A method for an electric propulsion satellite to enter a near-circular orbit

Abstract

An embodiment of the present invention discloses a method for a near-circular orbit satellite to enter the orbit by electric propulsion, including: within one orbital period, removing the satellite attitude adjustment time to obtain the time of a single thrust applied to the satellite and the finite thrust coefficient; calculating the semi-major axis change amount and the eccentricity vector change amount according to the initial values of the semi-major axis, eccentricity, and argument of perigee of the satellite orbit and the target values of the semi-major axis, eccentricity, and argument of perigee; calculating the total velocity increment of the first group and the second group of tangential thrusts and the latitude argument of the midpoint of the arc segment where the first group and the second group of tangential thrusts are applied according to the semi-major axis change amount and the eccentricity vector change amount; calculating the turn-on and turn-off latitude arguments of the first group and the second group of tangential thrusts according to the total velocity increment of the first group and the second group of tangential thrusts and the latitude argument of the midpoint of the arc segment where the first group and the second group of tangential thrusts are applied; calculating the number of turn-on times of the first group and the second group of tangential thrusts and selecting the execution strategy for the satellite electric propulsion orbital maneuver to enter the orbit.

Description

Technical Field

[0001] The present invention relates to the field of satellite orbit dynamics and control. More specifically, it relates to a method for electrically propelling a satellite into orbit in a nearly circular orbit. Background Art

[0002] At present, after a satellite is launched into space, the satellite's orbit may not be the final working orbit, and the satellite needs to adopt a certain orbit control strategy for a period of time to complete precise orbit injection. In order to reduce the mutual coupling of orbit control and facilitate the calculation of orbit control parameters, the adjustments in the orbit plane and outside the orbit plane can be separated. The adjustments outside the orbit plane include the adjustment of the orbit inclination and the right ascension of the ascending node. A normal thrust can be applied at the ascending node or the descending node, which can adjust the orbit inclination with little influence on the right ascension of the ascending node; a normal impulse can be applied at the highest or lowest point of the latitude argument, which can adjust the right ascension of the ascending node with little influence on the orbit inclination. For the semi-major axis, eccentricity, and argument of perigee in the orbit plane, applying a tangential thrust will change all three of these quantities simultaneously, making the control relatively difficult. The present invention proposes a method for electrically propelling a low-Earth near-circular orbit satellite into the orbit plane to simultaneously adjust the semi-major axis, eccentricity, and argument of perigee. Summary of the Invention

[0003] The purpose of the present invention is to provide a method for electrically propelling a satellite into orbit in a nearly circular orbit to solve at least one of the problems existing in the prior art.

[0004] To achieve the above object, the present invention adopts the following technical solutions:

[0005] The first aspect of the present invention provides a method for electrically propelling a satellite into orbit in a nearly circular orbit, including:

[0006] Within one orbital period, removing the satellite attitude adjustment time to obtain the time of a single thrust applied to the satellite and the finite thrust coefficient;

[0007] According to the initial values of the semi-major axis, eccentricity, and argument of perigee of the satellite orbit, and the target values of the semi-major axis, eccentricity, and argument of perigee, calculating the change amount of the semi-major axis and the change amount of the eccentricity vector;

[0008] According to the change amount of the semi-major axis and the change amount of the eccentricity vector, calculating the total velocity increment of the first group and the second group of tangential thrusts, and the latitude argument at the midpoint of the arc segment where the first group and the second group of tangential thrusts are applied;

[0009] According to the total velocity increment of the first group and the second group of tangential thrusts, and the latitude argument at the midpoint of the arc segment where the first group and the second group of tangential thrusts are applied, calculating the turn-on and turn-off latitude arguments of the first group and the second group of tangential thrusts;

[0010] Calculating the number of turn-ons of the first group and the second group of tangential thrusts and selecting an execution strategy for the satellite electric propulsion orbit maneuver to enter the orbit.

[0011] Optionally, the time of a single thrust applied to the satellite

[0012]

[0013] where t o is the orbital period, a is the semi-major axis, and μ e is the gravitational constant of the Earth, and μ e = 3.986×10 14 m 3 / s 2 Δt p is the satellite attitude adjustment time,

[0014] During the calculation of the single thrust, the change in the argument of latitude of the satellite is

[0015]

[0016] The finite thrust coefficient is obtained

[0017]

[0018] Optionally,

[0019] The change in the semi-major axis Δa = a t - a i

[0020] The change in the eccentricity vector includes

[0021] The first variable Δe x = e t cos(ω t ) - e i cos(ω i ),

[0022] The second variable Δe y = e t sin(ω t ) - e i sin(ω i ),

[0023] where a i 、e i and ω i are the initial values of the semi-major axis, the initial value of the eccentricity, and the initial value of the argument of perigee respectively; a t 、e t and ω t are the target values of the semi-major axis, the target value of the eccentricity, and the target value of the argument of perigee respectively.

