Autonomous position holding control method and apparatus for coplanar mounted electric thrusters

Abstract

The present application relates to the technical field of spacecraft attitude orbit control, and particularly relates to a kind of coplanar installation electric thruster autonomous position keeping control method and device.Method includes: when receiving autonomous position keeping control instruction, based on current star time and pre-determined position keeping control period, the orbit inclination vector variation in position keeping control period is determined;Based on the orbit inclination vector variation in position keeping control period, the total firing duration in position keeping control period is determined to determine the firing duration corresponding to each operating day;Based on the orbit inclination vector variation in position keeping control period, the firing point right ascension is determined;Based on the firing duration corresponding to each operating day and the firing point right ascension, the coplanar installation electric thruster is used to autonomously carry out position keeping control in each operating day.This scheme can predict the position keeping parameters of multiple days, and the calculation is simple and easy to implement.

Description

Technical Field

[0001] This invention relates to the field of spacecraft attitude and orbit control technology, and in particular to an autonomous position-holding control method and apparatus for coplanar mounted electric thrusters. Background Technology

[0002] Electric propulsion, a type of low-thrust propulsion system, is characterized by its light weight, high specific impulse, and low thrust. After a satellite is positioned, coplanar electric thrusters can be used to maintain its north-south position throughout its entire lifespan.

[0003] Existing position-keeping control methods for coplanar-mounted electric thrusters mostly rely on trajectory recursion to predict the position error for the next day. However, in practical applications, position-keeping control is typically required five days a week, and the trajectory recursion process is complex and computationally intensive. Therefore, existing position-keeping control methods for coplanar-mounted electric thrusters not only have short prediction times but also require significant computational resources.

[0004] Therefore, there is an urgent need for an autonomous position-holding control method for coplanar-mounted electric thrusters. Summary of the Invention

[0005] To address the problems of existing position-holding control methods for coplanar-mounted electric thrusters, which have both short prediction times and high computational loads, this invention provides an autonomous position-holding control method and apparatus for coplanar-mounted electric thrusters.

[0006] In a first aspect, embodiments of the present invention provide an autonomous position-holding control method for a coplanarly mounted electric thruster, comprising:

[0007] When an autonomous position-keeping control command is received, the orbital inclination vector change within the position-keeping control period is determined based on the current satellite time and a pre-determined position-keeping control period.

[0008] Based on the change in the orbital tilt vector within the position holding control cycle, the total ignition duration within the position holding control cycle is determined, thereby determining the ignition duration corresponding to each operating day within the position holding control cycle; wherein, the position holding control cycle is 7 days and the operating day is 5 days;

[0009] The right ascension of the ignition point is determined based on the change in the orbital inclination vector during the position holding control cycle.

[0010] Based on the ignition duration and right ascension of the ignition point for each operating day, the coplanar electric thrusters autonomously maintain north-south position control within each operating day.

[0011] Secondly, embodiments of the present invention also provide an autonomous position-keeping control device based on the method described in any embodiment of this specification, comprising:

[0012] The first determining unit is used to determine the change in orbital inclination vector within the position holding control period based on the current satellite time and a pre-determined position holding control period when an autonomous position holding control command is received.

[0013] The second determining unit is used to determine the total ignition duration within the position holding control cycle based on the change in the orbital inclination vector within the position holding control cycle, so as to determine the ignition duration corresponding to each operating day within the position holding control cycle; wherein, the position holding control cycle is 7 days and the operating day is 5 days;

[0014] The third determining unit is used to determine the right ascension of the ignition point based on the change in the orbital inclination vector during the position holding control cycle;

[0015] The control unit is used to autonomously maintain the north-south position of the electric thruster within each operating day, based on the ignition duration and the right ascension of the ignition point corresponding to each operating day.

[0016] Thirdly, embodiments of the present invention also provide a computing device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, it implements the method described in any embodiment of this specification.

[0017] Fourthly, embodiments of the present invention also provide a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the methods described in any embodiment of this specification.

[0018] This invention provides an autonomous position-holding control method and apparatus for coplanar-mounted electric thrusters. By using the current satellite time, a pre-determined position-holding control period, and current onboard orbital parameters, the method autonomously calculates the position-holding ignition start time and duration for multiple operational days, enabling the satellite to autonomously complete a multi-day position-holding process. Furthermore, the thrust directions of the coplanar-mounted electric thrusters in this scheme are radial and normal, which is more in line with practical satellite engineering applications. Therefore, this scheme is not only operable in predicting time and thrust direction, and more in line with practical engineering applications, but also simple to calculate, does not involve orbital recursion processes, and is more conducive to onboard implementation. Attached Figure Description

[0019] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0020] Figure 1 This is a flowchart of an autonomous position-holding control method for a coplanar mounted electric thruster according to an embodiment of the present invention;

[0021] Figure 2 This is a schematic diagram illustrating the north-south position maintenance according to an embodiment of the present invention;

[0022] Figure 3 This is a hardware architecture diagram of a computing device provided in an embodiment of the present invention;

[0023] Figure 4 This is a structural diagram of an autonomous position holding control device for a coplanarly mounted electric thruster provided in an embodiment of the present invention. Detailed Implementation

[0024] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0025] The following describes the specific implementation of the above concept.

