Satellite strict regression orbit accurate maintenance control method based on micro electric propulsion

Abstract

The application discloses a satellite strict regression orbit accurate maintenance control method based on micro electric propulsion, which is based on an electric propulsion system using micro thrust as an execution mechanism of orbit control, utilizes the characteristics that micro thrust has little interference to a satellite platform, executes orbit control tasks at any time, and timely corrects orbit deviation without affecting the normal work of the satellite platform. The current maintenance control method can be applied to long-time orbit maintenance control, realizes timely correction of high-precision satellite orbit deviation without affecting the normal work of the satellite platform, and the orbit maintenance control precision of the whole orbit can reach a level of 100 meters or even higher for the strict regression orbit maintenance control task.

Description

Technical Field

[0001] This invention relates to a precise control method for maintaining a satellite's strict return orbit based on micro-electric propulsion, belonging to the field of satellite orbit control technology. Background Technology

[0002] With the development of satellite orbit control technology, many spacecraft payloads have increasingly higher precision requirements for their target orbits. Traditional orbit holding methods mainly rely on chemical propulsion systems, but these systems have drawbacks such as high thrust, low precision, and low specific impulse. These drawbacks result in significant interference with the stable operation of the satellite platform during propulsion system operation. Therefore, orbit holding control tasks need to be processed in a time-sharing manner with payload operation tasks, which limits the available time for orbit holding control. If orbit control cannot be performed in a timely manner, it will cause a significant increase in orbital deviation. On the other hand, the thrust calculation of chemical thrusters is affected by temperature, pressure, and other factors, and the deviation between the theoretical calculation value and the actual effect is large. This results in a relatively large deviation from the target state after each control operation. Therefore, satellites using chemical propulsion typically achieve lower orbit holding control precision. In addition, the low specific impulse of chemical propulsion further limits its use in missions requiring large orbit holding control parameters for long-life satellites. In contrast, electric propulsion systems generally have thrust in the mN range, which is considered micro-thrust. Such thrust causes very little interference to satellites or spacecraft. Moreover, whether it is Hall-type or ion-type electric propulsion, its specific impulse is at least 5 to 10 times that of chemical propulsion, which is difficult to meet the requirements of satellite high-precision, long-term, and accurate orbital maintenance and control missions. Summary of the Invention

[0003] The technical problem solved by this invention is that, in the current technology, it is difficult to meet the requirements of high-precision, long-term accurate orbit maintenance control for satellites. Therefore, this invention proposes a method for precise orbit maintenance control of satellites based on micro-electric propulsion.

[0004] The present invention solves the above-mentioned technical problem through the following technical solution: A precise control method for maintaining a satellite's strict return orbit based on micro-electric propulsion includes: Obtain the time start point, three-dimensional position, and three-dimensional velocity sequence information of the strict regression reference orbit; Based on the selected forward strict regression orbit reference point closest to the on-board time and its corresponding orbital time position and velocity, calculate the orbital six-root number of the forward strict regression orbit reference point; Perform orbit recursion to calculate the orbital six-root number of the satellite at the position point corresponding to the selected orbital time. Based on the number of six orbital elements at the forward strict regression orbit reference point and the number of six orbital elements of the orbit to be controlled at the position point corresponding to the selected orbit time, the deviation of the number of orbital elements to be controlled is calculated and inertial filtering is performed to obtain the amount of deviation of the current orbital elements to be corrected, and the number of six orbital elements of the orbit to be controlled is corrected. The average rate of change and the corresponding thruster start-up control time are determined based on the number of six orbital elements of the corrected orbit to be controlled. The correspondence between the three-axis thrust and the orbital coordinate system is obtained by using the installation relationship of the three-axis thrusters between satellites, and the radial, circumferential and normal micro-propulsion control accelerations are calculated. Numerical integration is performed on the radial, circumferential, and normal micro-propulsion control accelerations, as well as the average rate of change of the six orbital elements, to obtain the correction amount of the thruster working time length in each direction of the track to be controlled. Based on the radial, circumferential and normal micro-propulsion control accelerations, and the correction amount of the thruster working time length in each direction of the track to be controlled, the working time of the thruster in each direction is obtained according to the principle of minimizing the total velocity increment modulus and value of the three axes. Perform control torque feedforward compensation within the thruster control cycle to obtain the thruster feedforward compensation torque in each direction.