[0024] Optionally, the total velocity increments of the first and second groups of tangential thrusts and the latitude argument of the midpoint of the arc segments where the first and second groups of tangential thrusts are applied are calculated based on the change in the semi-major axis and the change in the eccentricity vector, including two cases;

[0025] In the first case,

[0026]

[0027] wherein, the latitude argument u1 of the midpoint of the arc segment where the first group of tangential thrusts is applied can be arbitrarily selected;

[0028] In the second case,

[0029]

[0030] wherein, the directions of the velocity increments of the first and second groups of tangential thrusts are opposite, the latitude arguments of the midpoints of the arc segments where the first and second groups of tangential thrusts are applied differ by 180°, Δv1 is the total velocity increment of the first group of tangential thrusts, u1 is the latitude argument of the midpoint of the arc segment where the first group of tangential thrusts is applied, Δv2 is the total velocity increment of the second group of tangential thrusts, u2 is the latitude argument of the midpoint of the arc segment where the second group of tangential thrusts is applied, and V is the velocity of the satellite.

[0031] Optionally, the start latitude argument and the end latitude argument of the first group of tangential thrusts are respectively

[0032] u 1i = u1 - Δu / 2

[0033] u 1f = u1 + Δu / 2

[0034] The start latitude argument and the end latitude argument of the second group of tangential thrusts are respectively

[0035] u 2i = u2 - Δu / 2

[0036] u 2f = u2 + Δu / 2

[0037] Optionally, the number of start times of the first group of tangential thrusts and the second group of tangential thrusts are respectively

[0038] k1 = round(Δv1 / Δv)

[0039] k2 = round(Δv2 / Δv)

[0040] wherein, Δv is the velocity increment of the satellite caused by a single thrust, Δv = FΔt m / m, F is the thrust amplitude, m is the satellite mass, and round() represents the rounding function.

[0041] Optionally, the satellite electric propulsion orbit maneuvering and in-orbit execution strategy includes

[0042] The first strategy is that when k1≥k2,

[0043] In the first k2 orbits, when the latitude argument is u 1i Apply the first set of tangential thrust, and when the latitude argument is u 1f Stop applying the first set of tangential thrust;

[0044] When the latitude argument is u 2i Apply the second set of tangential thrust, and when the latitude argument is u 2f Stop applying the second set of tangential thrust;

[0045] In the (k1 - k2)-th orbit, when the latitude argument is u 1i Apply the first set of tangential thrust, and when the latitude argument is u 1f Stop applying the first set of tangential thrust.

[0046] Optionally, the satellite electric propulsion orbit maneuvering and in-orbit execution strategy includes the second strategy, when k1 < k2

[0047] In the first k1 orbits, when the latitude argument is u 1i Apply the first set of tangential thrust, and when the latitude argument is u 1f Stop applying the first set of tangential thrust;

[0048] When the latitude argument is u 2i Apply the second set of tangential thrust, and when the latitude argument is u 2f Stop applying the second set of tangential thrust;

[0049] In the (k2 - k1)-th orbit, when the latitude argument is u 2i Apply the second set of tangential thrust, and when the latitude argument is u 2f Stop applying the second set of tangential thrust.

[0050] The second aspect of the present invention provides a computer device, including a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the program, the method provided in the first aspect of the present invention is implemented.

[0051] The third aspect of the present invention provides a computer-readable storage medium, on which a computer program is stored. When the program is executed by a processor, the method provided in the first aspect of the present invention is implemented.

[0052] The beneficial effects of the present invention are as follows:

[0053] A method for a near-circular orbit satellite to enter orbit using electric propulsion provided by the present invention is used to synchronously adjust the semi-major axis, eccentricity, and argument of perigee of the near-circular orbit satellite using electric propulsion; and to achieve in-orbit entry within the orbital plane of the near-circular orbit satellite using electric propulsion. BRIEF DESCRIPTION OF THE DRAWINGS

[0054] The following further elaborates on the specific embodiments of the present invention with reference to the accompanying drawings.

