Decoupling iterative calibration method for radial shear joint control based on orbit perturbation model inversion

Abstract

The present invention discloses a radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion, which is based on the fact that there are few existing studies on the inverse process of the radial shear joint control process, and there is currently no radial shear joint control decoupling calibration method based on orbital perturbation model inversion. A radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion is provided, which has good reliability, strong operability, and is easy to promote and use. The radial shear velocity increment is calculated by inverting the orbital elements before and after satellite control and the orbital perturbation model, thereby realizing synchronous decoupling calibration of radial thrusters and tangential thrusters. Considering the limited calculation accuracy of the analytical model, numerical iteration is used in the calibration to ensure the effectiveness of the calibration method. The proposed method can improve the orbital control accuracy of the radial shear joint control of spacecraft, effectively save satellite fuel consumption, have certain economic benefits for the on-orbit operation of spacecraft, and have important guiding significance for mission implementation.

Description

Technical Field

[0001] The invention belongs to the technical field of spacecraft measurement and control, and relates to a radial shear joint control decoupling iterative calibration method based on orbit perturbation model inversion. Background Art

[0002] The on-orbit calibration of the orbit control thruster coefficient is based on the measured orbit elements before control, the theoretical orbit elements after control, and the measured orbit elements after control, and based on the relevant models and algorithms, a comprehensive calibration is performed on various error factors that mainly affect the orbit control effect, such as the orbit control thruster thrust error, satellite attitude error, etc. There are many factors that affect the orbit control effect, and the relationship between them is very complex. It is impossible to accurately decouple and separate them in engineering practice. Generally speaking, since the orbit control thruster is affected by factors such as on-orbit operating conditions and environment, there is a certain deviation in the thrust size of the thruster, and this deviation is often the main factor affecting the orbit control effect. Therefore, it is necessary to calibrate the thrust coefficient of the thruster after each orbit control, so that the calibrated coefficient can be used in subsequent orbit control to improve the orbit control accuracy.

[0003] With the continuous development of satellite application technology, satellite formation is a common satellite application mode. Under the limitation of satellite size and mass, satellite formation can greatly improve the working efficiency of single satellite payload. Compared with single satellite control, satellite formation flight has higher control accuracy requirements for orbit shape and satellite position. In order to achieve fine control of formation configuration, radial control and tangential control can be used. Since radial and tangential velocity increments have coupling effects on four orbital elements such as semi-major axis, eccentricity, perigee angle, and mean anomaly in the orbital plane, and the efficiency of radial control is relatively low compared to that of tangential control, tangential control or radial-tangential decoupling control is generally used. With the gradual increase in the configuration accuracy requirements of spacecraft formations, in order to improve the control timeliness, scholars have carried out a lot of research on radial-tangential joint control. The idea is mainly based on orbital perturbation model or relative motion model to analyze the influence of perturbation factors on formation configuration, and on this basis to achieve high-precision control of formation configuration. However, there are few studies on the inverse process of radial shear joint control, and there is currently no radial shear joint control decoupling calibration method based on orbital perturbation model inversion. Summary of the invention

[0004] The purpose of the present invention is to provide a radial-tangential joint control decoupling iterative calibration method based on orbit perturbation model inversion, so as to realize synchronous decoupling calibration of radial thrusters and tangential thrusters.

[0005] The technical solution adopted by the present invention is a radial shear joint control decoupling iterative calibration method based on orbit perturbation model inversion, which is implemented according to the following steps:

[0006] Step 1: Determine the parameters of the satellite orbit before the intermediate time of radial joint control, including the time T of the satellite orbit s , semi-major axis a s , eccentricity e s 、Inclination angle i s , right ascension of ascending node Ω s 、Arg of perigee ω s 、Mean anomaly angle M s ;