[0026] Please refer to Figure 1 This invention provides an autonomous position-holding control method for a coplanar mounted electric thruster, the method comprising:

[0027] Step 100: When an autonomous position-keeping control command is received, the orbital inclination vector change within the position-keeping control period is determined based on the current satellite time and a pre-determined position-keeping control period.

[0028] Step 102: Based on the change in the orbital inclination vector within the position holding control cycle, determine the total ignition duration within the position holding control cycle, and then determine the ignition duration corresponding to each operating day within the position holding control cycle; wherein, the position holding control cycle is 7 days and the operating day is 5 days;

[0029] Step 104: Determine the right ascension of the ignition point based on the change in the orbital inclination vector during the position holding control cycle;

[0030] Step 106: Based on the ignition duration and right ascension of the ignition point corresponding to each operation day, the coplanar-mounted electric thrusters are used to autonomously maintain the north-south position within each operation day; wherein, the thrust direction of the coplanar-mounted electric thrusters is the radial direction and the normal direction.

[0031] In this embodiment of the invention, the satellite autonomously calculates the start time and duration of position-holding ignition for multiple operational days using the current satellite time, a pre-determined position-holding control period, and current onboard orbital parameters, enabling the satellite to autonomously complete a multi-day position-holding process. Furthermore, the thrust directions of the coplanarly mounted electric thrusters in this scheme are radial and normal, which better aligns with practical satellite engineering applications. Therefore, this scheme is not only operable in predicting time and thrust direction, and more in line with practical engineering applications, but also simple to calculate, does not involve orbital recursion, and is more conducive to onboard implementation.

[0032] It should be noted that existing position-holding control methods for coplanar-mounted electric thrusters involve thrusters installed in a manner that directs the thrust in both tangential and normal directions. This differs from practical engineering applications, making it difficult to implement and impractical. In contrast, the control method proposed in this scheme directs the thrust of the coplanar-mounted electric thruster in both radial and normal directions. This method can be referenced from [the original text]. Figure 2 This scheme utilizes a two-degree-of-freedom vector adjustment mechanism to achieve coplanar mounting of the electric thruster, providing thrust in both the radial and normal directions, thus enabling control of the track inclination and eccentricity. The radial direction points towards the Earth's center, while the normal direction is perpendicular to the track plane.

[0033] For step 100:

[0034] In some implementations, step 100, "determining the change in orbital inclination vector within the position-holding control period based on the current satellite time and a predetermined position-holding control period," may include steps S1-S4:

[0035] S1, based on the current star time, the initial star time and J2000 time, calculate the Julian century, and determine the ascending node longitude based on the Julian century;

[0036] S2, based on the ecliptic longitude of the ascending node, calculates the rate of change of the orbital inclination vector over one year;

[0037] S3, transform the zero tilt vector of the instantaneous true coordinate system to the J2000 coordinate system, and obtain the magnitude of the orbit tilt vector in the x and y directions corresponding to the zero tilt vector in the instantaneous true coordinate system in the J2000 coordinate system;

[0038] S4, based on the rate of change of the track inclination vector within one year, the magnitude of the track inclination vector in the x and y directions corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, and the predetermined position holding control period, determines the change of the track inclination vector within the position holding control period.

[0039] In some implementations, the rate of change of the orbital inclination vector over a year is determined as follows, i.e., steps S1-S2 may include:

[0040] The Julian century can be calculated using the following formula:

[0041]

[0042] in,

[0043] t J2000 =t+473342400

[0044] In the formula, T represents the Julian century, and t J2000 t represents the number of seconds between the current satellite time and the J2000 time, and t represents the number of seconds between the current satellite time and the initial satellite time.

[0045] It should be noted that J2000.0 is an epoch used in astronomy. The prefix "J" indicates that it is a Julian epoch, not a Bessel epoch. It refers to the Julian date TT2451545.0, or TT time January 1, 2000, 12:00.

[0046] Calculate the ecliptic longitude of the ascending node using the following formula:

[0047] Ω l =2.182439 - 33.757 × T

[0048] In the formula, T represents the Julian century, and Ω l The ecliptic longitude of the ascending node;

[0049] The rate of change of the orbital inclination vector over a year can be calculated using the following formula:

[0050] dΔi′ x =0.097cosΩ l +0.852

[0051] dΔi′ y =0.13sinΩ l

[0052] In the formula, dΔi′ x and dΔi′ y These represent the rates of change of the orbital inclination vector in the x-direction and the rate of change in the y-direction over a year, respectively, Ω. l The ecliptic longitude of the ascending node;

[0053] The zero tilt vector of the instantaneous true coordinate system is transformed to the J2000 coordinate system using the following formula, i.e., step S3 includes:

[0054] Δi xTOD =KTodcof1*C PR (2,0)

[0055] Δi yTOD =KTodcof1*C PR (2,1)

[0056] In the formula, Δi xTOD and Δi yTOD These represent the x- and y-direction magnitudes of the orbital inclination vector corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, respectively. KTodcof1 is a pre-set coordinate system transformation adjustment coefficient, and C... PR (2,0) is the element in the 3rd row and 1st column of the precession nutation transformation matrix, where 2 and 0 are the row and column numbers, respectively. PR (2,1) is the element in the 3rd row and 2nd column of the precession nutation transformation matrix, where 2 and 1 are the row number and column number, respectively.