[0055] Figure 1 An exemplary method flow diagram showing an embodiment to which the present invention can be applied is shown.

[0056] Figure 2 A schematic diagram of the thruster power-on scheme showing an embodiment of the present invention is shown.

[0057] Figure 3 A schematic diagram of the orbital altitude time history showing an embodiment of the present invention is shown.

[0058] Figure 4 A schematic diagram of the eccentricity time history showing an embodiment of the present invention is shown.

[0059] Figure 5 A schematic diagram of the argument of perigee time history showing an embodiment of the present invention is shown. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0060] To more clearly illustrate the present invention, the following further elaborates on the present invention with reference to the embodiments and the accompanying Figures 1-5 drawings. Similar components in the drawings are denoted by the same reference numerals. Those skilled in the art should understand that the content specifically described below is illustrative rather than restrictive, and should not be used to limit the protection scope of the present invention.

[0061] The present invention provides a method for a near-circular orbit satellite to enter orbit using electric propulsion, including:

[0062] Step 1: Determine the satellite attitude adjustment time

[0063] Using electric propulsion to simultaneously adjust the semi-major axis, eccentricity, and argument of perigee, the orbital maneuver consists of several tangential thrusts and negative tangential thrusts, that is, positive tangential thrust and negative tangential thrust are respectively applied within one orbital period. A schematic diagram of the thruster power-on scheme within one orbital period is as Figure 2 shown. Since low-orbit satellites usually only have one set of thrusters, which are fixedly connected to the satellite body. Therefore, during the process of switching between positive thrust and negative thrust, it is necessary to adjust the attitude by 180° to achieve thruster direction adjustment.

[0064] Step 2: Within one orbital period, remove the satellite attitude adjustment time to obtain the time of a single thrust applied to the satellite and the finite thrust coefficient;

[0065] During one orbital period, the satellite attitude adjustment time Δt is removed twice p , and divided by 2 to obtain the time Δt of a single thrust m

[0066] Time of a single thrust applied to the satellite

[0067]

[0068] where t o is the orbital period, a is the semi-major axis, μ e is the gravitational constant of the Earth, μ e = 3.986×10 14 m 3 / s 2 Δt p is the satellite attitude adjustment time,

[0069] During the calculation of a single thrust, the change in the argument of latitude of the satellite is

[0070]

[0071] Obtain the finite thrust coefficient

[0072]

[0073] Step 3: Calculate the change in the semi-major axis and the change in the eccentricity vector according to the initial values of the semi-major axis, eccentricity, and argument of perigee of the satellite orbit, and the target values of the semi-major axis, eccentricity, and argument of perigee

[0074] Change in semi-major axis Δa = a t -a i

[0075] The change in the eccentricity vector includes

[0076] The first variable Δe x = e t cos(ω t ) - e i cos(ω i ),

[0077] The second variable Δe y = e t sin(ω t ) - e i sin(ω i ),

[0078] where a i 、e i and ω iThey are respectively the initial value of the semi-major axis, the initial value of the eccentricity, and the initial value of the argument of perigee; a t , e t and ω t They are respectively the target value of the semi-major axis, the target value of the eccentricity, and the target value of the argument of perigee.

[0079] Step Four: Calculate the total velocity increments of the first group and the second group of tangential thrusts and the latitude argument of the midpoint of the arc segments where the first group and the second group of tangential thrusts are applied based on the change in the semi-major axis and the change in the eccentricity vector; it includes two cases;

[0080] In a possible implementation manner, the first case,

[0081]

[0082] wherein, the latitude argument u1 of the midpoint of the arc segment where the first group of tangential thrusts is applied can be arbitrarily selected;

[0083] In a possible implementation manner, the second case,

[0084]

[0085] wherein, the directions of the velocity increments of the first group and the second group of tangential thrusts are opposite, the latitude arguments of the midpoints of the arc segments where the first group and the second group of tangential thrusts are applied differ by 180°, Δv1 is the total velocity increment of the first group of tangential thrusts, u1 is the latitude argument of the midpoint of the arc segment where the first group of tangential thrusts is applied, Δv2 is the total velocity increment of the second group of tangential thrusts, u2 is the latitude argument of the midpoint of the arc segment where the second group of tangential thrusts is applied, and V is the velocity of the satellite.

