A method for simultaneously maintaining the eccentricity vector and orbit angle of geosynchronous orbit

Abstract

The present invention discloses a method for simultaneously maintaining the eccentricity vector and orbital argument of a geosynchronous orbit, comprising the following steps: S1, determining the calculation timing of a position-keeping strategy; S2, calculating the position-keeping strategy; S3, executing a corresponding switch instruction of an electric propulsion engine at an orbital argument that meets corresponding conditions according to the calculated orbital strategy.

Description

Technical Field

[0001] The invention relates to a method for maintaining the east-west position of a geosynchronous satellite, and in particular to a method for simultaneously maintaining an eccentricity vector and an orbit angle of a geosynchronous orbit. Background Art

[0002] An ideal geostationary orbit is one where the satellite is stationary above the Earth when viewed from any point on the Earth. Three conditions must be met: 1. The satellite orbit period is consistent with the Earth's rotation period; 2. The orbit eccentricity is 0, so that the orbit is a perfect circle; 3. The orbit inclination is 0, so that the satellite orbit coincides with the Earth's instantaneous equatorial plane. A satellite that meets these three conditions has a sub-satellite point at a fixed point on the equator, also known as the fixed longitude of a geostationary orbit satellite. For multi-satellite co-located satellites, in addition to isolation through longitude, co-location can also be achieved through eccentricity vectors, orbit inclination vectors, or a combination of eccentricity vectors + orbit inclination vectors. The sub-satellite point trajectory of such satellites is no longer fixed to a point on the Earth's equator, but moves within a certain range. Such satellites are called geosynchronous satellites.

[0003] Satellites in orbit are subject to various environmental perturbations, such as the non-spherical shape of the Earth, the gravitational force of the sun and the moon, and the solar pressure. Under the influence of these environmental perturbations, the sub-satellite point of the satellite will gradually deviate from the nominal fixed longitude. Therefore, it is necessary to use the medium exchange method regularly based on the principle of momentum exchange to generate reverse thrust on the satellite to maintain the fixed longitude of the satellite, also known as "position keeping".

[0004] The "position keeping" of geosynchronous satellites is divided into "east-west keeping" and "north-south keeping". Among them, "east-west keeping" includes orbital altitude and orbital eccentricity keeping, which belongs to orbital plane control; "north-south keeping" and orbital inclination keeping belong to orbital plane control. As we all know, orbital plane control and out-of-plane control are decoupled from each other and do not affect each other. Synchronous keeping technology will focus on orbital plane control, including orbital angle keeping and eccentricity vector keeping, that is, the "east-west keeping" of geosynchronous satellites.

[0005] The speed increment required to maintain the orbital angle in the east-west hold does not exceed 2m / s per year. The speed increment required to maintain the eccentricity does not exceed 8m / s per year according to the different satellite surface mass ratio. Summary of the invention

[0006] The purpose of the present invention is to provide a method for simultaneously maintaining the eccentricity vector and orbit angle of a geosynchronous orbit, which can realize the capture and maintenance of parameters within the orbital plane and belongs to the minimum fuel consumption maintenance.

[0007] In order to achieve the above objectives, the present invention is implemented by the following technical solutions:

[0008] A method for simultaneously maintaining a geosynchronous orbit eccentricity vector and an orbital argument, characterized in that it comprises the following steps:

[0009] S1, determine the calculation timing of the position protection strategy;

[0010] S2, calculate the position protection strategy;

[0011] S3, according to the calculated orbital strategy, execute the corresponding switching instructions of the electric propulsion engine at the orbital angle that meets the corresponding conditions.

[0012] The step S1 includes: if the current satellite local time theta_s1 is greater than the agreed local time theta_s0, and the orbit keeping execution end flag: SK_step_t==0, SK_step_r==0, it is considered that the time for calculating the position keeping strategy has arrived, otherwise the time for calculating the position keeping strategy has not arrived.