[0007] Step 2: Set the initial value of the radial joint control decoupling iteration calibration, including the time T of the temporary track t =T s , semi-major axis a t =a s , eccentricity e t =e s 、Inclination angle i t =i s , right ascension of ascending node Ω t =Ω s 、Arg of perigee ω t =ω s 、Mean anomaly angle M t =M s , radial actual control speed increment Δv re =0, tangential actual control speed increment Δv te =0;

[0008] Step 3: Calculate the orbital satellite position vector before the intermediate time control of the radial joint control in the J2000.0 coordinate system and velocity vector

[0009] Where F1(t s ,a s ,e s ,i s ,Ω s ,ω s ,M s ) is based on the satellite orbit time T s , semi-major axis a s , eccentricity e s 、Inclination angle i s , right ascension of ascending node Ω s 、Arg of perigee ω s Peace anomaly M s Calculate the satellite position vector and velocity vector Function of

[0010] Step 4: Based on the temporary orbital eccentricity e t Peace anomaly M t, iteratively calculate the orbital anomaly angle E at the intermediate moment of radial joint control t

[0011] E t =M t +e t sinE t ;

[0012] Step 5: Calculate the temporary orbit true anomaly angle f at the intermediate moment of radial joint control t and the orbital average angular velocity n t ;

[0013] Step 6: Determine the satellite orbit after the intermediate time of radial joint control. The parameters include the time T of the satellite orbit e , semi-major axis a e , eccentricity e e , inclination angle i e , right ascension of ascending node Ω e , the argument of perigee ω e , mean anomaly angle M e ;

[0014] Step 7: Calculate the satellite position vector after the intermediate time of radial joint control in the J2000.0 coordinate system and velocity vector

[0015] Step 8: Calculate the change of the semi-major axis Δa, the change of the eccentricity Δe and the change of the mean anomaly ΔM of the radial joint control orbit;

[0016] Step 9: Calculate the radial control velocity increment correction δv based on the orbital perturbation model inversion re and the tangential control speed increment correction δv te ;

[0017] Step 10: Calculate the actual radial control speed increment Δv re and the actual tangential control speed increment Δv te ;

[0018] Step 11: Calculate the temporary orbit satellite position vector in the J2000.0 coordinate system and the velocity vector

[0019] Step 12: Based on the time T of the satellite's temporary orbit t ,Location speed Calculate the temporary orbit elements of the satellite;

[0020] Step 13: Calculate the comprehensive deviation J between the temporary orbit and the controlled satellite orbit;

[0021] Step 14: Determine whether the comprehensive deviation J meets the accuracy requirement. If it does, that is, J<δ, where δ is the iterative calculation threshold, which can be manually selected as needed, then proceed to step 15, otherwise go to step 4;

[0022] Step 15: Decouple and calibrate the radial thruster and tangential thruster coefficients, and then realize the decoupled iterative calibration of radial and tangential joint control.

[0023] The present invention is also characterized in that:

[0024] The calculation formula for step 5 is:

[0025]

[0026] Among them, arctan2(*,*) is the inverse tangent calculation function, and μ is the earth's gravitational constant.

[0027] The calculation formula for step 7 is:

[0028]

[0029] The calculation formula for step 8 is:

[0030]

[0031] The specific calculation process of step 9 is:

[0032] When the radial cutting joint control intermediate moment position is not at the perigee or apogee, that is, sinf t ≠0,

[0033]

[0034] When the radial cutting joint control intermediate moment position is at the perigee or apogee, that is, sinf t = 0,

[0035]

[0036] The calculation formula for step 10 is:

[0037]

[0038] The calculation formula for step 11 is:

[0039]

[0040] Where M is the transformation matrix from the satellite radial method coordinate system to the J2000.0 coordinate system at the intermediate moment of radial joint control.

[0041] The calculation formula for step 12 is:

[0042]

[0043] in, According to the satellite orbit time T t ,Location speed Function to calculate the elements of a satellite's orbit.