[0086] Step Five: Calculate the turn-on and turn-off latitude arguments of the first group and the second group of tangential thrusts based on the total velocity increments of the first group and the second group of tangential thrusts and the latitude arguments of the midpoints of the arc segments where the first group and the second group of tangential thrusts are applied;

[0087] The turn-on latitude argument and the turn-off latitude argument of the first group of tangential thrusts are respectively

[0088] u 1i = u1 - Δu / 2

[0089] u 1f = u1 + Δu / 2

[0090] The turn-on latitude argument and the turn-off latitude argument of the second group of tangential thrusts are respectively

[0091] u 2i = u2 - Δu / 2

[0092] u 2f= u2 + Δu / 2

[0093] Step Six, calculate the number of startup times of the first group and the second group of tangential thrusts and select the satellite electric propulsion orbital maneuvering into-orbit execution strategy.

[0094] Preferably, the number of startup times of the first group of tangential thrust and the second group of tangential thrust are respectively

[0095] k1 = round(Δv1 / Δv)

[0096] k2 = round(Δv2 / Δv)

[0097] where Δv is the satellite velocity increment caused by a single thrust, Δv = FΔt m / m, F is the thrust amplitude, m is the satellite mass, and round() represents the rounding function.

[0098] In a possible implementation, the satellite electric propulsion orbital maneuvering into-orbit execution strategy includes

[0099] The first strategy, when k1 ≥ k2,

[0100] In the first k2 orbits, when the latitude argument is u 1i , start applying the first group of tangential thrusts (the thrust direction is determined by the sign of Δv1), and when the latitude argument is u 1f , stop applying the first group of tangential thrusts; when the latitude argument is u 2i , start applying the second group of tangential thrusts (the thrust direction is determined by the sign of Δv2), and when the latitude argument is u 2f , stop applying the second group of tangential thrusts.

[0101] In the (k1 - k2)th orbit, when the latitude argument is u 1i , start applying the first group of tangential thrusts (the thrust direction is determined by the sign of Δv1), and when the latitude argument is u 1f , stop applying the first group of tangential thrusts.

[0102] In a possible implementation, the satellite electric propulsion orbital maneuvering into-orbit execution strategy includes the second strategy, when k1 < k2

[0103] In the first k1 orbits, when the latitude argument is u 1i , start applying the first group of tangential thrusts (the thrust direction is determined by the sign of Δv1), and when the latitude argument is u 1f , stop applying the first group of tangential thrusts; when the latitude argument is u 2i , start applying the second group of tangential thrusts (the thrust direction is determined by the sign of Δv2), and when the latitude argument is u 2f , stop applying the second group of tangential thrusts.

[0104] In the k2 - k1 circle, when the latitude argument is u 2i , start applying the second set of tangential thrusts (the thrust direction is determined by the sign of Δv2), and when the latitude argument is u 2f , stop applying the second set of tangential thrusts.

[0105] A method for electrically propelled orbit injection in the orbital plane of a low - earth near - circular orbit satellite proposed by the present invention can achieve simultaneous adjustment of the semi - major axis, eccentricity, and argument of perigee.

[0106] In a specific embodiment, the initial and target values of the satellite orbit are shown in Table 1 below. The satellite mass is 300 kg, and the thrust amplitude of the electric propulsion is 12 mN.

[0107] Table 1 Initial and target values of the satellite orbit

[0108] Orbital elements Initial value Target value Orbit altitude, km 500 510 Eccentricity 0.003 0.001 Argument of perigee, deg 20 90

[0109] Step 1: Determine the satellite attitude adjustment time Δt p

[0110] According to engineering experience, the time for the satellite to adjust its attitude by 180° is Δt p = 1104 sec.

[0111] Step 2: Determine the time of a single - thrust and the finite - thrust coefficient

[0112] The time of a single - thrust applied to the satellite

[0113]

[0114] Δt m = 1735 sec

[0115] where t o is the orbital period, a is the semi - major axis, μ e is the gravitational constant of the earth, μ e = 3.986×10 14 m 3 / s 2 Δt p is the satellite attitude adjustment time,

[0116] During a single - thrust process, the change in the latitude argument of the satellite is

[0117]

[0118] Δu = 110°

[0119] The finite - thrust coefficient

[0120]

[0121] k u = 0.853

[0122] Step 3: Calculate the semi-major axis change amount and the eccentricity vector change amount based on the initial values of the semi-major axis, eccentricity, and argument of perigee, and the target values of the semi-major axis, eccentricity, and argument of perigee

[0123] Calculate the semi-major axis change amount Δa and the first variable Δe of the eccentricity vector change amount according to the initial value target values of the semi-major axis, eccentricity, and argument of perigee in Table 1 x and the second variable Δe y