[0005] The starting point of the strictly regressive reference orbit is set to t0. Utilizing the orbital characteristic of a regression period of T, the distance t from the satellite is found by consulting the orbital reference point sequence table. i The most recent forward strict regression orbit reference point P C And determine the corresponding orbital time T. C Location [x] c ,y c ,z c ] and speed [vx c ,vy c ,vz c ], used to calculate the orbital root number of the forward strict regression orbit reference point [a c ,e c i c , Ω c ,ω c M c ].

[0006] Based on the orbital six-root number [a] of the forward strict regression orbital reference point c ,e c i c , Ω c ,ω c M c Perform orbit recursion to obtain the satellite's time T in the selected orbit. C The location point P corresponding to the time. S The orbital six-root number [as ,e s i s , Ω s ,ω s M s The method for calculating the deviation of the track elements to be controlled based on the obtained parameters is as follows:

[0007]

[0008]

[0009]

[0010]

[0011]

[0012] In the formula, : Track major semi-axis, Track inclination, Eccentricity , Ascending node right ascension, :Aspect of perigee : Average anomaly; subscript c represents the orbital element number corresponding to the reference orbital point; subscript s represents the current satellite's orbital element number.

[0013] The method for smoothing the noise of the track element deviation using inertial filtering is as follows:

[0014] In the formula, This represents the amount of deviation in the current shooting trajectory that needs to be corrected. The calculation amount of the current control track element deviation is related to... Related, These are the selected filter parameters.

[0015] The correspondence between the three-axis thrusters and the orbital coordinate system is obtained by utilizing the installation relationship of the three-axis thrusters between satellites. The radial, circumferential, and normal micro-propulsion control accelerations S, T, and W are calculated, respectively. The average rate of change and micro-propulsion control accelerations S, T, and W within the control time period are determined using the six orbital elements of the corrected orbit. The method for obtaining the correction amount of the thruster working time length in each direction of the orbit to be controlled by numerical integration is as follows: ; ; ; ; ;

[0016] In the formula, For the track semi-major, is the Earth's gravitational constant.

[0017] Before calculating the correction amount of the thruster working time length in each direction of the track to be controlled, an overshoot is introduced to limit the control accuracy of the six track members of the track to be corrected. The relationship between the overshoot and the control accuracy of the six track members of the track to be corrected is: the overshoot cannot exceed m times the control accuracy range. The value of m is set based on prior historical data and the requirements of the micro-electric propulsion mission.

[0018] The method for controlling the accuracy of the six track elements in the overshoot correction track is as follows: Set the control precision to The rate of change is f, if ,but: ,like ,but: ; After obtaining the correction amount of the working time length of the thrusters in each direction of the track to be controlled, the working time of the thrusters in each direction is obtained according to the principle of minimizing the sum of the three-axis total velocity increment modulus. The principle for minimizing the sum of the total velocity increment moduli of the three axes is to use the sum of the total velocity increment moduli of the three axes. Minimum.

[0019] The control torque feedforward compensation during the thruster control cycle is used to eliminate disturbing torques during thruster operation via open-loop control. The method for thruster feedforward compensation is as follows: Obtain the angular momentum in that direction at the start of thruster operation; Determine the thruster's current jet duration in the direction of operation and the target angular momentum in that direction at the end of the thruster's operation. Calculate the thruster's feedforward compensation torque using the following method:

[0020] In the formula: This represents the angular momentum in this direction at the start of thruster operation without feedforward compensation. Let be the angular momentum in that direction at the moment the thruster finishes its operation. This refers to the duration of jet propulsion in that direction; After obtaining the thruster feedforward compensation torque through reverse propulsion, a simulation test is performed to determine whether the preset angular momentum in that direction can be output at the end of the thruster's operation. Based on the determination result, the calculated thruster feedforward compensation torque is adjusted.