[0044] The calculation process of the comprehensive deviation J between the temporary orbit and the controlled satellite orbit in step 13 is as follows: When the orbital elements are used for evaluation, the comprehensive deviation J is

[0045] J=λ a |a t -a e |+λ e |e t -e e |+λ i |i t -i e |+λ Ω |Ω t -Ω e |+λ ω ω t -ω e |+λ M |M t -M e |

[0046] where λ a , e , i , Ω , ω , M They are the calculation weight coefficients of semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of perigee and mean anomaly, which can be selected manually as needed;

[0047] When position and speed are used for evaluation, the comprehensive deviation where λ r , v They are the calculation weight coefficients for position and speed, which can be selected manually as needed.

[0048] The calculation formula for step 15 is:

[0049]

[0050] where Δv rs is the theoretical speed increment for radial control, Δv ts Theoretical velocity increment for tangential control, k rs is the radial thruster coefficient used in this control, k ts is the tangential thruster coefficient used in this control, Δv rs , Δvts , k rs and k ts Obtained from the radial cutting joint control track change control parameters.

[0051] The radial-tangential joint control decoupling iterative calibration method based on orbital perturbation model inversion of the present invention calculates radial velocity increments by inverting the satellite orbits and orbital perturbation models before and after control, thereby realizing synchronous decoupling calibration of radial thrusters and tangential thrusters. Considering the limited calculation accuracy of the analytical model, numerical iteration is used in the calibration to ensure the effectiveness of the calibration method, greatly improve the orbital control accuracy, and the method has high reliability, strong operability, and is easy to promote and use. BRIEF DESCRIPTION OF THE DRAWINGS

[0052] Figure 1 It is a flow chart of the radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion of the present invention. DETAILED DESCRIPTION

[0053] The present invention is described in detail below with reference to the accompanying drawings and specific embodiments.

[0054] In the following embodiment, a satellite A performs formation flight around a satellite B, and the satellite A adopts the radial joint control mode when maintaining control of the formation configuration.

[0055] Embodiment 1:

[0056] The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion of the present invention is as follows: Figure 1 When the radial cutting joint control intermediate moment position is not at the perigee or apogee, the specific implementation steps are as follows:

[0057] Step 1: Determine the satellite orbit before the intermediate time of radial joint control. The parameters include the time T of the satellite orbit s , semi-major axis a s , eccentricity e s , inclination angle i s , right ascension of ascending node Ω s , the argument of perigee ω s , mean anomaly angle M s ;

[0058] Step 2: Set the initial value of the radial joint control decoupling iteration calibration, including the time T of the temporary track t =T s , semi-major axis a t =a s , eccentricity e t =e s , inclination angle i t =i s , right ascension of ascending node Ω t =Ω s, the argument of perigee ω t =ω s , mean anomaly angle M t =M s , radial actual control speed increment Δv re =0, tangential actual control speed increment Δv te =0;

[0059] Step 3: Calculate the orbital satellite position vector before the intermediate time control of the radial joint control in the J2000.0 coordinate system and velocity vector

[0060]

[0061] Where F1(t s ,a s ,e s ,i s ,Ω s ,ω s ,M s ) is based on the satellite orbit time T s , semi-major axis a s , eccentricity e s 、Inclination angle i s , right ascension of ascending node Ω s 、Arg of perigee ω s Peace anomaly M s Calculate the satellite position vector and velocity vector Function of

[0062] Step 4: Based on the temporary orbital eccentricity e t Peace anomaly M t , iteratively calculate the orbital anomaly angle E at the intermediate moment of radial joint control t

[0063] E t =M t +e t sin E t ;

[0064] Step 5: Calculate the temporary orbit true anomaly angle f at the intermediate moment of radial joint control t and the average angular velocity of the orbit n t

[0065]

[0066] Among them, arctan2(*,*) is the inverse tangent calculation function, and μ is the earth's gravitational constant;

[0067] Step 6: Determine the satellite orbit after the intermediate time of radial joint control. The parameters include the time T of the satellite orbit e , semi-major axis a e , eccentricity e e , inclination angle i e , right ascension of ascending node Ω e , the argument of perigee ω e , mean anomaly angle M e ;