[0013] The current satellite local time theta_s1 is greater than the agreed local time theta_s0 as follows:

[0014] Mod(theta_s1-theta_s0+12,24)-12>0

[0015] theta_s1=12 / pi×[Lon+alphaG-atan2(ys,xs)]-12

[0016] Among them: Mod is the remainder function, ys is the projection of the current sun position on the y-axis of the J2000 coordinate system; xs is the projection of the current sun position on the x-axis of the J2000 coordinate system; Lon is the geographic longitude of the current sub-satellite point of the satellite; alphaG is the Greenwich sidereal hour angle at the current moment; atan2() is the inverse tangent function.

[0017] The step S2 comprises:

[0018] S2.1, calculate the tangential velocity increment delta_V required to maintain the eccentricity;

[0019] S2.2, determine the orbital argument of the position and orbit control, if delta_V<0.000000001, select the orbital argument ufire_m_EW corresponding to the ignition point, ufire_m_EW=π, the corresponding angle is 180°, V_r=0, V_r is the radial ignition speed increment; if delta_V>=0.000000001, execute step S2.2.1;

[0020] S2.2.1, calculate the temporary ignition point: u_std = atan2(eyMEAN, exMEAN);

[0021] eyMEAN and exMEAN represent the y-component and x-component of the satellite orbit eccentricity vector, respectively;

[0022] S2.2.2, calculate the error between the current geographic accuracy and the target geographic accuracy:

[0023] delta_lon=lonMEAN-lon_Target, where lon_Target is the longitude of the target point;

[0024] S2.2.3, Calculate orbit angle control accuracy:

[0025] u_yuzhi=-0.125×ddu×864002, where ddu is the drift acceleration of the satellite under natural perturbation;

[0026] S2.2.4, calculate the relative drift velocity dn0 of the satellite at the current moment:

[0027] dn0 = -1.5 × we × (smaMEAN - sma0) / sma0, where smaMEAN is the semi-major axis of the satellite's horizontal orbit at the current moment;

[0028] S2.2.5, calculate the velocity increment V_t along the trajectory:

[0029] If delta_lon+u_yuzhi<0

[0030] V_t=-1 / 3×sma0×(sqrt(2×ddu×(delta_lon+u_yuzhi))-dn0);

[0031] If delta_lon+u_yuzhi>=0

[0032] V_t=0;

[0033] S2.2.6, calculate the angle theta of the ignition point:

[0034] If delta_V>|V_t|

[0035] theta=pi / 2-asin(|V_t| / delta_V)

[0036] If delta_V<=|V_t|

[0037] theta = 0;

[0038] S2.2.7, calculate the orbital argument ufire_m_EW corresponding to the ignition point:

[0039] ufire_m_EW=u_std+sign_vr×theta

[0040] Wherein, sign_vr=1 means that the satellite-z-plane thruster is used, and the radial velocity increment is negative; sign_vr=-1 means that the satellite z-plane thruster is used, and the radial velocity increment is positive;

[0041] S2.2.8, calculate the radial velocity increment V_r:

[0042] V_r=sign_vr×delta_V×cos(pi / 2-theta)×2

[0043] S2.2.9, calculate the orbital arguments corresponding to the start and end times of ignition:

[0044] Calculate the thrust ignition duration: t_thrust_t = mass × |V_t| / F;

[0045] Calculate radial ignition duration: t_thrust_r = mass × |V_r| / F;

[0046] Where: mass is the mass of the satellite; F is the thrust;

[0047] Calculate the orbital argument corresponding to the start time of the ignition:

[0048] ufire_t_0=mod(ufire_m_EW-0.5×we×t_thrust_t,2×pi);

[0049] ufire_t_t=mod(ufire_m_EW+0.5×we×t_thrust_t,2×pi);

[0050] Among them, ufire_t_0 is the orbit angle corresponding to the start time of the track ignition;

[0051] ufire_t_t is the orbit width corresponding to the end time of the ignition in the track direction;

[0052] Calculate the orbital angle corresponding to the radial ignition start time:

[0053] ufire_r_0=mod(ufire_m_EW-0.5×we×t_thrust_r,2×pi);

[0054] ufire_r_t=mod(ufire_m_EW+0.5×we×t_thrust_r,2×pi);

[0055] ufire_r_0: The radial ignition start time corresponds to the orbital argument;

[0056] ufire_r_t: The orbital argument corresponding to the radial ignition termination time;

[0057] S2.2.10, Update track keeping strategy generation flag:

[0058] If |V_t|>yuzhi_2, where yuzhi_2>yuzhi_1, then SK_step_t=1, indicating that track control is required; otherwise: SK_step_t=0, indicating that track control is not required this time;

[0059] If |V_r|>yuzhi_2, where yuzhi_2>yuzhi_1, then SK_step_r=1, indicating that radial orbit control is required; otherwise: SK_step_r=0, indicating that radial orbit control is not required this time.

[0060] The step S3 comprises:

[0061] S3.1, perform directional and radial orbit control:

[0062] If mod(Lon+alphaG-ufire_t_0+pi,2×pi)-pi>0

[0063] And mod(Lon+alphaG-ufire_t_0+pi,2×pi)-pi<=we×Crl_period

[0064] And t_thrust_t≠0, and SK_step_t==1

[0065] Among them: Crl_period is the control period of the satellite control system;

[0066] Then the satellite's east side orbit control engine is started to ignite to slow down the satellite. The engine ignition duration is t_thrust_t;

[0067] If mod(Lon+alphaG-ufire_r_0+pi,2×pi)-pi>0

[0068] And mod(Lon+alphaG-ufire_r_0+pi,2×pi)-pi<=we×Crl_period

[0069] And t_thrust_r≠0, and SK_step_r==1

[0070] Then the satellite's east side orbit control engine is started and ignited. The engine ignition duration is t_thrust_r.

[0071] S3.2. Update the track keeping strategy generation flag:

[0072] If SK_step_t==1, and mod(Lon+alphaG-ufire_t_t+pi,2×pi)-pi>we×Crl_period, then:

[0073] SK_step_t = 0;

[0074] If SK_step_r == 1, and mod(Lon+alphaG-ufire_r_t+pi,2×pi)-pi>we×Crl_period, then:

[0075] SK_step_r=0.

[0076] Compared with the prior art, the present invention has the following advantages:

[0077] First, when the speed increment required for orbital argument maintenance is greater than the speed increment required for eccentricity maintenance (or the east-west position keeping accuracy requirement is low and the eccentricity vector does not need to be kept with high accuracy), only the direction position keeping can be used to simultaneously achieve orbital argument and eccentricity vector maintenance, which saves the most fuel.

[0078] Second, when the velocity increment required for orbital argument maintenance is less than the velocity increment required for eccentricity maintenance (it is necessary to maintain the eccentricity vector with high precision, such as multi-satellite co-location maintenance), part of the eccentricity vector is maintained by the track velocity increment, and the remaining part of the eccentricity vector is maintained by the radial velocity increment, which saves the most fuel;

[0079] Third, the directional orbit control for orbit angle maintenance and the radial orbit control for eccentricity vector maintenance are carried out simultaneously to reduce the number of position maintenance, minimize the number of position maintenance, and reduce the impact of attitude disturbance during position maintenance on the load work. If combined with the oblique layout design of the thruster, it can further reduce fuel consumption;

[0080] Fourthly, the present invention is applicable to the capture and maintenance of a fixed position of any point in the area with a longitude of 80° to 155° (China and its surrounding areas, orbit) without consuming additional fuel, and is the most fuel-efficient. BRIEF DESCRIPTION OF THE DRAWINGS

[0081] Figure 1 A flow chart of a method for simultaneously maintaining a geosynchronous orbit eccentricity vector and an orbital argument according to the present invention;

[0082] Figure 2 It is a schematic diagram of the present invention using a satellite-z-plane thruster to generate radial velocity increments;

[0083] Figure 3 It is a schematic diagram of the present invention using a satellite z-plane thruster to generate radial velocity increments. DETAILED DESCRIPTION

[0084] The present invention will be further described below by describing a preferred specific embodiment in detail in conjunction with the accompanying drawings.