[0021] After obtaining the feedforward compensation torque of each thruster, orbit control is executed within the thruster control cycle according to the calculated thruster operating time. During the orbit control process, the angular momentum of the actuator in the direction of the disturbance torque is monitored. When the accumulated angular momentum exceeds the maximum control angular momentum in the current direction, the control is initiated. Half of the angular momentum is controlled; suspend orbit control until the effect of angular momentum is eliminated, then resume the corresponding orbit control; maximum control angular momentum. The settings are based on prior historical data and the requirements of the micro-electric propulsion mission.

[0022] The advantages of this invention compared to the prior art are: (1) The present invention provides a precise control method for satellite strict return orbit based on micro electric propulsion. The micro-thrust electric propulsion system is used as the actuator for orbit control. On the one hand, it takes advantage of the small interference of micro thrust on the satellite platform, and can perform orbit control tasks at any time without affecting the normal operation of the satellite platform, so as to correct the orbit deviation in a timely manner. On the other hand, compared with chemical propellants, electric propulsion has a significantly increased specific impulse, strong orbit control capability, and is more suitable for long-term orbit holding control. (2) In view of the problem of determining the pipeline deviation of the strict regression orbit, the present invention adopts the method of strictly aligning the orbit parameters of the strict regression reference orbit point with the satellite orbit parameters in time, obtaining the actual satellite orbit parameters to calculate the deviation, and then using inertial filtering to smooth the calculation noise, so as to obtain the orbit parameter deviation to be corrected that can be used for control. (3) This invention uses the average rate of change of orbital elements to calculate the thruster start-up control time. Previously, the system used a micro-thruster with millinewton-level thrust for high-precision orbital maintenance control, resulting in a long control time. The early optimal pulse control calculation method based on infinite thrust is no longer applicable to the current scenario. Therefore, a control calculation method based on minimizing the three-axis velocity increment modulus is presented. Simultaneously, considering the impact of orbital control interference on the satellite's attitude, it is necessary to segment the long-term control operation when necessary to ensure satellite attitude stability. Attached Figure Description

[0023] Figure 1 Block diagram of a satellite strict return-to-orbit control system based on micro-electric propulsion provided for this invention; Figure 2 This is a schematic diagram of the thruster installation direction in an embodiment of the present invention. Detailed Implementation

[0024] A precise orbit-keeping control method for satellites based on micro-electric propulsion is proposed. This method utilizes a micro-thrust electric propulsion system as the orbit control actuator, leveraging the low interference of micro-thrust on the satellite platform to perform orbit control tasks continuously and correct orbital deviations in a timely manner without affecting the normal operation of the satellite platform. This current orbit-keeping control method is applicable to long-term orbit-keeping control, achieving high-precision timely correction of satellite orbital deviations without affecting the normal operation of the satellite platform. For strict orbit-keeping control tasks, its orbit-keeping control accuracy across the entire orbit can reach hundreds of meters or even higher.

