Fastest fuel-optimal on-board autonomous planning method for transferring anisoplanar target on elliptic orbit to coplanar orbit

Abstract

The application discloses a method for quickly changing orbit of an elliptical orbit out-of-plane target to the same orbit with the least fuel control on a satellite, and a method for independently solving the least fuel control on a satellite for quickly changing orbit from an out-of-plane orbit to the same orbit, which comprises the following steps: judging whether to perform normal orbit plane correction or orbit shape correction according to orbit elements of an initial satellite and a target; solving a least fuel normal orbit plane correction control strategy according to a normal orbit plane element difference between the satellite and the target; solving a least fuel orbit shape correction control strategy according to an in-orbit plane element difference between the satellite and the target; and iteratively correcting the least fuel control strategy by using a high-precision orbit extrapolation model. The method is suitable for orbit changing of any elliptical orbit satellite in near-earth space to any elliptical orbit target, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for solving a task planning on a satellite.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace, in particular to a most fuel-efficient control on-board autonomous planning method for fast orbit transfer of an out-of-plane target on an elliptical orbit to the same orbit. BACKGROUND

[0002] With the deepening of space applications, more and more spacecrafts use elliptical orbits. In the case of failure or failure of an elliptical orbit spacecraft (hereinafter referred to as the target) and the out-of-plane orbit of the maintenance spacecraft (hereinafter referred to as the satellite), the satellite needs to first quickly transfer to the same orbit as the failed / faulty target, and then implement high-precision rendezvous. Due to the huge fuel consumption of out-of-plane and large-scale orbit transfer, considering the power system constraints of the satellite, the problem of planning and controlling the most fuel-efficient on-board autonomous solution for fast transfer of an out-of-plane target on an elliptical orbit is studied. Through recent research in related fields, more research has been done on circular orbit transfer and rendezvous, and less research has been done on autonomous orbit transfer of an out-of-plane target on an elliptical orbit. SUMMARY

[0003] The purpose of the present application is to provide a most fuel-efficient control on-board autonomous planning method for fast orbit transfer of an out-of-plane target on an elliptical orbit to the same orbit, to reasonably set the multi-gradient control boundaries of each orbit element according to the orbit evolution characteristics, to propose an analytical solution method of one or joint control or separate control according to the coupling characteristics of each orbit element control, and to use the analytical solution as the initial value and use the high-precision orbit extrapolation model to quickly and accurately iterate the initial value to achieve the most fuel-efficient control under on-board autonomous solution.

[0004] In order to meet the demand for fast orbit transfer of an out-of-plane target on an elliptical orbit to the same orbit, the power system constraints and the most fuel-efficient control principle of the spacecraft are considered, and the orbit plane normal correction and the orbit shape correction are performed respectively. The analytical solution of the most fuel-efficient orbit plane normal correction and the most fuel-efficient orbit shape correction is used as the initial value, and a high-precision orbit extrapolation model is used to iterate the accurate control sequence. Here, the orbit plane normal correction refers to the correction of inclination and ascending node right ascension, and the orbit shape correction refers to the correction of semi-major axis and eccentricity vector (i.e. eccentricity, perigee amplitude).

[0005] In order to illustrate the above-mentioned application, the following technical solutions are adopted in this patent: a most fuel-efficient control on-board autonomous planning method for fast orbit transfer of an out-of-plane target on an elliptical orbit to the same orbit, the specific steps are as follows:

[0006] Step 1: According to the principle of the most fuel-efficient control, under the two-body model, according to the initial orbit elements of the satellite and the target, it is determined whether to perform orbit plane normal correction first or orbit shape correction first;

[0007] Step two, under the two-body model: for the orbit normal correction, according to the orbit normal orbital element difference of the satellite and the target, the analytical solution initial value of the most fuel-efficient orbit normal correction control strategy is solved;

[0008] Step three, under the two-body model: for the orbit shape correction, according to the orbit in-plane orbital element difference of the satellite and the target, the analytical solution initial value of the most fuel-efficient orbit shape correction control strategy is solved;

[0009] Step four, considering the influence of space perturbation, taking the analytical solution of the orbit normal correction in step two and the orbit shape correction in step three as the initial value, the Newton iteration method is used to solve the final value of the most fuel-efficient control strategy by using the high-precision orbit extrapolation model.