[0057] In some implementations, the change in the track inclination vector during the position holding control cycle is calculated using the following formula, i.e., step S4 may include:

[0058] Δi x =dΔi x ′ / 365*TJetPeriod*Kcomp+(i·sin(Ω)-Δi xTOD R2D

[0059] Δi y =dΔi y ′ / 365*TJetPeriod*Kcomp+(-i·cos(Ω)-Δi yTOD R2D

[0060] In the formula, Δi x and Δi y dΔi represents the change in the orbital inclination vector in the x-direction and the change in the y-direction, respectively, during the position holding control cycle. x ′ and dΔi y ′ represents the rate of change of the orbital inclination vector in the x and y directions respectively over one year; TJetPeriod is the predetermined position-keeping control period, which is 7; Kcomp is the correction value, which defaults to 1; i is the satellite's orbital inclination; Ω is the right ascension of the satellite's ascending node; Δi xTOD and Δi yTODR2D represents the magnitudes of the orbital inclination vector in the x and y directions corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, respectively, and R2D is the conversion between radians and degrees.

[0061] Regarding step 102:

[0062] In some implementations, step 102 may include steps A1-A3:

[0063] A1, based on the change in track inclination vector during the position holding control cycle, calculate the magnitude of track inclination change, and determine the required velocity increment in the normal direction based on the magnitude of track inclination change.

[0064] In this embodiment of the invention, the magnitude of the change in track inclination Δi is calculated using the following formula. jet The required velocity increment Δv in the normal direction jet :

[0065]

[0066]

[0067] In the formula, Δi x and Δi y These represent the changes in the track inclination vector in the x-direction and the changes in the y-direction, respectively, during the position holding control cycle.

[0068] A2 determines the total ignition duration within the position-holding control cycle based on the required velocity increment in the normal direction, the satellite's mass, and the satellite's nominal thrust.

[0069] In this embodiment of the invention, the total ignition duration ΔT within the position holding control cycle is calculated using the following formula. jet :

[0070]

[0071] In the formula, Msa is the satellite mass, and F jetnorm The nominal thrust is calculated using the south main component as the nominal value by default, corresponding to 0.0297 N.

[0072] A3. Divide the total ignition duration within the position holding control cycle into 5 equal parts to obtain the ignition duration corresponding to each operating day.

[0073] In this embodiment, position holding control is typically required five days a week, so the number of operating days is 5. The total ignition duration ΔT is calculated as follows. jet Dividing the data into 5 equal parts, we can obtain the ignition duration ΔT for each operating day. id .

[0074] In some implementations, after dividing the total ignition duration within the position holding control cycle into five equal parts in A3, before obtaining the ignition duration corresponding to each operating day, the following steps are also included:

[0075] Obtain the time limit for coplanar mounted electric thrusters;

[0076] Determine whether the total ignition time, after being divided equally, exceeds the time limit.

[0077] If the limit is exceeded, the time limit will be set as the ignition time corresponding to each operation day.

[0078] If the total ignition time is not exceeded, the total ignition time will be divided equally and used as the ignition time for each operation day.

[0079] In this embodiment, to ensure the safe operation of the coplanar-mounted electric thruster, a daily working time limit is generally set for the coplanar-mounted electric thruster. Under normal circumstances, the total ignition time after being divided equally is less than the limit. However, in order to further protect the coplanar-mounted electric thruster, it is necessary to determine whether the total ignition time after being divided equally exceeds the limit. If it exceeds, the limit is determined as the ignition time corresponding to each operating day. If it does not exceed, the total ignition time after being divided equally is used as the ignition time corresponding to each operating day.

[0080] Regarding step 104:

[0081] In some implementations, the ignition point ascension is calculated using the following formula:

[0082] OmgF=atan2(Δi x ,Δi y )

[0083] In the formula, OmgF is the right ascension of the ignition point, and Δi x and Δi y These represent the changes in the track inclination vector in the x-direction and the changes in the y-direction, respectively, during the position holding control cycle.

[0084] Regarding step 106:

[0085] In this step, the ignition duration ΔT corresponding to each operating day in step 102 is obtained. id Following the ignition point right ascension determined in step 104, ignition is performed once at the ascending node and once at the descending node on each operating day. The sum of the two ignition times must equal the corresponding ignition time ΔT for each operating day. id Therefore, on each operational day, when the satellite reaches the right ascension of the two ignition points, it activates the coplanar-mounted electric thrusters for position holding control, until the total duration reaches the ignition duration ΔT. id Stop at a certain time to allow for independent operation to maintain north-south position.

[0086] In some implementations, steps 094-098 may be included before step 100:

[0087] Step 094: Based on the satellite's fixed position, determine the direction of the satellite's mean longitude drift caused by the perturbation acceleration and the change in the satellite's semi-major axis during the position-holding control period under the influence of Earth's perturbation force; wherein, the position-holding control period is 7 days and the operation day is 5 days;

[0088] Step 096: Determine the direction of satellite drift caused by the tangential thrust generated by the coplanar electric thrusters when the center of mass of the entire satellite deviates by 1 mm in the +X or -X direction, as well as the amount of change of the semi-major axis of the satellite during the position holding control period. Based on the direction of satellite drift in longitude caused by perturbation acceleration and the amount of change of the semi-major axis of the satellite during the position holding control period under the action of Earth perturbation, determine the target offset direction and target offset of the center of mass of the entire satellite.