[0085] like Figures 1 to 3 As shown, a method for simultaneously maintaining a geosynchronous orbit eccentricity vector and an orbital angle comprises the following steps:

[0086] S1, the onboard computer determines the calculation timing of the position-keeping strategy;

[0087] S2, calculate the position protection strategy;

[0088] S3, according to the calculated orbital strategy, execute the corresponding switching instructions of the electric propulsion engine at the orbital angle that meets the corresponding conditions.

[0089] The step S1 includes: if the current satellite local time theta_s1 is greater than the agreed local time theta_s0, and the orbit keeping execution end flag: SK_step_t==0, SK_step_r==0, it is considered that the time for calculating the position keeping strategy has arrived, otherwise the time for calculating the position keeping strategy has not arrived.

[0090] It should be noted that some geosynchronous satellites are not equipped with GNSS (Global Navigation Satellite System) and need to add orbital parameters on the ground. The timing of adding orbital parameters on the ground must be agreed upon: it needs to be before theta_s0 and the difference with theta_s0 should be less than 3 hours.

[0091] The preceding satellite local time theta_s1 is greater than the agreed local time theta_s0 as follows:

[0092] Mod(theta_s1-theta_s0+12,24)-12>0

[0093] theta_s1=12 / pi×[Lon+alphaG-atan2(ys,xs)]-12

[0094] Among them: ys is the y-axis projection of the current sun position in the J2000 coordinate system; xs is the x-axis projection of the current sun position in the J2000 coordinate system; Lon is the geographic longitude of the current sub-satellite point of the satellite; alphaG is the Greenwich sidereal hour angle at the current moment; atan2() is the inverse tangent function;

[0095] atan2(y,x) ranges from -pi to pi, specifically:

[0096] When x>0, atan2(y,x)=atan(y / x), the range is (-pi / 2,pi / 2);

[0097] When x < 0, atan2(y, x) = pi + atan(y / x), and the range is (pi / 2, 1.5 * pi);

[0098] When x = 0 and y > 0, atan2(y, x) = pi / 2;

[0099] When x = 0 and y < 0, atan2(y, x) = -pi / 2;

[0100] Finally, through the modulo function, the value range of atan2(y, x) is converted to [0, 2 * pi).

[0101] In a specific embodiment, the step S2 includes:

[0102] S2.1, calculate the magnitude of the tangential velocity increment delta_V required to maintain the eccentricity:

[0103] delta_V = sqrt(exMEAN2 + eyMEAN2) × sma0 × we / 2

[0104] Where:

[0105] exMEAN is the x-axis component of the eccentricity vector, with a dimension of 1;

[0106] eyMEAN is the y-axis component of the eccentricity vector, with a dimension of 1;

[0107] sma0 = 42166289 is the semi-major axis of the satellite's nominal orbit, in units of m;

[0108] we = 7.292115847907396 × 10^-5 is the angular velocity of the Earth's rotation, in units of rad / s;

[0109] S2.2, determine the argument of latitude of the orbit where the orbit maintenance control is located. If delta_V < yuzhi_1, and yuzhi_1 = 0.000000001, it is considered that the current eccentricity is close to 0, and there is no need to select the optimal ignition point to perform orbit maintenance. Select the argument of latitude corresponding to the ignition point, select the argument of latitude ufire_m_EW corresponding to the ignition point, ufire_m_EW = pi, and the corresponding angle is 180°, V_r = 0, where V_r is the radial ignition velocity increment. If delta_V >= yuzhi_1, then execute step S2.2.1;

[0110] S2.2.1, calculate the temporary ignition point: u_std = atan2(eyMEAN, exMEAN);

[0111] S2.2.2, calculate the error between the current geographical accuracy and the target geographical accuracy:

[0112] delta_lon=lonMEAN-lon_Target, where lon_Target is the longitude of the target point;

[0113] S2.2.3, Calculate orbit angle control accuracy:

[0114] u_yuzhi=-0.125×ddu×86400 2 , where ddu is the drift acceleration of the satellite under natural perturbation. For the area with longitude from 80° to 155°, ddu<0, indicating accelerated drift toward the west;

[0115] S2.2.4, calculate the relative drift velocity dn0 of the satellite at the current moment:

[0116] dn0 = -1.5 × we × (smaMEAN - sma0) / sma0, where smaMEAN is the semi-major axis of the satellite's horizontal orbit at the current moment;

[0117] S2.2.5, calculate the velocity increment V_t along the trajectory:

[0118] If delta_lon+u_yuzhi<0

[0119] V_t=-1 / 3×sma0×(sqrt(2×ddu×(delta_lon+u_yuzhi))-dn0);

[0120] If delta_lon+u_yuzhi>=0

[0121] V_t=0;

[0122] S2.2.6, calculate the angle theta of the ignition point:

[0123] If delta_V>|V_t|

[0124] theta=pi / 2-asin(|V_t| / delta_V)

[0125] If delta_V<=|V_t|

[0126] theta = 0;

[0127] S2.2.7, calculate the orbital argument ufire_m_EW corresponding to the ignition point:

[0128] ufire_m_EW=u_std+sign_vr×theta

[0129] Wherein, sign_vr = 1 means that the radial velocity increment is negative when the satellite-z thruster is used (see Figure 2); sign_vr = -1 means that the radial velocity increment is positive when the satellite z-plane thruster is used (see Figure 3 );

[0130] S2.2.8, calculate the radial velocity increment V_r:

[0131] V_r=sign_vr×delta_V×cos(pi / 2-theta)×2

[0132] S2.2.9, calculate the orbital arguments corresponding to the start and end times of ignition:

[0133] Calculate the thrust ignition duration: t_thrust_t = mass × |V_t| / F;

[0134] Calculate radial ignition duration: t_thrust_r = mass × |V_r| / F;

[0135] Where: mass is the mass of the satellite; F is the thrust;

[0136] Calculate the orbital argument corresponding to the start time of the ignition:

[0137] ufire_t_0=mod(ufire_m_EW-0.5×we×t_thrust_t,2×pi);

[0138] ufire_t_t=mod(ufire_m_EW+0.5×we×t_thrust_t,2×pi);

[0139] Calculate the orbital angle corresponding to the radial ignition start time:

[0140] ufire_r_0=mod(ufire_m_EW-0.5×we×t_thrust_r,2×pi);

[0141] ufire_r_t=mod(ufire_m_EW+0.5×we×t_thrust_r,2×pi); S2.2.10, update the track keeping strategy generation flag:

[0142] If |V_t|>yuzhi_2, where yuzhi_2>yuzhi_1, then SK_step_t=1, indicating that track control is required; otherwise: SK_step_t=0, indicating that track control is not required this time;

[0143] If |V_r|>yuzhi_2, where yuzhi_2>yuzhi_1, then SK_step_r=1, indicating that radial orbit control is required; otherwise: SK_step_r=0, indicating that radial orbit control is not required this time.

[0144] The step S3 comprises:

[0145] S3.1, perform directional and radial orbit control:

[0146] If mod(Lon+alphaG-ufire_t_0+pi,2×pi)-pi>0

[0147] And mod(Lon+alphaG-ufire_t_0+pi,2×pi)-pi<=we×Crl_period

[0148] And t_thrust_t≠0, and SK_step_t==1

[0149] Among them: Crl_period is the control period of the satellite control system, generally between 0.05s and 1s;

[0150] Then the satellite's east side orbit control engine is started to ignite to slow down the satellite. The engine ignition duration is t_thrust_t;