[0068] Step 7: Calculate the satellite position vector after the intermediate time of radial joint control in the J2000.0 coordinate system and the velocity vector

[0069]

[0070] Step 8: Calculate the change in the semi-major axis of the radial joint control orbit Δa, the change in eccentricity Δe, and the change in the mean anomaly ΔM

[0071]

[0072] Step 9: Calculate the radial control velocity increment correction δv based on the orbital perturbation model inversion re and the tangential control speed increment correction δv te

[0073]

[0074] Step 10: Calculate the actual radial control speed increment Δv re and the actual tangential control speed increment Δv te

[0075]

[0076] Step 11: Calculate the temporary orbit satellite position vector in the J2000.0 coordinate system and the velocity vector

[0077]

[0078] Where M is the transformation matrix from the satellite radial method coordinate system to the J2000.0 coordinate system at the intermediate moment of radial joint control;

[0079] Step 12: Based on the time T of the satellite's temporary orbit t ,Location speed Calculate the temporary orbit elements of the satellite

[0080]

[0081] in, It is based on the satellite orbit time T t ,Location speed Functions for calculating satellite orbit elements;

[0082] Step 13: Calculate the comprehensive deviation J between the temporary orbit and the controlled satellite orbit, and use the position and velocity to evaluate the comprehensive deviation where λ r , v They are the calculation weight coefficients of position and speed, which can be selected manually as needed;

[0083] Step 14: Determine whether the comprehensive deviation J meets the accuracy requirement. If it does, that is, J<δ, where δ is the iterative calculation threshold, which can be manually selected as needed, then proceed to step 15, otherwise go to step 4;

[0084] Step 15: Decouple and calibrate radial thruster and tangential thruster coefficients

[0085]

[0086] where Δv rs is the theoretical speed increment for radial control, Δv ts Theoretical velocity increment for tangential control, k rs is the radial thruster coefficient used in this control, k ts is the tangential thruster coefficient used in this control, Δv rs , Δv ts , k rs and k ts Obtained from the radial cutting joint control track change control parameters.

[0087] Example 2

[0088] When the intermediate moment position of radial cutting joint control is at the perigee or apogee, the specific implementation steps are as follows:

[0089] Step 1: Determine the satellite orbit before the intermediate time of radial joint control. The parameters include the time T of the satellite orbit s , semi-major axis a s , eccentricity e s , inclination angle i s , right ascension of ascending node Ω s , the argument of perigee ω s , mean anomaly angle M s ;

[0090] Step 2: Set the initial value of the radial joint control decoupling iteration calibration, including the time T of the temporary track t =T s , semi-major axis a t =as , eccentricity e t =e s , inclination angle i t =i s , right ascension of ascending node Ω t =Ω s , the argument of perigee ω t =ω s , mean anomaly angle M t =M s , radial actual control speed increment Δv re =0, tangential actual control speed increment Δv te =0;

[0091] Step 3: Calculate the orbital satellite position vector before the intermediate time control of the radial joint control in the J2000.0 coordinate system and velocity vector

[0092]

[0093] Where F1(t s ,a s ,e s ,i s ,Ω s ,ω s ,M s ) is based on the satellite orbit time T s , semi-major axis a s , eccentricity e s 、Inclination angle i s , right ascension of ascending node Ω s 、Arg of perigee ω s Peace anomaly M s Calculate the satellite position vector and velocity vector Function of

[0094] Step 4: Based on the temporary orbital eccentricity e t Peace anomaly M t , iteratively calculate the orbital anomaly angle E at the intermediate moment of radial joint control t

[0095] E t =M t +e t sinE t ;

[0096] Step 5: Calculate the temporary orbit true anomaly angle f at the intermediate moment of radial joint control t and the average angular velocity of the orbit n t

[0097]

[0098] Among them, arctan2(*,*) is the inverse tangent calculation function, and μ is the earth's gravitational constant;