[0010] Further, the order judgment method of the orbit normal correction and the orbit shape correction in step one is as follows:

[0011] The coefficients of the fuel ratio representing the orbit normal correction and the orbit shape correction are calculated respectively 、 :

[0012]

[0013]

[0014] Wherein:

[0015] is the average orbit angular velocity, semi-major axis and eccentricity of the satellite;

[0016] is the average orbit angular velocity, semi-major axis and eccentricity of the target;

[0017] The subscript s represents the satellite, and the subscript t represents the target;

[0018] If , the orbit normal correction should be performed first; if , the orbit shape correction should be performed first.

[0019] Further, in step two, for the orbit normal correction, the current inclination of the satellite and the ascending node right ascension are corrected to the control target value and , and the inclination to-be-controlled quantity and the ascending node right ascension to-be-controlled quantity of the satellite are:

[0020]

[0021]

[0022] where j represents the jth iteration solution;

[0023] If and , the joint correction of inclination and right ascension of ascending node is performed; otherwise, the separate correction of inclination and right ascension of ascending node is performed.

[0024] Further, for the joint correction of inclination and right ascension of ascending node, the single-pulse control is adopted, and the most fuel-saving control quantity initial value and the control timing (the latitude amplitude) are:

[0025]

[0026] For the separate correction of inclination, the single-pulse control is adopted, and the most fuel-saving control quantity initial value and the control timing are:

[0027]

[0028] For the separate correction of right ascension of ascending node, the single-pulse control is adopted, and the most fuel-saving control quantity initial value and the control timing are:

[0029]

[0030] wherein:

[0031] is the average orbit inclination, the perigee amplitude, and the latitude amplitude of the target;

[0032] is the average orbit angular velocity, the semi-major axis, and the eccentricity of the target;

[0033] In the above three formulas, the closer to the satellite apogee is selected for correction, which is more fuel-saving.

[0034] Further, in step three, for the orbit shape correction, the current semi-major axis , eccentricity , and perigee amplitude of the satellite are corrected to the control target values , and , and then the semi-major axis, eccentricity, and perigee amplitude of the satellite are the controlled quantities:

[0035]

[0036]

[0037]

[0038] where j represents the jth iteration solution;

[0039] If and and , the semi-major axis-eccentricity-argument of perigee joint correction is carried out; otherwise, the semi-major axis-eccentricity joint correction and the argument of perigee single correction are carried out.

[0040] Further, for semi-major axis, eccentricity and argument of perigee joint correction, two-pulse control is adopted, and by judging the relationship between the semi-major axis to be controlled and the eccentricity vector to be controlled , the control strategy is determined;

[0041] If , the two-pulse control quantity initial value and the control time (latitude amplitude angle) are:

[0042]

[0043] If , the two-pulse control quantity initial value and the control time are:

[0044]

[0045] Wherein:

[0046]

[0047] For semi-major axis and eccentricity joint control, two-pulse control strategy is adopted, and the control quantity initial value and the control time (true anomaly) are:

[0048]

[0049] Wherein, represents the perigee altitude difference and the apogee altitude difference to be controlled in the jth iteration:

[0050]

[0051] Wherein: represents the expected value of the satellite perigee altitude and apogee altitude;

[0052] represents the current actual value of the satellite perigee altitude and apogee altitude;

[0053] For the perigee amplitude correction alone, the single pulse control is adopted, and the most fuel-efficient control quantity initial value And control timing (true perigee angle) Is:

[0054]

[0055] Wherein: The average orbit angular velocity of the target, the semi-major axis and the eccentricity.

[0056] Further, for the analytical solution of the orbit plane normal correction in step two and the orbit shape correction in step three, the fast iteration method using the high-precision orbit extrapolation model as the initial value is as follows:

[0057] (1) Convert the control quantity initial value solved by the analytical solution And the control timing Or Into finite thrust, substitute into the high-precision orbit extrapolation model (HPOP model), extrapolate to the control target formation time , solve the orbit root number to be controlled At the moment ;

[0058] (2) Determine whether to stop iteration:

[0059] Set the orbit root number control threshold , if , it means that all orbit roots of the satellite except the phase are controlled to the position, and the iteration is stopped;

[0060] (3) Otherwise, update the to-be-controlled quantity according to the following formula, and re-solve the analytical solution of the orbit plane normal correction in step two and the orbit shape correction in step three as the initial value and the high-precision orbit extrapolation model iteration in step (1).

