The combined posture control method according to claim 6, characterized in that: The filtering estimation of satellite remanent magnetic moment includes: Calculate the residual magnetic torque to estimate T rem =(Jω-Jω lst ) / Δt-T g -T c , where ω and ω lst are the current cycle value and the previous cycle value of the three-axis angular velocity of the satellite body relative to the inertial system, Δt is the cycle duration, T c Theoretical magnetron torque estimate calculated for the previous cycle; Calculate the magnetic field direction vector B v =B / B|, where B = [B x B y B z ] T is the three-axis component representation of the geomagnetic induction intensity vector in the satellite coordinate system, |·| represents the vector length calculated for the vector; If thrust control was not performed in the previous cycle, calculate the filtered update value of the satellite residual magnetic moment: M rem =M remlst +K Mrem ((M remlst ×B v )×B v -(T rem ×B v ) / B), Among them, M remlst is the calculated value of the satellite remanent magnetic moment in the previous period, K Mrem is the filter coefficient value; If thrust control was performed in the previous cycle, the satellite residual magnetic moment maintains the estimated value of the previous cycle: M rem =M remlst ; Update the residual magnetic moment value M for the next cycle calculation remlst =M rem .

8. The joint posture control method according to claim 1, characterized in that: The magnetic moment compensation and limiting are as follows: The three-axis magnetic moment calculated by the small perturbation magnetic moment distribution is written as a vector form M c =[M x M y M z ] T , calculate the three-axis magnetic moment M after magnetic moment compensation out =M c -M rem ;M rem is the residual magnetic moment filtering estimation result; Then the compensated three-axis magnetic moment M out After limiting, output M outlim That is, if the absolute value of the magnetic moment of an axis does not exceed the maximum magnetic moment value allowed by the axis, it will be output according to the magnetic moment value; otherwise, it will be output according to the maximum magnetic moment value allowed by the axis while retaining the positive and negative signs of the magnetic moment.

9. The joint posture control method according to claim 1, characterized in that: The theoretical magnetic control torque is estimated as follows: T c =M outlim ×B; M outlim is the value after the three-axis magnetic moment is limited, and B is the representation of the geomagnetic induction intensity vector in the satellite body coordinate system.

10. A high-precision magnetic control and thruster combined attitude control method with optimal energy consumption, which is applied to gravity field measurement satellites and high-quietness scientific exploration satellites, all spacecraft requiring magnetic control or spacecraft specific working modes.

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