[0089] Step 098: Adjust the position of the satellite's center of mass based on the target offset direction and target offset amount.

[0090] In traditional methods, after a satellite reaches its designated orbit, coplanar-mounted electric thrusters are used to maintain its north-south position throughout its entire lifespan. Then, chemical thrusters are used to maintain its east-west position, ultimately achieving orbit control. This decoupling of north-south and east-west position maintenance requires a significant amount of chemical propellant for the chemical thrusters. To reduce propellant consumption, the north-south and east-west position maintenance tasks can be coupled, using coplanar-mounted electric thrusters to perform both north-south and east-west position maintenance simultaneously.

[0091] In this embodiment, the position of the satellite's center of mass is adjusted based on the target offset direction and the target offset amount. When the satellite performs north-south position-keeping control in orbit, the offset of the satellite's center of mass will generate tangential thrust, thereby raising or lowering the semi-major axis. This can precisely offset the change in the satellite's semi-major axis during the position-keeping control period under the influence of Earth's perturbation. Therefore, the east-west position-keeping control task is completed at the same time as the north-south position-keeping control. Not only is there no need to waste extra time on east-west position-keeping, but the tangential thrust generated during north-south position-keeping control can also be used to adjust the east-west position, which can also reduce the consumption of chemical propellant.

[0092] Regarding step 094:

[0093] In this embodiment of the invention, in order to determine the target offset direction and target offset amount of the whole satellite's center of mass shift, it is necessary to determine the direction of the satellite's mean longitude drift caused by the perturbation acceleration and the amount of change of the satellite's semi-major axis during the position-maintaining control period under the action of the Earth's perturbation force, based on the satellite's fixed position.

[0094] For example, when the satellite's fixed position is at 134 degrees east longitude, the local drift acceleration is -0.00163 degrees / day^2. Since the drift acceleration is negative, the perturbation acceleration causes the satellite's mean longitude to drift westward. Under the influence of the Earth's perturbation force, the satellite's semi-major axis rises by about 127m during the position-maintaining control period.

[0095] Regarding steps 096 and 098:

[0096] You can refer to this. Figure 2 When the center of mass of the satellite shifts, the electric thrusters generate tangential thrust during tilt control. When the satellite shifts westward and its semi-major axis rises due to perturbation, the east-west position of the satellite can be adjusted by the required direction and magnitude of tangential thrust so that the satellite can operate in the predetermined orbit.

[0097] In some implementations, the change in the semi-major axis of the satellite during the position-holding control period caused by the tangential thrust generated by the coplanarly mounted electric thrusters for every 1 mm deviation of the satellite's center of mass is determined by the following steps B1-B4:

[0098] Step B1: Obtain the tangential and normal thrust of the primary and backup electric thrusters, which are coplanarly mounted, when the center of mass of the entire satellite deviates by 1 mm.

[0099] In this embodiment, the coplanarly mounted electric thrusters include a primary electric thruster and a backup electric thruster, and the thrust of the primary electric thruster and the backup electric thruster are different. It is necessary to obtain the tangential thrust and normal thrust of the primary electric thruster and the tangential thrust and normal thrust of the backup electric thruster respectively.

[0100] Step B2: Based on the tangential thrust, normal thrust, and the pre-set normal daily average velocity increment, calculate the corresponding tangential daily average velocity increment when the primary electric thruster is started and when the backup electric thruster is started.

[0101] In some implementations, the tangential daily average velocity increment is calculated using the following formula:

[0102]

[0103] In the formula, ΔV t ΔV represents the daily average velocity increment in the tangential direction. n For the daily average velocity increment in the normal direction, F t For tangential thrust, F nThis is the normal thrust.

[0104] In this embodiment, the daily average normal velocity increment is determined as follows: the tilt drift velocity is approximately 0.75° to 0.95° per year, and the velocity increment Δv maintained at the north-south position is between 41 m / s and 51 m / s. The average value Δv = 46 m / s is used in the design. Dividing Δv by 365 days yields the daily average normal velocity increment ΔV. n It is 0.126 m / s.

[0105] In this embodiment, since the thrust of the primary electric thruster and the backup electric thruster are different, it is necessary to use the above formula to calculate the tangential daily average velocity increment when using the primary electric thruster alone and the tangential daily average velocity increment when using the backup electric thruster alone.

[0106] Step B3: Based on the vitality formula and the daily average tangential velocity increment, calculate the daily average semi-major axis change when the primary electric thruster is started and when the backup electric thruster is started.

[0107] In some implementations, the daily average change in the semi-major axis is calculated using the following formula:

[0108]

[0109] In the formula, ΔV t denoted as the daily average velocity increment, v as the satellite velocity, μ as the gravitational constant, r as the satellite radius vector, a as the semi-major axis, and Δa as the daily average change in the semi-major axis.

[0110] From the above formula, we can obtain the relationship Δa=27.4×ΔV t .