[0124] Δa = 10 km

[0125] Δe x = -0.00282

[0126] Δe y = -2.61×10 -5

[0127] Step 4: Determine the total velocity increment Δv1 of the first group of tangential thrusts, the latitude argument u1 at the midpoint of the arc segment where the first group of tangential thrusts is applied, the total velocity increment Δv2 of the second group of tangential thrusts, the latitude argument u2 at the midpoint of the arc segment where the second group of tangential thrusts is applied, the relative semi-major axis change amount Δa, and the eccentricity vector change amount Δe x and Δe y The relationship is

[0128]

[0129] According to the above relationship, use the corresponding formula to calculate the total velocity increment Δv1 of the first group of tangential thrusts, the latitude argument u1 at the midpoint of the arc segment where the first group of tangential thrusts is applied, the total velocity increment Δv2 of the second group of tangential thrusts, and the latitude argument u2 at the midpoint of the arc segment where the second group of tangential thrusts is applied

[0130] Δv1 = 9.05 m / s

[0131] u1 = 180.53°

[0132] Δv2 = -3.52 m / s

[0133] u2 = 0.53°

[0134] Step 5: Determine the on-off latitude arguments of the first group of tangential thrusts and the second group of tangential thrusts

[0135] The on-off latitude arguments of the first group of tangential thrusts are respectively

[0136] u1i = 125.53°

[0137] u 1f = 235.53°

[0138] The start latitude amplitude angle and shutdown latitude amplitude angle of the second set of tangential thrusts are respectively

[0139] u 2i = 305.53°

[0140] u 2f = 55.53°

[0141] Step 6, determine the number of startup times of the first set of tangential thrust and the second set of tangential thrust

[0142] Within one orbital period, the satellite velocity increment caused by a single thrust is Δv = 0.0695 m / s. Then the number of startup times of the first set of tangential thrust and the second set of tangential thrust are

[0143] k1 = 130

[0144] k2 = 51

[0145] Step 7, determine the satellite electric propulsion orbital maneuvering and in-orbit execution strategy.

[0146] According to Step 6, the relationship between k1 and k2 is k1 ≥ k2. Then the electric propulsion orbital maneuvering and in-orbit execution strategy is as follows:

[0147] In the first k2 = 51 laps, at the latitude amplitude angle u 1i = 125.53°, start applying the first set of tangential thrust (thrust direction is positive). When the latitude amplitude angle u 1f = 235.53°, stop applying the first set of tangential thrust;

[0148] At the latitude amplitude angle u 2i = 305.53°, start applying the second set of tangential thrust (thrust direction is negative). When the latitude amplitude angle u 2f = 55.53°, stop applying the second set of tangential thrust;

[0149] In the (k1 - k2) = 79th lap, at the latitude amplitude angle u 1i = 125.53°, start applying the first set of tangential thrust (thrust direction is positive). When the latitude amplitude angle u 1f = 235.53°, stop applying the first set of tangential thrust.

[0150] As Figure 3 shown in the orbital altitude time history, through the electric propulsion thrust, the orbital altitude of the satellite is changed from 500 km to 510 km, achieving the expected orbital altitude adjustment.

[0151] As Figure 4 shown in the eccentricity time history, through the electric propulsion thrust, the eccentricity of the satellite changes from 0.003 to 0.001, achieving the expected eccentricity adjustment.

[0152] As Figure 5 shown in the argument of perigee time history, through the electric propulsion thrust, the argument of perigee of the satellite changes from 20° to 90°, achieving the expected argument of perigee adjustment.

[0153] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by terms such as "upper", "lower", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation. Therefore, it should not be construed as a limitation to the present invention. Unless otherwise clearly specified and defined, the terms "mounted", "connected", and "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0154] It should also be noted that in the description of the present invention, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprising", "including" or any other variation thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements not only includes those elements, but also includes other elements not expressly listed, or also includes elements inherent to such process, method, article or device. Without further limitation, an element defined by the statement "including a..." does not exclude the existence of additional identical elements in the process, method, article or device including the said element.

[0155] Obviously, the above embodiments of the present invention are only examples for clearly illustrating the present invention, rather than limitations on the implementation manners of the present invention. For those of ordinary skill in the art, other different forms of changes or variations can be made on the basis of the above description. It is impossible to list all the implementation manners here. Any obvious changes or variations derived from the technical solutions of the present invention still fall within the protection scope of the present invention.