[0099] Step 6: Determine the satellite orbit after the intermediate time of radial joint control. The parameters include the time T of the satellite orbit e , semi-major axis a e , eccentricity e e , inclination angle i e , right ascension of ascending node Ω e , the argument of perigee ω e , mean anomaly angle M e ;

[0100] Step 7: Calculate the satellite position vector after the intermediate time of radial joint control in the J2000.0 coordinate system and the velocity vector

[0101]

[0102] Step 8: Calculate the change in the semi-major axis of the radial joint control orbit Δa, the change in eccentricity Δe, and the change in the mean anomaly ΔM

[0103]

[0104] Step 9: Calculate the radial control velocity increment correction δv based on the orbital perturbation model inversion re and the tangential control speed increment correction δv te

[0105] When the radial cutting joint control intermediate moment position is not at the perigee or apogee, that is, sinf t ≠0 hours

[0106]

[0107] When the radial cutting joint control intermediate moment position is at the perigee or apogee, that is, sinf t =0 o'clock

[0108]

[0109] Step 10: Calculate the actual radial control speed increment Δv re and the actual tangential control speed increment Δv te

[0110]

[0111] Step 11: Calculate the temporary orbit satellite position vector in the J2000.0 coordinate system and the velocity vector

[0112] Where M is the transformation matrix from the satellite radial method coordinate system to the J2000.0 coordinate system at the intermediate moment of radial joint control;

[0113] Step 12: Based on the time T of the satellite's temporary orbit t ,Location speed Calculate the temporary orbit elements of the satellite

[0114]

[0115] in, It is based on the satellite orbit time T t ,Location speed Functions for calculating satellite orbit elements;

[0116] Step 13: Calculate the comprehensive deviation J between the temporary orbit and the controlled satellite orbit. When the orbital elements are used for evaluation, the comprehensive deviation J is

[0117] J=λ a |a t -a e |+λ e |e t -e e |+λ i |i t -i e |+λ Ω |Ω t -Ω e |+λ ω |ω t -ω e |+λ M |M t -M e |, where λ a ,

[0118] λ e , i , Ω , ω , M They are the calculation weight coefficients of semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of perigee and mean anomaly, which can be selected manually as needed;

[0119] When position and speed are used for evaluation, the comprehensive deviation where λ r , v They are the calculation weight coefficients of position and speed, which can be selected manually as needed;

[0120] Step 14: Determine whether the comprehensive deviation J meets the accuracy requirement. If it does, that is, J<δ, where δ is the iterative calculation threshold, which can be manually selected as needed, then proceed to step 15, otherwise go to step 4;

[0121] Step 15: Decouple and calibrate radial thruster and tangential thruster coefficients

[0122]

[0123] where Δv rs is the theoretical speed increment for radial control, Δv ts Theoretical velocity increment for tangential control, k rs is the radial thruster coefficient used in this control, k ts is the tangential thruster coefficient used in this control, Δv rs , Δv ts , k rs and k ts Obtained from the radial cutting joint control track change control parameters.