[0111] Similarly, by substituting the daily average tangential velocity increment corresponding to the use of the primary electric thruster alone into the formula, we can obtain the daily average semi-major axis change when the primary electric thruster is started; by substituting the daily average tangential velocity increment corresponding to the use of the backup electric thruster alone into the formula, we can obtain the daily average semi-major axis change when the backup electric thruster is started.

[0112] Step B4: Based on the daily average semi-major axis change and the number of days in the position holding control period, determine the amount of semi-major axis change caused by the tangential thrust generated by the coplanar electric thrusters when the center of mass of the satellite deviates by 1 mm during the position holding control period.

[0113] In this embodiment, multiplying the daily average semi-major axis change when the primary electric thruster is activated by 7 yields the semi-major axis change caused by the tangential thrust generated by the primary electric thruster during the position-holding control period when the satellite's center of mass deviates by 1 mm. Multiplying the daily average semi-major axis change when the backup electric thruster is activated by 7 yields the semi-major axis change caused by the tangential thrust generated by the backup electric thruster during the position-holding control period when the satellite's center of mass deviates by 1 mm.

[0114] Therefore, the effect of each 1mm deviation of the entire star's center of mass on the semi-major axis is shown in Table 1 below.

[0115] Table 1

[0116]

[0117]

[0118] In some implementations, the target offset direction and target offset amount of the entire star's center of mass are determined in the following manner:

[0119] When the center of mass of the satellite is offset in the +X and -X directions respectively, the direction of satellite drift caused by the tangential thrust generated by the coplanar electric thrusters is determined. The direction of the offset of the center of mass of the satellite corresponding to the direction opposite to the direction of satellite drift in longitude caused by perturbation acceleration is determined as the target offset direction of the center of mass of the satellite.

[0120] The change in the semi-major axis of the satellite during the position-holding control period under the influence of Earth's perturbation is divided by the change in the semi-major axis of the satellite during the position-holding control period caused by the tangential thrust generated by the coplanar electric thrusters when the satellite's center of mass deviates from the target offset direction by 1 mm, thus yielding the target offset of the satellite's center of mass.

[0121] Continuing with the example of a fixed position at 134 degrees east longitude, step 100 has determined that when the satellite's fixed position is at 134 degrees east longitude, the perturbation acceleration causes the satellite's mean longitude to drift westward. Under the influence of the Earth's perturbation force, the satellite's semi-major axis rises by approximately 127 meters during the position maintenance control period.

[0122] When the center of mass of the entire satellite shifts in the +X direction, the tangential thrust generated by the coplanar electric thrusters causes the satellite to accelerate westward. When the center of mass of the entire satellite shifts in the -X direction, the tangential thrust generated by the coplanar electric thrusters causes the satellite to accelerate eastward. The perturbation acceleration causes the satellite's mean longitude to drift westward. Therefore, the target shift direction of the center of mass of the entire satellite is towards the -X direction.

[0123] As can be seen from Table 1 above, the primary electric thruster is usually used. As can be seen from the table above, the change in the semi-major axis of the satellite during the position-keeping control period of 127m under the action of Earth perturbation is divided by the change in the semi-major axis of 11.55m during the position-keeping control period corresponding to the primary electric thruster in Table 1, and the target offset of the entire satellite's center of mass is obtained as 11mm.

[0124] Therefore, when the satellite is on Earth, the satellite's center of mass is adjusted to shift 11mm in the -X direction. When in orbit, during the autonomous north-south position maintenance period using steps 100-106, the effect of electric thruster ignition is opposite to the direction of Earth's natural perturbation, which can exactly cancel out the change in the satellite's semi-major axis during the position maintenance control period under the influence of Earth's perturbation. Thus, the east-west position maintenance control task is completed at the same time as the north-south position maintenance control. Not only is there no need to waste extra time on east-west position maintenance, but the tangential thrust generated during the north-south position maintenance control can also be used to adjust the east-west position, which can also reduce the consumption of chemical propellant.

[0125] like Figure 3 , Figure 4 As shown, this embodiment of the invention provides an autonomous position-holding control device for a coplanarly mounted electric thruster. The device embodiment can be implemented through software, hardware, or a combination of both. From a hardware perspective, as... Figure 3 The diagram shown is a hardware architecture diagram of a computing device housing an autonomous position-holding control device for a coplanarly mounted electric thruster, according to an embodiment of the present invention. (Except for...) Figure 3 In addition to the processor, memory, network interface, and non-volatile memory shown, the computing device in the embodiment may also include other hardware, such as a forwarding chip responsible for processing packets. Taking software implementation as an example, such as... Figure 4 As shown, a device in a logical sense is formed by the CPU of its computing device reading the corresponding computer program from the non-volatile memory into memory and running it. This embodiment provides an autonomous position-holding control device for a coplanarly mounted electric thruster based on any embodiment of this specification, comprising:

[0126] The first determining unit 401 is used to determine the change in orbital inclination vector within the position holding control period based on the current satellite time and a pre-determined position holding control period when an autonomous position holding control command is received.

[0127] The second determining unit 402 is used to determine the total ignition duration within the position holding control cycle based on the change in the track inclination vector within the position holding control cycle, so as to determine the ignition duration corresponding to each operating day within the position holding control cycle; wherein, the position holding control cycle is 7 days and the operating day is 5 days;

[0128] The third determining unit 403 is used to determine the right ascension of the ignition point based on the change in the orbital inclination vector during the position holding control cycle.