[0025] The satellite's precise orbit-holding control method based on micro-electric propulsion, with specific control steps designed as follows: Obtain the time start point, three-dimensional position, and three-dimensional velocity sequence information of the strict regression reference orbit; Based on the selected forward strict regression orbit reference point closest to the on-board time and its corresponding orbital time position and velocity, calculate the orbital six-root number of the forward strict regression orbit reference point; Perform orbit recursion to calculate the orbital six-root number of the satellite at the position point corresponding to the selected orbital time. Based on the number of six orbital elements at the forward strict regression orbit reference point and the number of six orbital elements of the orbit to be controlled at the position point corresponding to the selected orbit time, the deviation of the number of orbital elements to be controlled is calculated and inertial filtering is performed to obtain the amount of deviation of the current orbital elements to be corrected, and the number of six orbital elements of the orbit to be controlled is corrected. The average rate of change and the corresponding thruster start-up control time are determined based on the number of six orbital elements of the corrected orbit to be controlled. The correspondence between the three-axis thrust and the orbital coordinate system is obtained by using the installation relationship of the three-axis thrusters between satellites, and the radial, circumferential and normal micro-propulsion control accelerations are calculated. Numerical integration is performed on the radial, circumferential, and normal micro-propulsion control accelerations, as well as the average rate of change of the six orbital elements, to obtain the correction amount of the thruster working time length in each direction of the track to be controlled. Based on the radial, circumferential and normal micro-propulsion control accelerations, and the correction amount of the thruster working time length in each direction of the track to be controlled, the working time of the thruster in each direction is obtained according to the principle of minimizing the total velocity increment modulus and value of the three axes. Perform control torque feedforward compensation within the thruster control cycle to obtain the thruster feedforward compensation torque in each direction.

[0026] The starting time of the strict regression reference orbit is set to t0. Utilizing the orbital characteristics of a strict regression reference orbit with a regression period of T, the distance t from the satellite is found by consulting the orbital reference point sequence table. i The most recent forward strict regression orbit reference point P C And determine the corresponding orbital time T.C Location [x] c ,y c ,z c ] and speed [vx c ,vy c ,vz c ], used to calculate the orbital root number of the forward strict regression orbit reference point [a c ,e c i c , Ω c ,ω c M c ].

[0027] Based on the orbital six-root number [a] of the forward strict regression orbital reference point c ,e c i c , Ω c ,ω c M c Perform orbit recursion to obtain the satellite's time T in the selected orbit. C The location point P corresponding to the time. S The orbital six-root number [a s ,e s i s , Ω s ,ω s M s The method for calculating the deviation of the track elements to be controlled based on the obtained parameters is as follows:

[0028]

[0029]

[0030]

[0031]

[0032]

[0033] In the formula, : Track major semi-axis, Track inclination, Eccentricity , Ascending node right ascension, :Aspect of perigee : Average anomaly; subscript c represents the orbital element number corresponding to the reference orbital point; subscript s represents the current satellite's orbital element number.

[0034] The method for smoothing the noise of the track element deviation using inertial filtering is as follows:

[0035] In the formula, This represents the amount of deviation in the current shooting trajectory that needs to be corrected. The calculation amount of the current control track element deviation is related to... Related, These are the selected filter parameters.

[0036] The correspondence between the three-axis thrusters and the orbital coordinate system is obtained by utilizing the installation relationship of the three-axis thrusters between satellites. The radial, circumferential, and normal micro-propulsion control accelerations S, T, and W are calculated, respectively. The average rate of change and micro-propulsion control accelerations S, T, and W within the control time period are determined using the six orbital elements of the corrected orbit. The method for obtaining the correction amount of the thruster working time length in each direction of the orbit to be controlled by numerical integration is as follows: ; ; ; ; ;

[0037] In the formula, For the track semi-major, is the Earth's gravitational constant; the definitions of other orbital parameters are the same as before.

[0038] Before calculating the correction amount of the thruster working time length in each direction of the track to be controlled, an overshoot is introduced to limit the control accuracy of the six track members of the track to be corrected. The relationship between the overshoot and the control accuracy of the six track members of the track to be corrected is: the overshoot cannot exceed m times the control accuracy range. The value of m is set based on prior historical data and the requirements of the micro-electric propulsion mission.

[0039] The method for controlling the accuracy of the six track elements in the track with overshoot to be corrected is as follows: Set the control precision to The rate of change is f, if ,but: ,like ,but: ; After obtaining the correction amount of the working time length of the thrusters in each direction of the track to be controlled, the working time of the thrusters in each direction is obtained according to the principle of minimizing the sum of the three-axis total velocity increment modulus. The principle for minimizing the sum of the total velocity increment moduli of the three axes is to use the sum of the total velocity increment moduli of the three axes. Minimum.