Claims

1. A method for a near-circular orbit satellite to enter orbit by electric propulsion, characterized in that Including: During one orbital period, removing the satellite attitude adjustment time to obtain the time of a single thrust applied to the satellite and the finite thrust coefficient; Calculating the semi-major axis change amount and the eccentricity vector change amount based on the initial values of the semi-major axis, eccentricity, and argument of perigee of the satellite orbit, and the target values of the semi-major axis, eccentricity, and argument of perigee; Calculating the total velocity increment of the first and second groups of tangential thrusts and the latitude argument of the midpoint of the thrust application arc segment of the first and second groups of tangential thrusts based on the semi-major axis change amount and the eccentricity vector change amount; Calculating the turn-on and turn-off latitude arguments of the first and second groups of tangential thrusts based on the total velocity increment of the first and second groups of tangential thrusts and the latitude argument of the midpoint of the thrust application arc segment of the first and second groups of tangential thrusts; Calculating the number of turn-ons of the first and second groups of tangential thrusts and selecting the satellite electric propulsion orbit maneuvering into-orbit execution strategy.

2. The method according to claim 1, wherein The time of a single thrust applied to the satellite Among them, is the orbital period, , is the semi-major axis, is the gravitational constant of the Earth, , is the satellite attitude adjustment time, During the calculation of the single thrust, the change amount of the satellite's latitude argument is Obtaining the finite thrust coefficient 。 3. The method according to claim 2, wherein Semi-major axis change amount The eccentricity vector change amount includes First variable , Second variable , Among them, , and are the initial value of the semi-major axis, the initial value of the eccentricity, and the initial value of the argument of perigee, respectively; , and are the target value of the semi-major axis, the target value of the eccentricity, and the target value of the argument of perigee, respectively.

4. The method according to claim 3, wherein Calculating the total velocity increment of the first and second groups of tangential thrusts and the latitude argument of the midpoint of the thrust application arc segment of the first and second groups of tangential thrusts based on the semi-major axis change amount and the eccentricity vector change amount includes The first case ， Among them, the latitude argument of the midpoint of the first set of tangential thrust application arc segments can be selected optionally, is the total velocity increment of the first set of tangential thrusts, is the latitude argument of the midpoint of the first set of tangential thrust application arc segments, is the total velocity increment of the second set of tangential thrusts, is the latitude argument of the midpoint of the second set of tangential thrust application arc segments, is the velocity of the satellite; The second case Wherein, the directions of the velocity increments of the first and second groups of tangential thrusts are opposite, and the latitude arguments of the midpoints of the thrust application arcs of the first and second groups of tangential thrusts differ by 180°.

5. The method according to claim 4, wherein The turn-on latitude argument and the turn-off latitude argument of the first group of tangential thrusts are respectively The turn-on latitude argument and the turn-off latitude argument of the second group of tangential thrusts are respectively 。 6. The method according to claim 4, wherein The number of turn-ons of the first group of tangential thrusts and the second group of tangential thrusts are respectively Among them, is the satellite velocity increment caused by a single thrust, , is the thrust amplitude, is the satellite mass, represents the rounding function.

7. The method according to claim 6, wherein The satellite electric propulsion orbit maneuvering into-orbit execution strategy includes The first strategy, when is In front ring, when the latitude argument is , start applying the first set of tangential thrusts, and when the latitude argument is , stop applying the first set of tangential thrusts; When the latitude argument is , the second set of tangential thrusts starts to be applied. When the latitude argument is , the application of the second set of tangential thrusts stops; At the th circle, when the latitude argument is , start applying the first set of tangential thrusts, and when the latitude argument is , stop applying the first set of tangential thrusts.

8. The method according to claim 6, characterized in that The satellite electric propulsion orbit maneuvering and injection execution strategy includes a second strategy, when When In front ring, when the latitude argument is , start applying the first set of tangential thrusts, and when the latitude argument is , stop applying the first set of tangential thrusts; When the latitude argument is , the second set of tangential thrusts starts to be applied, and when the latitude argument is , the application of the second set of tangential thrusts stops; At the th circle, when the latitude argument is , start applying the second set of tangential thrusts, and when the latitude argument is , stop applying the second set of tangential thrusts.

9. A computer device, comprising a memory, a processor, and a computer program stored on the memory and executable on the processor, characterized in that When the processor executes the program, it implements the method described in any one of claims 1-8.

10. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the program is executed by the processor, it implements the method described in any one of claims 1-8.

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