[0151] If mod(Lon+alphaG-ufire_r_0+pi,2×pi)-pi>0

[0152] And mod(Lon+alphaG-ufire_r_0+pi,2×pi)-pi<=we×Crl_period

[0153] And t_thrust_r≠0, and SK_step_r==1

[0154] Then the satellite's east side orbit control engine is started and ignited, and the engine ignition duration is t_thrust_r;

[0155] S3.2. Update the track keeping strategy generation flag:

[0156] If SK_step_t==1, and mod(Lon+alphaG-ufire_t_t+pi,2×pi)-pi>we×Crl_period, then:

[0157] SK_step_t = 0;

[0158] If SK_step_r == 1, and mod(Lon+alphaG-ufire_r_t+pi,2×pi)-pi>we×Crl_period, then:

[0159] SK_step_r=0.

[0160] In summary, the method for simultaneously maintaining the eccentricity vector and orbit angle of a geosynchronous orbit of the present invention can realize the capture and maintenance of parameters within the orbital plane, which belongs to the minimum fuel consumption maintenance.

[0161] Although the content of the present invention has been described in detail through the above preferred embodiments, it should be appreciated that the above description should not be considered as a limitation of the present invention. After reading the above content, it will be apparent to those skilled in the art that various modifications and substitutions of the present invention will occur. Therefore, the protection scope of the present invention should be limited by the appended claims.

Claims

1. A method for simultaneously maintaining the eccentricity vector and orbit angle of a geosynchronous orbit, characterized in that: The steps include: S1, determine the calculation timing of the position protection strategy; S2, calculate the position protection strategy; S3, according to the calculated orbit strategy, executing the corresponding switch command of the electric propulsion engine at the orbit angle that meets the corresponding conditions; The step S2 comprises: S2.1, calculate the tangential velocity increment delta_V required to maintain the eccentricity; S2.2, determine the orbital argument of the position and orbit control, if delta_V<0.000000001, select the orbital argument ufire_m_EW corresponding to the ignition point, ufire_m_EW=π, the corresponding angle is 180°, V_r=0, V_r is the radial ignition speed increment; if delta_V>=0.000000001, execute step S2.2.1; S2.2.1, calculate the temporary ignition point: u_std = atan2(eyMEAN, exMEAN); eyMEAN and exMEAN represent the y-component and x-component of the satellite orbit eccentricity vector, respectively; S2.2.2, calculate the error between the current geographic accuracy and the target geographic accuracy: delta_lon=lonMEAN-lon_Target, where lon_Target is the target fixed point longitude; S2.2.3, calculate the orbit angle control accuracy: u_yuzhi=-0.125×ddu×86400 2 , where ddu is the drift acceleration of the satellite under natural perturbation; S2.2.4, calculate the relative drift velocity dn0 of the satellite at the current moment: dn0 = -1.5 × we × (smaMEAN - sma0) / sma0, where smaMEAN is the semi-major axis of the satellite's horizontal orbit at the current moment; S2.2.5, calculate the velocity increment V_t along the trajectory: If delta_lon+u_yuzhi<0 V_t=-1 / 3×sma0×(sqrt(2×ddu×(delta_lon+u_yuzhi))-dn0); If delta_lon+u_yuzhi>=0 V_t=0; S2.2.6, calculate the angle theta of the ignition point: If delta_V>|V_t| theta=pi / 2-asin(|V_t| / delta_V) If delta_V<=|V_t| theta = 0; S2.2.7, calculate the orbital argument ufire_m_EW corresponding to the ignition point: ufire_m_EW=u_std+sign_vr×theta Wherein, sign_vr=1 means that the satellite-z-plane thruster is used, and the radial velocity increment is negative; sign_vr=-1 means that the satellite z-plane thruster is used, and the radial velocity increment is positive; S2.2.8, calculate the radial velocity increment V_r: V_r=sign_vr×delta_V×cos(pi / 2-theta)×2 S2.2.9, calculate the orbital arguments corresponding to the start and end times of ignition: Calculate the thrust ignition duration: t_thrust_t = mass × |V_t| / F; Calculate radial ignition duration: t_thrust_r = mass × |V_r| / F; Where: mass is the mass of the satellite; F is the thrust; Calculate the orbital argument corresponding to the start time of the ignition: ufire_t_0=mod(ufire_m_EW-0.5×we×t_thrust_t,2×pi); ufire_t_t=mod(ufire_m_EW+0.5×we×t_thrust_t,2×pi); Among them, ufire_t_0 is the orbit angle corresponding to the start time of the track ignition; ufire_t_t is the orbital argument corresponding to the end time of the ignition in the track direction; Calculate the orbital angle corresponding to the radial ignition start time: ufire_r_0=mod(ufire_m_EW-0.5×we×t_thrust_r,2×pi); ufire_r_t=mod(ufire_m_EW+0.5×we×t_thrust_r,2×pi); ufire_r_0: The radial ignition start time corresponds to the orbital argument; ufire_r_t: The orbital argument corresponding to the radial ignition termination time; S2.2.10, Update track keeping strategy generation flag: If |V_t|>yuzhi_2, where yuzhi_2>yuzhi_1, then SK_step_t=1, indicating that track control is required; otherwise: SK_step_t=0, indicating that track control is not required this time; If |V_r|>yuzhi_2, where yuzhi_2>yuzhi_1, then SK_step_r=1, indicating that radial orbit control is required; otherwise: SK_step_r=0, indicating that radial orbit control is not required this time.