Claims

1. A radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion, characterized in that: Follow these steps to implement: Step 1: Determine the parameters of the satellite orbit before the intermediate time of radial joint control, including the time T of the satellite orbit s , semi-major axis a s , eccentricity e s 、Inclination angle i s , right ascension of ascending node Ω s 、Arg of perigee ω s 、Mean anomaly angle M s ; Step 2: Set the initial value of the radial joint control decoupling iteration calibration, including the time T of the temporary track t =T s , semi-major axis a t =a s , eccentricity e t =e s 、Inclination angle i t =i s , right ascension of ascending node Ω t =Ω s 、Arg of perigee ω t =ω s 、Mean anomaly angle M t =M s , radial actual control speed increment Δv re =0, tangential actual control speed increment Δv te =0; Step 3: Calculate the orbital satellite position vector before the intermediate time control of the radial joint control in the J2000.0 coordinate system and the velocity vector Where F1(t s ,a s ,e s ,i s ,Ω s ,ω s ,M s ) is based on the satellite orbit time T s , semi-major axis a s , eccentricity e s 、Inclination angle i s , right ascension of ascending node Ω s 、Arg of perigee ω s Peace anomaly M s Calculate the satellite position vector and the velocity vector Function of Step 4: Based on the temporary orbital eccentricity e t Peace anomaly M t , iteratively calculate the orbital anomaly angle E at the intermediate moment of radial joint control t It is t =M t +e t Himself t ; Step 5: Calculate the temporary orbit true anomaly angle f at the intermediate moment of radial joint control t and the average angular velocity of the orbit n t ; Step 6: Determine the satellite orbit after the intermediate time of the radial joint control. The parameters include the time T of the satellite orbit e , semi-major axis a e , eccentricity e e , inclination angle i e , right ascension of ascending node Ω e , the argument of perigee ω e , mean anomaly angle M e ; Step 7: Calculate the satellite position vector after the intermediate time of radial joint control in the J2000.0 coordinate system and the velocity vector Step 8: Calculate the change of the semi-major axis Δa, the change of the eccentricity Δe and the change of the mean anomaly ΔM of the radial joint control orbit; Step 9: Calculate the radial control velocity increment correction δv based on the orbital perturbation model inversion re and the tangential control speed increment correction δv te ; Step 10: Calculate the actual radial control speed increment Δv re and the actual tangential control speed increment Δv te ; Step 11: Calculate the temporary orbit satellite position vector in the J2000.0 coordinate system and the velocity vector Step 12: Based on the time T of the satellite's temporary orbit t ,Location speed Calculate the temporary orbit elements of the satellite; Step 13: Calculate the comprehensive deviation J between the temporary orbit and the controlled satellite orbit; Step 14: Determine whether the comprehensive deviation J meets the accuracy requirement. If it does, that is, J<δ, where δ is the iterative calculation threshold, which can be manually selected as needed, then proceed to step 15, otherwise go to step 4; Step 15: Decouple and calibrate the radial thruster and tangential thruster coefficients, and then realize the decoupled iterative calibration of radial and tangential joint control.

2. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 5 is: Among them, arctan2(*,*) is the inverse tangent calculation function, and μ is the earth's gravitational constant.

3. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 7 is:

4. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 8 is:

5. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The specific calculation process of step 9 is: When the radial cutting joint control intermediate moment position is not at the perigee or apogee, that is, sinf t ≠0, When the radial cutting joint control intermediate moment position is at the perigee or apogee, that is, sinf t = 0, 6. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 10 is:

7. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 11 is: Where M is the transformation matrix from the satellite radial method coordinate system to the J2000.0 coordinate system at the intermediate moment of radial joint control.

8. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 12 is: in, According to the satellite orbit time T t ,Location speed Function to calculate the elements of a satellite's orbit.

9. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation process of the comprehensive deviation J between the temporary orbit and the controlled satellite orbit in step 13 is as follows: When the orbit elements are used for evaluation, the comprehensive deviation J is J=λ a |a t -a e |+λ e |e t -e e |+λ i |i t -i e |+λ Ω |Oh t -Oh e |+λ ω |oh t -oh e |+λ M |M t -M e | where λ a , e , i , Ω , ω , M They are the calculation weight coefficients of semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of perigee and mean anomaly, which can be selected manually as needed; When position and speed are used for evaluation, the comprehensive deviation where λ r , v They are the calculation weight coefficients for position and speed, which can be selected manually as needed.

10. The radial shear joint control decoupling iterative calibration method based on orbital perturbation model inversion according to claim 1 is characterized in that: The calculation formula of step 15 is: where Δv rs is the theoretical speed increment for radial control, Δv ts Theoretical velocity increment for tangential control, k rs is the radial thruster coefficient used in this control, k ts is the tangential thruster coefficient used in this control, Δv rs , Δv ts , k rs and k ts Obtained from the radial-cut joint control track change control parameters.

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