[0061] .

[0062] The beneficial effects of the present application are: to provide a fast orbit transfer on- satellite autonomous planning method for anisoplanar targets on an elliptical orbit to the same orbit with the most fuel-efficient control, to set a multi-gradient control boundary for each orbit root number according to the orbit evolution characteristics; to propose an analytical solution method of one or joint control or separate control according to the coupling characteristics of each orbit root number control; to use the analytical solution as the initial value, and to use the high-precision orbit extrapolation model to perform fast and high-precision iteration on the initial value, and to realize the most fuel-efficient control under on-satellite autonomous solution.

[0063] The present application is based on the analytical solution of the relative orbit root number variable orbit problem derived from the Gauss perturbation equation, and proposes a fast variable orbit strategy for out-of-plane targets on elliptical orbits based on a multi-gradient control boundary, and achieves high-precision control through high-precision orbit extrapolation model iteration.

[0064] The present application comprises: judging whether to first perform normal orbit plane correction or orbit shape correction according to the initial orbit roots of the satellite and the target; solving the most fuel-efficient normal orbit plane correction control strategy according to the normal orbit root difference of the orbit plane of the satellite and the target; solving the most fuel-efficient orbit shape correction control strategy according to the in-plane orbit root difference of the satellite and the target; and iteratively correcting the most fuel-efficient control strategy by using a high-precision orbit extrapolation model. The method is applicable to variable orbit of any elliptical orbit satellite to any elliptical orbit target in near-earth space, and the control accuracy can reach ten meters; and the calculation amount is small, and is suitable for on-board autonomous mission planning solution. DETAILED DESCRIPTION

[0065] Firstly, taking the most fuel-efficient control as the principle, the initial instantaneous orbit roots of the satellite and the target are taken as the input under the two-body model to judge whether to first perform normal orbit plane correction or first perform orbit shape correction.

[0066] Secondly, under the two-body model: for normal orbit plane correction, the analytical solution of the most fuel-efficient normal orbit plane correction control strategy is solved according to the normal orbit root difference of the orbit plane of the satellite and the target; for orbit shape correction, the analytical solution of the most fuel-efficient orbit shape correction control strategy is solved according to the in-plane orbit root difference of the satellite and the target.

[0067] Finally, considering the influence of space perturbation, the Newton iteration method is used to accurately iterate the finite thrust accurate control sequence by using the high-precision orbit extrapolation model (i.e. HPOP model) with the analytical solution as the initial value.

[0068] Further, the specific implementation steps of the technical solution are as follows:

[0069] Step 1: judging whether to first perform normal orbit plane correction or first perform orbit shape correction

[0070] Since the normal orbit plane correction and the orbit shape correction of the spacecraft are independent of each other, the principle of saving fuel is used to judge whether the satellite first performs normal orbit plane correction to save more fuel or first performs orbit shape correction to save more fuel. The following coefficients are calculated to represent the fuel ratio of normal orbit plane correction and orbit shape correction 、 :

[0071] (1)

[0073] (2)

[0075] where:

[0076] are the mean orbital angular velocity, semi-major axis and eccentricity of the satellite;

[0077] are the mean orbital angular velocity, semi-major axis and eccentricity of the target;

[0078] The subscript s denotes the satellite, and the subscript t denotes the target.

[0079] If , it indicates that the orbit normal correction is more fuel-efficient, and the orbit normal correction is performed first.

[0080] If , it indicates that the orbit shape correction is more fuel-efficient, and the orbit shape correction is performed first.

[0081] Step 2: Solve the most fuel-efficient control analytical solution of the orbit normal correction and the orbit shape correction

[0082] Under the two-body problem, there is an analytical solution of the most fuel-efficient control for both the orbit normal correction and the orbit shape correction. Solve the analytical solution as the initial value.