[0040] The control torque feedforward compensation during the thruster control cycle is used to eliminate disturbing torques during thruster operation via open-loop control. The method for thruster feedforward compensation is as follows: Obtain the angular momentum in that direction at the start of thruster operation; Determine the thruster's current jet duration in the direction of operation and the target angular momentum in that direction at the end of the thruster's operation. Calculate the thruster's feedforward compensation torque using the following method:

[0041] In the formula: This represents the angular momentum in this direction at the start of thruster operation without feedforward compensation. Let be the angular momentum in that direction at the moment the thruster finishes its operation. This refers to the duration of jet propulsion in that direction; After obtaining the thruster feedforward compensation torque through reverse propulsion, a simulation test is performed to determine whether the preset angular momentum in that direction can be output at the end of the thruster's operation. Based on the determination result, the calculated thruster feedforward compensation torque is adjusted.

[0042] After obtaining the feedforward compensation torque of each thruster, orbit control is executed within the thruster control cycle according to the calculated thruster operating time. During the orbit control process, the angular momentum of the actuator in the direction of the disturbance torque is monitored. When the accumulated angular momentum exceeds the maximum control angular momentum in the current direction, the control is initiated. Half of the angular momentum is controlled; suspend orbit control until the effect of angular momentum is eliminated, then resume the corresponding orbit control; maximum control angular momentum. The settings are based on prior historical data and the requirements of the micro-electric propulsion mission.

[0043] The following description, in conjunction with the accompanying drawings and preferred embodiments, provides further details: In the current embodiment, the satellite strict return orbit precise maintenance control method based on micro-electric propulsion uses a micro-thrust electric propulsion system as the orbit control actuator. Leveraging the low interference of micro-thrust on the satellite platform, orbit control tasks can be performed at any time without affecting the normal operation of the satellite platform, thus enabling timely correction of orbital deviations. Furthermore, compared to chemical propellants, electric propulsion significantly increases specific impulse, resulting in stronger orbit control capabilities and making it more suitable for long-term orbit maintenance control.

[0044] By utilizing the characteristics of a strictly regressive reference orbit with a regression period of T, the distance t from the satellite can be found by consulting the orbital reference point sequence table. i The most recent forward strict regression orbit reference point P C By using orbital recursion and time alignment, the orbital element deviation can be calculated as follows:

[0045]

[0046]

[0047]

[0048]

[0049]

[0050] Since this deviation has short-term stability, inertial filtering can be used to smooth the noise of the orbital element deviation, that is:

[0051] The amount to be corrected for the deviation of the orbital elements.

[0052] During the calculation process, a certain overshoot needs to be considered based on the track parameters to be corrected and their rate of change. The overshoot should not exceed m times the control accuracy range of the parameter (m can be adjusted according to actual needs), using the semi-major axis control quantity. For example, if its control precision is The rate of change is f, if ,but: ,like ,but: .

[0053] By utilizing the relationship between the mean rate of change of orbital elements of a near-circular orbital satellite and the radial, circumferential, and normal micro-propulsion control accelerations S, T, and W, the thruster operating time in each direction required to correct orbital parameter errors is obtained through numerical integration. ; ; ; ; ;

[0054] And the sum of the total velocity increment moduli of the three axes Based on the principle of minimizing costs, the required thruster operating time is calculated. The overall satellite strict return-to-orbit maintenance control block diagram is designed as follows: Figure 1 As shown.

[0055] For handling micro-propulsion disturbances: If the disturbance torque is too large, attitude feedforward compensation control is required. The control torque is calculated as follows:

[0056] If the angular momentum in the direction of the disturbance torque gradually accumulates and exceeds the maximum control angular momentum in that direction... Half of (i.e.: If this is the case, control in that direction needs to be paused until the influence of angular momentum is eliminated, and then the corresponding orbit control can be resumed. This is equivalent to breaking down a single control operation into a control sequence.