[0129] The control unit 404 is used to autonomously maintain the north-south position of the electric thruster within each operating day based on the ignition duration and right ascension of the ignition point corresponding to each operating day.

[0130] In one embodiment of the present invention, when the first determining unit 401 determines the change in orbital inclination vector within the position holding control period based on the current satellite time and a pre-determined position holding control period, it is configured to perform the following:

[0131] Based on the current star time, the initial star time, and the J2000 time, calculate the Julian century to determine the ascending node longitude;

[0132] Calculate the rate of change of orbital inclination vector over one year based on the ecliptic longitude of the ascending node;

[0133] Transform the zero tilt vector in the instantaneous true coordinate system to the J2000 coordinate system to obtain the magnitude of the orbit tilt vector in the x and y directions corresponding to the zero tilt vector in the instantaneous true coordinate system in the J2000 coordinate system.

[0134] Based on the rate of change of the track inclination vector within one year, the magnitude of the track inclination vector in the x and y directions corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, and a predetermined position holding control period, the change in the track inclination vector within the position holding control period is determined.

[0135] In one embodiment of the present invention, the rate of change of the orbital inclination vector over one year in the first determining unit 401 is determined in the following manner:

[0136] The Julian century can be calculated using the following formula:

[0137]

[0138] in,

[0139] t J2000 =t+473342400

[0140] In the formula, T represents the Julian century, and t J2000 t represents the number of seconds between the current satellite time and the J2000 time, and t represents the number of seconds between the current satellite time and the initial satellite time.

[0141] Calculate the ecliptic longitude of the ascending node using the following formula:

[0142] Ω l =2.182439 - 33.757 × T

[0143] In the formula, T represents the Julian century, and Ω l The ecliptic longitude of the ascending node;

[0144] The rate of change of the orbital inclination vector over a year can be calculated using the following formula:

[0145] dΔi x =0.097cosΩ l +0.852

[0146] dΔi y =0.13sinΩ l

[0147] In the formula, dΔi x ′ and dΔi y ′ represent the rates of change of the orbital inclination vector in the x-direction and y-direction, respectively, over a year, Ω l The ecliptic longitude of the ascending node;

[0148] In the first determining unit 401, the zero tilt vector of the instantaneous true coordinate system is transformed to the J2000 coordinate system using the following formula:

[0149] Δi xTOD =KTodcof1*C PR (2,0)

[0150] Δi yTOD =KTodcof1*C PR (2,1)

[0151] In the formula, Δi xTOD and Δi yTOD These represent the x- and y-direction magnitudes of the orbital inclination vector corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, respectively. KTodcof1 is a pre-set coordinate system transformation adjustment coefficient, and C... PR (2,0) is the element in the 3rd row and 1st column of the precession nutation transformation matrix, where 2 and 0 are the row and column numbers, respectively. PR (2,1) is the element in the 3rd row and 2nd column of the precession nutation transformation matrix, where 2 and 1 are the row number and column number, respectively.

[0152] In one embodiment of the present invention, the change in the track inclination vector during the position holding control cycle in the first determining unit 401 is calculated using the following formula:

[0153] Δi x =dΔi′ x / 365*TJetPeriod*Kcomp+(i·sin(Ω)-Δi xTOD R2D

[0154] Δiy =dΔi′ y / 365*TJetPeriod*Kcomp+(-i·cos(Ω)-Δi yTOD R2D

[0155] In the formula, Δi x and Δi y These represent the changes in the orbital inclination vector in the x-direction and y-direction, respectively, during the position-holding control cycle, dΔi′ x and dΔi′ y These represent the rates of change of the orbital inclination vector in the x and y directions over one year, respectively. TJetPeriod is the pre-determined position-keeping control period, Kcomp is the correction value (default is 1), i is the satellite's orbital inclination, Ω is the right ascension of the satellite's ascending node, and Δi... xTOD and Δi yTOD R2D represents the magnitudes of the orbital inclination vector in the x and y directions corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, respectively, and R2D is the conversion between radians and degrees.

[0156] In one embodiment of the present invention, the second determining unit 402 is used to perform:

[0157] Based on the change in the track inclination vector during the position holding control cycle, the magnitude of the track inclination change is calculated, and the required normal velocity increment is determined based on the magnitude of the track inclination change.

[0158] Based on the required velocity increment in the normal direction, the satellite's mass, and the satellite's nominal thrust, determine the total ignition duration within the position-holding control cycle;

[0159] Divide the total ignition time within the position holding control cycle into 5 equal parts to obtain the ignition time corresponding to each operating day.

[0160] In one embodiment of the present invention, after the second determining unit 402 divides the total ignition duration within the position holding control cycle into five equal parts, and before obtaining the ignition duration corresponding to each operating day, the second determining unit 402 further includes:

[0161] Obtain the time limit for coplanar mounted electric thrusters;

[0162] Determine whether the total ignition time, after being divided equally, exceeds the time limit.

[0163] If the limit is exceeded, the time limit will be set as the ignition time corresponding to each operation day.

[0164] If the total ignition time is not exceeded, the total ignition time will be divided equally and used as the ignition time for each operation day.