[0083] 1. Orbit normal correction analytical solution

[0084] (1) Solve the inclination and the ascending node right ascension to be controlled

[0085] For the orbit normal correction, the current inclination of the satellite is corrected and the ascending node right ascension to the control target value and , then the inclination to be controlled of the satellite and the ascending node right ascension to be controlled are:

[0086] (3)

[0088] (4)

[0090] For the orbit transfer to the same orbit as the target problem, the target values and are the inclination and the ascending node right ascension of the target, i.e. , ;

[0091] For the orbit transfer to perform some configuration of the target, such as companion flying, etc. and Solve according to configuration requirements;

[0092] In the above formula, j represents the jth iteration solution.

[0093] (2) Determine whether to use joint correction or separate correction

[0094] Under the two-body problem, according to the Gauss perturbation equation, the most fuel-efficient correction opportunity of the inclination is at the ascending-descending node (i.e. at the latitude amplitude angle of 0° and 180°); the most fuel-efficient correction opportunity of the ascending node right ascension is at the south-north point (i.e. at the latitude amplitude angle of 90° and 270°); the joint correction of the inclination and the ascending node right ascension is more fuel-efficient than the separate correction of the two, and the most fuel-efficient correction opportunity of the joint correction has uniqueness, which is the orbit intersection of the satellite and the target.

[0095] However, under the space perturbation, when the orbital plane difference between the satellite and the target is small, i.e. the inclination and the ascending node right ascension difference between the satellite and the target is small to a certain extent, the orbit intersection of the satellite and the target becomes unstable under the space perturbation, and the joint correction is prone to control opportunity oscillation problem, so it is appropriate to correct separately.

[0096] Therefore, the boundary parameters for determining whether to use joint correction or separate correction for the inclination and the ascending node right ascension are: the inclination joint correction boundary parameter .

[0097] If and , it indicates that the controlled quantities of the inclination and the ascending node right ascension are both large, and the joint correction of the inclination and the ascending node right ascension is performed;

[0098] Otherwise, it indicates that at least one of the controlled quantities of the inclination and the ascending node right ascension is small, and the joint correction may fall into the control opportunity oscillation problem, so the separate correction of the inclination and the ascending node right ascension is performed.

[0099] Generally, for near-earth orbit spacecraft, it is appropriate to:

[0100] (5)

[0102] (6)

[0104] (3) Analytical solutions of joint correction and separate correction

[0105] For the joint correction of the inclination and the ascending node right ascension, the most fuel-efficient control quantity initial value and the control opportunity (latitude amplitude angle) are:

[0106] (7)

[0108] Wherein: is the average orbit inclination, argument of perigee, argument of latitude of the target;

[0109] For inclination correction alone, the single pulse control is adopted, and the most fuel-saving control amount initial value and control timing (the argument of latitude) are:

[0110] (8)

[0112] For the ascending node right ascension correction alone, the single pulse control is adopted, and the most fuel-saving control amount initial value and control timing (the argument of latitude) are:

[0113] (9)

[0115] In the above three formulas, the latitude argument closer to the satellite apogee is selected for correction, which is more fuel-saving.

[0116] If the single pulse control amount exceeds the maximum control amount of the satellite single control , the multi-pulse correction is selected, and the multi-pulse correction number K is:

[0117] (10)

[0119] Wherein: .

[0120] For the multi-pulse control condition, after each control is completed, the actual control result is repeated to solve as above.

[0121] 2, Analytical solution of orbit shape correction

[0122] (1) Solve the semi-major axis and eccentricity vector control amount

[0123] For orbit shape correction, the current semi-major axis , eccentricity , argument of perigee of the satellite are corrected to the control target value , and , then the semi-major axis, eccentricity and argument of perigee control amount of the satellite are:

[0124] (11)

[0126] (12)

[0128] (13)

[0130] For small elliptical orbits, because of the small eccentricity, the eccentricity and the argument of perigee can also be expressed by the eccentricity vector in the LVLH coordinate system :

[0131] (14)

[0133] Where:

[0134] is the eccentricity vector The components in the LVLH coordinate system

[0135] are the eccentricity and the argument of perigee of the spacecraft.