[0057] Example 1: This embodiment proposes a precise control method for maintaining a satellite's strict return orbit based on micro-electric propulsion. This method is implemented using a precise control system for maintaining a satellite's strict return orbit, which includes modules such as a GNSS receiver, navigation information processing and control calculations, satellite attitude control, and a micro-electric propulsion system. The GNSS receiver module is responsible for measuring real-time orbital position and velocity information. The navigation information processing and control calculation module is responsible for calculating pipeline deviation and track control strategies. The satellite attitude control module ensures the stability of the satellite's attitude during normal operation and orbit control. The micro-electric propulsion system module, including propellant, storage and supply system and thruster, can provide triaxial thrust (F... x F y F z ).

[0058] S1: Based on the time start point t0, three-dimensional position, and three-dimensional velocity sequence information of the strictly regressed reference orbit, and utilizing the characteristics of an orbit with a regression period of T, the distance from the satellite at time t can be found by consulting the orbit reference point sequence table. i The most recent forward strict regression orbit reference point P C The corresponding orbital time T C Location [x] c ,y c ,z c ] and speed [vx c ,vy c ,vz c ], and from this, the number of orbital roots of the reference point is calculated [a c ,e c i c , Ω c ,ω c M c And by recursively calculating the satellite's position at T... C Time P S The orbital number of the point [a] s ,e s i s , Ω s ,ω s Ms Therefore, the calculated amount of the track element deviation to be controlled can be obtained as follows:

[0059]

[0060]

[0061]

[0062]

[0063]

[0064] Since this deviation has short-term stability, inertial filtering can be used to smooth the noise of the orbital element deviation, that is:

[0065] in, This represents the amount of deviation in the current shooting trajectory that needs to be corrected. The computational complexity of the current orbital elements. These are the selected filter parameters.

[0066] S2: Due to the use of micro-propulsion for high-precision track maintenance control, the control time is relatively long. Therefore, the thruster start-up control time needs to be calculated based on the average rate of change of the track elements. The correspondence between the three-axis thrust and the track coordinate system is obtained using the installation relationship. The thruster installation direction is as follows: Figure 2 As shown. The radial, circumferential, and normal micro-propulsion control accelerations S, T, and W are obtained. Based on the relationship between the mean rate of change of orbital elements and acceleration for a near-circular orbital satellite, the thruster operating time in each direction required to correct the aforementioned orbital parameter errors is obtained through numerical integration. ; ; ; ; ;

[0067] in: For the track semi-major, is the Earth's gravitational constant.

[0068] S3: In the actual calculation process, it is also necessary to consider a certain overshoot based on the track parameters to be corrected and their rate of change. The overshoot should not exceed m times the control accuracy range of the parameter (m can be adjusted according to actual needs), using the semi-major axis control quantity. For example, if its control precision is The rate of change is f, if ,but: ,like ,but: After calculating the actual control quantities, the track parameter correction requirements and their influence from accelerations in different directions are determined by the sum of the three-axis total velocity increment moduli. Based on the principle of minimizing the time required for thruster operation, calculate the required operating time.

[0069] S4: During the thrust operation of the micro-propulsion system, disturbance torque will still be generated. On the one hand, if the disturbance torque is too large, it will affect the steady-state accuracy of the celestial body's attitude, requiring attitude feedforward compensation control. This control is an open-loop control, which means that the disturbance torque can be estimated based on the change in angular momentum in that direction without feedforward compensation, and the control torque feedforward compensation will be activated during the next control period to counteract the influence of the jet disturbance torque.

[0070] in: This represents the angular momentum in this direction at the start of thruster operation without feedforward compensation. Let be the angular momentum in that direction at the moment the thruster finishes its operation. This represents the duration of jet propulsion in that direction.