2. The method for simultaneously maintaining the geosynchronous orbit eccentricity vector and orbital argument according to claim 1, characterized in that: The step S1 includes: if the current satellite local time theta_s1 is greater than the agreed local time theta_s0, and the orbit keeping execution end flag: SK_step_t==0, SK_step_r==0, it is considered that the time for calculating the position keeping strategy has arrived, otherwise the time for calculating the position keeping strategy has not arrived.

3. The method for simultaneously maintaining the geosynchronous orbit eccentricity vector and orbital argument according to claim 2, characterized in that: The current satellite local time theta_s1 is greater than the agreed local time theta_s0: Mod(theta_s1-theta_s0+12,24)-12>0 theta_s1=12 / pi×[Lon+alphaG-atan2(ys,xs)]-12 Among them: Mod is the remainder function, ys is the projection of the current sun position on the y-axis of the J2000 coordinate system; xs is the projection of the current sun position on the x-axis of the J2000 coordinate system; Lon is the geographic longitude of the current sub-satellite point of the satellite; alphaG is the Greenwich sidereal hour angle at the current moment; atan2() is the inverse tangent function.

4. The method for simultaneously maintaining the geosynchronous orbit eccentricity vector and orbital argument according to claim 1, characterized in that: The step S3 comprises: S3.1, perform directional and radial orbit control: If mod(Lon+alphaG-ufire_t_0+pi,2×pi)-pi>0 And mod(Lon+alphaG-ufire_t_0+pi,2×pi)-pi<=we×Crl_period And t_thrust_t≠0, and SK_step_t==1 Among them: Crl_period is the control period of the satellite control system; Then the satellite's east side orbit control engine is started to ignite to slow down the satellite. The engine ignition duration is t_thrust_t; If mod(Lon+alphaG-ufire_r_0+pi,2×pi)-pi>0 And mod(Lon+alphaG-ufire_r_0+pi,2×pi)-pi<=we×Crl_period And t_thrust_r≠0, and SK_step_r==1 Then start the satellite's east side position keeping orbit control engine ignition, the engine ignition duration is t_thrust_r, S3.2, update the orbit keeping strategy generation flag: If SK_step_t==1, and mod(Lon+alphaG-ufire_t_t+pi,2×pi)-pi>we×Crl_period, then: SK_step_t = 0; If SK_step_r == 1, and mod(Lon+alphaG-ufire_r_t+pi,2×pi)-pi>we×Crl_period, then: SK_step_r=0.

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