[0165] In one embodiment of the present invention, the right ascension of the ignition point in the third determining unit 403 is calculated using the following formula:

[0166] OmgF=atan2(Δi x ,Δi y )

[0167] In the formula, OmgF is the right ascension of the ignition point, and Δi x and Δi y These represent the changes in the track inclination vector in the x-direction and the changes in the y-direction, respectively, during the position holding control cycle.

[0168] It is understood that the structures illustrated in the embodiments of the present invention do not constitute a specific limitation on an autonomous position-holding control device for a coplanarly mounted electric thruster. In other embodiments of the present invention, an autonomous position-holding control device for a coplanarly mounted electric thruster may include more or fewer component units than illustrated, or combine certain component units, or split certain component units, or arrange different component units. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0169] The information interaction and execution process between the modules in the above-mentioned device are based on the same concept as the method embodiment of the present invention, and the specific details can be found in the description of the method embodiment of the present invention, and will not be repeated here.

[0170] This invention also provides a computing device, including a memory and a processor. The memory stores a computer program, and when the processor executes the computer program, it implements an autonomous position-keeping control method for a coplanarly mounted electric thruster according to any embodiment of this invention.

[0171] This invention also provides a computer-readable storage medium storing a computer program, which, when executed by a processor, causes the processor to perform an autonomous position-holding control method for a coplanarly mounted electric thruster according to any embodiment of this invention.

[0172] Specifically, a system or apparatus equipped with a storage medium may be provided, on which software program code implementing the functions of any of the embodiments described above is stored, and the computer (or CPU or MPU) of the system or apparatus may read and execute the program code stored in the storage medium.

[0173] In this case, the program code read from the storage medium can itself implement the function of any of the above embodiments, and therefore the program code and the storage medium storing the program code constitute part of the present invention.

[0174] Examples of storage media used to provide program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Alternatively, program code can be downloaded from a server computer via a communication network.

[0175] Furthermore, it should be clear that not only can the program code read by the computer be executed, but also the operating system or other components operating on the computer can be instructed based on the program code to perform some or all of the actual operations, thereby realizing the function of any of the embodiments described above.

[0176] Furthermore, it is understood that the program code read from the storage medium is written to the memory set in the expansion board inserted into the computer or to the memory set in the expansion module connected to the computer. Then, based on the instructions of the program code, the CPU or other components installed on the expansion board or expansion module execute some and all of the actual operations, thereby realizing the function of any of the embodiments described above.

[0177] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus.

[0178] Those skilled in the art will understand that all or part of the steps of the above method embodiments can be implemented by hardware related to program instructions. The aforementioned program can be stored in a computer-readable storage medium. When the program is executed, it performs the steps of the above method embodiments. The aforementioned storage medium includes various media that can store program code, such as ROM, RAM, magnetic disk, or optical disk.

[0179] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. An autonomous position holding control method of a coplanar installation electric thruster, characterized by, include: The perturbation acceleration caused by the satellite's fixed position determines the direction of the satellite's mean longitude drift and the change in the satellite's semi-major axis during the position-holding control period under the influence of Earth's perturbation force; wherein, the position-holding control period is 7 days and the operation day is 5 days; Determine the direction of satellite drift caused by the tangential thrust generated by the coplanar electric thrusters when the center of mass of the entire satellite deviates by 1 mm in the +X or -X direction, as well as the amount of change of the semi-major axis of the satellite during the position-holding control period. Based on the direction of satellite drift in longitude caused by perturbation acceleration and the amount of change of the semi-major axis of the satellite during the position-holding control period under the action of Earth perturbation, determine the target offset direction and target offset of the center of mass of the entire satellite. Adjust the position of the satellite's center of mass based on the target offset direction and target offset amount; When an autonomous position-keeping control command is received, the orbital inclination vector change within the position-keeping control period is determined based on the current satellite time and a pre-determined position-keeping control period. Based on the change in the orbital inclination vector within the position holding control cycle, the total ignition duration within the position holding control cycle is determined, so as to determine the ignition duration corresponding to each operating day within the position holding control cycle. The right ascension of the ignition point is determined based on the change in the orbital inclination vector during the position holding control cycle. Based on the ignition duration and right ascension of the ignition point for each operation day, the coplanar electric thrusters are used to autonomously maintain the north-south position within each operation day. The change in the semi-major axis of the satellite during the position-holding control period caused by the tangential thrust generated by the coplanar electric thrusters when the center of mass of the entire satellite deviates by 1 mm is determined as follows: The tangential and normal thrusts of the primary and backup electric thrusters, which are coplanarly mounted, are obtained when the center of mass of the entire satellite deviates by 1 mm. Based on tangential thrust, normal thrust, and a pre-set normal daily average velocity increment, calculate the corresponding tangential daily average velocity increment when the primary electric thruster is started and when the backup electric thruster is started. Based on the vitality formula and the daily average tangential velocity increment, the daily average semi-major axis change is calculated when the primary electric thruster is started and when the backup electric thruster is started. Based on the daily average semi-major axis change and the number of days in the position holding control period, the tangential thrust generated by the coplanar electric thrusters caused by the 1mm deviation of the satellite's center of mass is determined as the change in the satellite's semi-major axis during the position holding control period.