[0071] On the other hand, even if the attitude disturbance torque is small and the celestial body's attitude can meet the requirements, over time, the angular momentum of the actuators used for attitude control will gradually accumulate in the direction of the disturbance torque. If the accumulated amount exceeds the maximum control angular momentum in that direction... Half of (i.e.: If this happens, control needs to be paused, and the corresponding orbital control will continue to be executed after the influence of angular momentum is eliminated.

[0072] This embodiment discloses a method and system for precise orbit maintenance control of satellites based on micro-electric propulsion. The proposed system construction idea and control method have clear principles and well-defined processes, fully leveraging the advantages of low interference torque and high specific impulse of the micro-thrust electric propulsion system. It can achieve three-axis omnidirectional thrust control in the radial (S), circumferential (T), and normal (W) directions without affecting the normal operation of the satellite platform. It possesses good engineering feasibility and effectively solves the problems of large control deviations, significant interference with satellite attitude, and untimely control maintenance in orbit control achieved by traditional chemical propulsion. By selecting an appropriate electric propulsion thrust, such as 10mN or even less, high-precision orbit control can be achieved. Furthermore, due to the high specific impulse of the electric propulsion system, the available total control impulse is large, making it more suitable for orbit maintenance control during long-term satellite operation. This method and system are applicable to high-precision, high-stability satellite orbit control applications.

[0073] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

[0074] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for precise control of a satellite's strict return orbit based on micro-electric propulsion, characterized in that... include: Obtain the time start point, three-dimensional position, and three-dimensional velocity sequence information of the strict regression reference orbit; Based on the selected forward strict regression orbit reference point closest to the on-board time and its corresponding orbital time position and velocity, calculate the orbital six-root number of the forward strict regression orbit reference point; Perform orbit recursion to calculate the orbital six-root number of the satellite at the position point corresponding to the selected orbital time. Based on the number of six orbital elements at the forward strict regression orbit reference point and the number of six orbital elements of the orbit to be controlled at the position point corresponding to the selected orbit time, the deviation of the number of orbital elements to be controlled is calculated and inertial filtering is performed to obtain the amount of deviation of the current orbital elements to be corrected, and the number of six orbital elements of the orbit to be controlled is corrected. The average rate of change and the corresponding thruster start-up control time are determined based on the number of six orbital elements of the corrected orbit to be controlled. The correspondence between the three-axis thrust and the orbital coordinate system is obtained by using the installation relationship of the three-axis thrusters between satellites, and the radial, circumferential and normal micro-propulsion control accelerations are calculated. Numerical integration is performed on the radial, circumferential, and normal micro-propulsion control accelerations, as well as the average rate of change of the six orbital elements, to obtain the correction amount of the thruster working time length in each direction of the track to be controlled. Based on the radial, circumferential and normal micro-propulsion control accelerations, and the correction amount of the thruster working time length in each direction of the track to be controlled, the working time of the thruster in each direction is obtained according to the principle of minimizing the total velocity increment modulus and value of the three axes. Perform control torque feedforward compensation within the thruster control cycle to obtain the thruster feedforward compensation torque in each direction.

2. The method for precise maintenance control of a satellite's strict return orbit based on micro-electric propulsion according to claim 1, characterized in that: The starting point of the strictly regressive reference orbit is set to t0. Utilizing the orbital characteristic of a regression period of T, the distance t from the satellite is found by consulting the orbital reference point sequence table. i The most recent forward strict regression orbit reference point P C And determine the corresponding orbital time T. C Position and velocity V are used to calculate the orbital six-root number of the forward strictly regressed orbit reference point.

3. The method for precise maintenance control of a satellite's strict return orbit based on micro-electric propulsion according to claim 2, characterized in that: Based on the orbital root number [a] of the forward strict regression orbital reference point c ,e c i c , Ω c ,ω c M c Perform orbit recursion to obtain the satellite's time T in the selected orbit. C The location point P corresponding to the time. S The orbital six-root number [a s ,e s i s , Ω s ,ω s M s The method for calculating the deviation of the track elements to be controlled based on the obtained parameters is as follows: In the formula, For the major semi-axis of the track, For the track inclination angle, Eccentricity , Right ascension of the ascending node, The perigee argument, The angle of approach is the mean angle; the subscript 'c' represents the orbital element corresponding to the reference orbit point; the subscript 's' represents the current orbital element of the satellite.