[0136] The to-be-controlled quantity of the eccentricity vector at this time can be expressed as :

[0137] (15)

[0139] Where:

[0140] is the mean orbital eccentricity and the argument of perigee of the satellite;

[0141] For the problem of orbit transfer to the same orbit as the target, the semi-major axis, eccentricity, and argument of perigee target values , , of the satellite are the semi-major axis of the target, the eccentricity , , ;

[0142] For the problem of orbit transfer to perform some formation flying around the target, etc. , and are solved and set according to the formation requirements;

[0143] In the above formula, j represents the jth iteration solution.

[0144] (2) Determine whether to use joint correction or separate correction

[0145] Under the two-body problem, according to the Gauss perturbation equation, the most fuel-efficient correction time of the semi-major axis is at the perigee (i.e. at the true anomaly of 0°); the most fuel-efficient correction time of the eccentricity is at the perigee and apogee (i.e. at the true anomaly of 0° and 180°); the most fuel-efficient correction time of the argument of perigee is at the line of apsides (i.e. at the true anomaly of 90° and 270°); joint correction of the semi-major axis, eccentricity and argument of perigee is more fuel-efficient than separate correction of the three, and the most fuel-efficient correction time of joint correction has infinitely many solutions, which can be flexibly adjusted according to engineering constraints.

[0146] However, under spatial perturbation, similar to the normal correction of the orbital plane, when the satellite and the target have small differences in orbital shape, i.e. the semi-major axis and eccentricity vector of the satellite and the target have small differences to a certain extent, the joint correction control time becomes unstable under spatial perturbation, and the joint correction is prone to control time oscillation problem, and the fuel saving advantage of joint correction is no longer obvious. At this time, separate correction is appropriate.

[0147] Therefore, the boundary parameters for determining whether to use joint correction or separate correction for the semi-major axis, eccentricity and argument of perigee are as follows: the joint correction boundary parameter of the semi-major axis is , the joint correction boundary parameter of the eccentricity is , and the joint correction boundary parameter of the argument of perigee is .

[0148] If and and , it indicates that the to-be-controlled quantities of the semi-major axis, eccentricity and argument of perigee are all large, and joint correction of the semi-major axis, eccentricity and argument of perigee is performed.

[0149] Otherwise, it indicates that at least one of the to-be-controlled quantities of the semi-major axis, eccentricity and argument of perigee is small, and joint correction may be prone to control time oscillation problem, and separate correction of the semi-major axis, eccentricity and argument of perigee is performed.

[0150] In general, for near-earth orbit spacecraft, the following can be taken:

[0151] (16)

[0153] (17)

[0155] (18)

[0157] (3) Analytical solutions of joint correction and separate correction

[0158] For the joint correction of semi-major axis, eccentricity and argument of perigee, two-impulse control is adopted. The control strategy is determined by judging the relationship between the semi-major axis to be controlled and the eccentricity vector to be controlled.

[0159] If , the initial value of the two-impulse control quantity with the least fuel consumption and the control timing (latitude amplitude angle) are:

[0160] (19)

[0162] If , the initial value of the two-impulse control quantity with the least fuel consumption and the control timing (latitude amplitude angle) are:

[0163] (20)

[0165] For the joint control of semi-major axis and eccentricity, two-impulse control strategy is adopted. The initial value of the control quantity and the control timing (true anomaly) are:

[0166] (21)

[0168] where represents the perigee height difference and the apogee height difference to be controlled in the jth iteration:

[0169] (22)

[0171] where:

[0172] represents the expected value of the satellite's perigee height and apogee height;

[0173] represents the current actual value of the satellite's perigee height and apogee height;

[0174] For the single correction of the argument of perigee, single-impulse control is adopted. The initial value of the control quantity with the least fuel consumption and the control timing (true anomaly) are:

[0175] (23) ​​​​​​​​

[0177] If the single-pulse control quantity exceeds the maximum control quantity of a single satellite control... Then, multi-pulse correction is selected, and the number of multi-pulse corrections K is:

[0178] (twenty four)

[0180] in: .

[0181] For multi-pulse control, the above solution is repeated after each control operation, based on the actual control results.