2. The method according to claim 1, characterized in that, The determination of the orbital inclination vector change within the position-holding control period based on the current satellite time and a pre-determined position-holding control period includes: Based on the current star time, the initial star time, and the J2000 time, calculate the Julian century, and determine the ascending node ecliptic longitude based on the Julian century; Based on the ecliptic longitude of the ascending node, calculate the rate of change of the orbital inclination vector over one year; Transform the zero tilt vector in the instantaneous true coordinate system to the J2000 coordinate system to obtain the magnitude of the orbit tilt vector in the x and y directions corresponding to the zero tilt vector in the instantaneous true coordinate system in the J2000 coordinate system. Based on the rate of change of the track inclination vector within one year, the magnitude of the track inclination vector in the x and y directions corresponding to the zero inclination vector in the instantaneous true coordinate system in the J2000 coordinate system, and a predetermined position holding control period, the change in the track inclination vector within the position holding control period is determined.

3. The method according to claim 2, characterized in that, The rate of change of the orbital inclination vector over one year is determined as follows: The Julian century can be calculated using the following formula: in, In the formula, For the Julian century, This represents the number of seconds between the current satellite time and the J2000 time. This represents the number of seconds between the current star time and the initial star time. Calculate the ecliptic longitude of the ascending node using the following formula: In the formula, For the Julian century, The ecliptic longitude of the ascending node; The rate of change of the orbital inclination vector over a year can be calculated using the following formula: In the formula, and These represent the rates of change of the orbital inclination vector in the x-direction and the rate of change in the y-direction over a year, respectively. The ecliptic longitude of the ascending node; The zero tilt vector of the instantaneous true coordinate system is transformed to the J2000 coordinate system using the following formula: In the formula, and These represent the magnitudes of the orbital inclination vector in the x and y directions, respectively, of the zero inclination vector in the instantaneous true coordinate system and the corresponding orbital inclination vector in the J2000 coordinate system. For the pre-set coordinate system transformation adjustment coefficient, This is the element in the 3rd row and 1st column of the precession nutation transformation matrix, where 2 and 0 are the row and column numbers, respectively. This is the element in the 3rd row and 2nd column of the precession nutation transformation matrix, where 2 and 1 are the row number and column number, respectively.

4. The method according to claim 2, characterized in that, The change in track inclination vector during the position holding control cycle is calculated using the following formula: In the formula, and These represent the changes in the track inclination vector in the x-direction and the y-direction, respectively, during the position holding control cycle. and These represent the rates of change of the orbital inclination vector in the x-direction and the rate of change in the y-direction over a year, respectively. To maintain a control cycle at a predetermined location, This is a correction value; the default value is 1. The orbital inclination of the satellite, The right ascension of the ascending node of the satellite, and These represent the magnitudes of the orbital inclination vector in the x and y directions, respectively, of the zero inclination vector in the instantaneous true coordinate system and the corresponding orbital inclination vector in the J2000 coordinate system. This is for converting between radians and degrees.

5. The method according to claim 1, characterized in that, The determination of the total ignition duration within the position holding control cycle based on the change in the orbital inclination vector during the position holding control cycle, and the determination of the ignition duration corresponding to each operating day within the position holding control cycle, includes: Based on the change in the track inclination vector during the position holding control cycle, the magnitude of the track inclination change is calculated, and the required normal velocity increment is determined based on the magnitude of the track inclination change. Based on the required velocity increment in the normal direction, the satellite's mass, and the satellite's nominal thrust, the total ignition duration within the position holding control cycle is determined. Divide the total ignition duration within the position holding control cycle into 5 equal parts to obtain the ignition duration corresponding to each operating day.

6. The method according to claim 5, characterized in that, After dividing the total ignition duration within the position holding control cycle into 5 equal parts, and before obtaining the ignition duration corresponding to each operating day, the method further includes: The time limit for obtaining the coplanar mounted electric thruster; Determine whether the total ignition time, after being divided equally, exceeds the stated time limit. If the limit is exceeded, the time limit will be determined as the ignition time corresponding to each operation day. If the total ignition time is not exceeded, the total ignition time will be divided equally and used as the ignition time for each operation day.

7. The method according to claim 6, characterized in that, The right ascension of the ignition point is calculated using the following formula: In the formula, To ignite the red ascension, and These represent the changes in the track inclination vector in the x-direction and the changes in the y-direction, respectively, during the position holding control cycle.

8. An autonomous position-keeping control device based on any one of the methods described in 1-7, characterized in that, include: The first determining unit is used to determine the change in orbital inclination vector within the position holding control period based on the current satellite time and a pre-determined position holding control period when an autonomous position holding control command is received. The second determining unit is used to determine the total ignition duration within the position holding control cycle based on the change in the orbital inclination vector within the position holding control cycle, so as to determine the ignition duration corresponding to each operating day within the position holding control cycle; wherein, the position holding control cycle is 7 days and the operating day is 5 days; The third determining unit is used to determine the right ascension of the ignition point based on the change in the orbital inclination vector during the position holding control cycle; The control unit is used to autonomously maintain the north-south position of the electric thruster within each operating day, based on the ignition duration and the right ascension of the ignition point corresponding to each operating day.

9. A computing device comprising a memory and a processor, wherein the memory stores a computer program, and the processor, when executing the computer program, implements the method as described in any one of claims 1-7.

10. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the method of any one of claims 1-7.

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