4. The method for precise maintenance control of a satellite's strict return orbit based on micro-electric propulsion according to claim 3, characterized in that: The method for smoothing the noise of the track element deviation using inertial filtering is as follows: In the formula, This represents the amount of deviation in the current shooting trajectory that needs to be corrected. This is the calculation amount for the current track element deviation to be controlled. These are the selected filter parameters.

5. The precise control method for maintaining a satellite's strict return orbit based on micro-electric propulsion according to claim 1, characterized in that: The correspondence between the three-axis thrusters and the orbital coordinate system is obtained by utilizing the installation relationship of the three-axis thrusters between satellites. The radial, circumferential, and normal micro-propulsion control accelerations S, T, and W are calculated, respectively. The average rate of change and micro-propulsion control accelerations S, T, and W within the control time period are determined using the six orbital elements of the corrected orbit. The method for obtaining the correction amount of the thruster working time length in each direction of the orbit to be controlled by numerical integration is as follows: ； ； ； ； ； In the formula, For the track semi-major, is the Earth's gravitational constant.

6. The method for precise maintenance control of a satellite's strict return orbit based on micro-electric propulsion according to claim 5, characterized in that: Before calculating the correction amount of the thruster working time length in each direction of the track to be controlled, an overshoot is introduced to limit the control accuracy of the six track members of the track to be corrected. The relationship between the overshoot and the control accuracy of the six track members of the track to be corrected is: the overshoot cannot exceed m times the control accuracy range, where m is a positive integer.

7. The method for precise maintenance control of a satellite's strict return orbit based on micro-electric propulsion according to claim 6, characterized in that: The method for controlling the accuracy of the six track elements in the overshoot correction track is as follows: Set the control precision to The rate of change is f, if ,but: ,like ,but: ; After obtaining the correction amount of the working time length of the thrusters in each direction of the track to be controlled, the working time of the thrusters in each direction is obtained according to the principle of minimizing the sum of the three-axis total velocity increment modulus. The principle for minimizing the sum of the total velocity increment moduli of the three axes is that the sum of the total velocity increment moduli of the three axes is... Minimum.

8. The method for precise maintenance control of a satellite's strict return orbit based on micro-electric propulsion according to claim 7, characterized in that: The control torque feedforward compensation during the thruster control cycle is used to eliminate disturbing torques during thruster operation via open-loop control. The method for thruster feedforward compensation is as follows: Obtain the angular momentum in that direction at the start of thruster operation; Determine the thruster's current jet duration in the direction of operation and the target angular momentum in that direction at the end of the thruster's operation. Calculate the thruster's feedforward compensation torque using the following method: In the formula: This represents the angular momentum in this direction at the start of thruster operation without feedforward compensation. Let be the angular momentum in that direction at the moment the thruster finishes its operation. This refers to the duration of jet propulsion in that direction; After obtaining the thruster feedforward compensation torque through reverse propulsion, a simulation test is performed to determine whether the preset angular momentum in that direction can be output at the end of the thruster's operation. Based on the determination result, the calculated thruster feedforward compensation torque is adjusted.

9. The precise control method for maintaining a satellite's strict return orbit based on micro-electric propulsion according to claim 8, characterized in that: After obtaining the feedforward compensation torque of each thruster, orbit control is executed within the thruster control cycle according to the calculated thruster operating time. During the orbit control process, the angular momentum of the actuator in the direction of the disturbance torque is monitored. When the accumulated angular momentum exceeds the maximum control angular momentum in the current direction, the control is initiated. Half of the angular momentum is controlled; suspend orbit control until the effect of angular momentum is eliminated, then resume the corresponding orbit control; maximum control angular momentum. The settings are based on prior historical data and the requirements of the micro-electric propulsion mission.

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