[0182] Step 3: Iterate the exact solution of finite thrust control using a high-precision orbit extrapolation model.

[0183] Using the analytical solution of the most fuel-efficient control for orbital plane normal correction and orbital shape correction under the two-body problem as the initial value, the Newton iteration method is adopted, and the exact solution of finite thrust control is iterated using the high-precision orbital extrapolation model (HPOP model).

[0184] 1. Determine the control effect of the current control sequence.

[0185] Initial values ​​of the most fuel-efficient pulse thrust control sequence solved under the two-body model. Converted into finite thrust Control duration The duration of thrust application is evenly distributed before and after the control timing. This utilizes finite thrust. Controlling duration and control the timing (or Substitute the data into the high-precision orbit extrapolation model (HPOP model) and extrapolate to the time of control target formation. The target orbit at that moment is obtained. and satellite orbit .

[0186] Solve The satellite's relative orbital elements to be controlled at any given time :

[0187] (25)

[0189] 2. Determine if control is in place and stop iteration.

[0190] Set the threshold for controlling the number of track elements. Determine whether the current control sequence is in place:

[0191] like At the same time, it is established that all orbital elements of the satellite are controlled except for the phase, that is, the satellite orbit plane and orbit formation are controlled, and then the iteration is stopped;

[0192] Otherwise, go to the following steps and continue iteration.

[0193] 3. Reset the control target to solve the analytical solution

[0194] Update the satellite orbital element control target in step 2 according to the following formula, and re-solve the analytical solution of step 2 and the HPOP extrapolation of steps 1 and 2 in step 3.

[0195]

[0196] (26).

[0197] In order to make the physical law and effect of the patent easy to understand, the following will be further described in combination with specific embodiments.

[0198] Task requirements and inputs

[0199] The patent is aimed at the rapid orbit transfer of spatial elliptical orbit out-of-plane targets to the same orbit, and develops control sequence planning solution suitable for on-board autonomous rapid solution and high-precision control. The task requirements and input conditions are as follows.

[0200] 1. Initial orbit of satellite and target

[0201] The initial orbit elements of the satellite and the target are shown in Table 1.

[0202] Table 1

[0203]

[0204] 2. Satellite control system constraints

[0205] The satellite control system constraints are shown in Table 2.

[0206] Table 2

[0207]

[0208] 3. Task objectives and control thresholds

[0209] The task objective of the satellite is to control the satellite and the target to have the same orbital elements except for the phase.

[0210] The threshold settings for the control of each orbital element of the satellite are shown in Table 3.

[0211] Table 3

[0212]

[0213] Task solving result

[0214] The whole process control sequence of the satellite relative to the target from the different planes to the same orbit is shown in Table 4. It can be seen that the control effect of each solving reaches the threshold value of the orbit root control to the position required in Table 3.

[0215] Table 4

[0216]

[0217] The relative orbit plane roots of the satellite before and after control and the target are shown in Table 5. It can be seen from Table 5 that the satellite and the target achieve the transfer from the different planes to the same orbit.

[0218] Table 5

[0219]

[0220] The above only describes the preferred embodiments of the present application. It should be noted that for ordinary skilled in the art, without departing from the principles of the present application, a number of improvements and refinements can be made, and these improvements and refinements should be considered as the protection scope of the present application.

Claims

1. A method for fast orbit transfer of an out-of-plane target to the same orbit with minimum fuel consumption on the control satellite, characterized in that, The specific steps are as follows: Step one, according to the principle of minimum fuel control, under the two-body model, according to the initial orbital elements of the satellite and the target, it is determined whether to carry out normal orbit correction first or to carry out orbit shape correction first; Step two, under the two-body model: for normal orbit correction, according to the difference of the normal orbit elements of the satellite and the target, the analytical solution initial value of the minimum fuel normal orbit correction control strategy is solved; Step three, under the two-body model: for orbit shape correction, according to the difference of the in-plane orbit elements of the satellite and the target, the analytical solution initial value of the minimum fuel orbit shape correction control strategy is solved; Step four, considering the influence of space perturbation, taking the analytical solution of the normal orbit correction in step two and the orbit shape correction in step three as the initial value, using the Newton iteration method, the high-precision orbit extrapolation model is used to iteratively solve the final value of the minimum fuel control strategy.

2. The method of claim 1, wherein the method is characterized in that, The sequence judgment method of normal orbit correction and orbit shape correction in step one is as follows: coefficients of the fuel ratio representing the normal correction of the orbital plane and the shape correction of the orbit are calculated respectively , : Wherein: the mean orbital angular velocity, semi-major axis and eccentricity of the satellite; Target mean orbital angular velocity, semi-major axis, and eccentricity; The subscript s represents the satellite, and the subscript t represents the target; If , the normal correction of the orbital plane should be performed first; if , the shape correction of the orbit should be performed first.

3. The method of claim 1, wherein the method is characterized in that: In step two, for the normal correction of the orbit plane, the current inclination of the satellite is corrected and the right ascension of the ascending node to the control target value and then the inclination of the satellite to be controlled and the right ascension of the ascending node to be controlled are: Wherein, j represents the jth iteration solution; If and , a combined correction of the inclination and the right ascension of the ascending node is made; otherwise a separate correction of the inclination and the right ascension of the ascending node is made.

4. The method of claim 3, wherein the method is characterized in that: For the combined correction of inclination and longitude of ascending node, the initial value of the control quantity for the minimum fuel control with single pulse is and control timing : For the pitch correction alone, the initial value of the control quantity for the minimum fuel control with single pulse control is and the control timing is : For the correction of the right ascension of the ascending node alone, the initial value of the control quantity for the minimum fuel control with single pulse is used and the control timing is: Wherein: Target mean orbital inclination, argument of perigee, argument of latitude; Target mean orbital angular velocity, semi-major axis, and eccentricity; Among the above three formulae, the one with the closest distance to the satellite's apogee is selected is modified, and fuel is saved more.

5. The method of claim 1, wherein the method is characterized in that: In step three, for the orbit shape correction, the current semi-major axis of the satellite is corrected , eccentricity , argument of perigee to the control target value , and , the semi-major axis, eccentricity and argument of perigee of the satellite are controlled quantities: Wherein, j represents the jth iteration solution; If and and , the semi-major axis-eccentricity-amplitude of the argument of the perigee joint correction is performed; otherwise, the semi-major axis-eccentricity joint correction and the argument of the perigee individual correction are performed.

6. The method of claim 5, wherein the method is characterized in that: For the joint correction of semi-major axis, eccentricity and argument of perigee, two impulse control is adopted The control strategy is determined by judging the relationship between the semi-major axis to be controlled and the eccentricity vector to be controlled If , the two-pulse control amount initial value that is the most fuel-efficient , and the control timing are: If , the two-pulse control amount initial value that is the most fuel-efficient , and the control timing are: Wherein: For the joint control of semi-major axis and eccentricity, two-pulse control strategy is adopted, and the initial value of control quantity and control timing is wherein, represents the perigee height difference and the apogee height difference of the jth iteration wherein: denotes the expected values of the satellite's perigee and apogee heights; represent the current actual values of the satellite's perigee and apogee heights; For the perigee amplitude angle correction alone, the single pulse control is adopted, and the initial value of the control quantity for the most fuel-saving control is and the control timing : where: The mean orbital angular velocity, semi-major axis, and eccentricity of the target.

7. The method of claim 1 or 4 or 6, characterized in that: For the high-precision orbit extrapolation model, the analytical solution of the normal orbit correction in step two and the orbit shape correction in step three is taken as the initial value, and the fast iteration method is as follows: (1) The control variable initial value solved by analytical solution and control timing or are converted into finite thrust, substituted into the high-precision orbit extrapolation model, extrapolated to the control target formation time , and the orbit root number of the satellite relative to the target at time is solved ; (2) Determine whether to stop iteration: Set the number of orbit control to the threshold of the position , if , it means that all the orbit parameters of the satellite are controlled to the position except the phase, and the iteration is stopped; (3) Otherwise, update the control quantity according to the following formula, and re-solve the analytical solution of the normal orbit correction in step two and the orbit shape correction in step three as the initial value and the high-precision orbit extrapolation model iteration in step (1); 。

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