Satellite-borne high-precision orbit extrapolation method and device and electronic equipment

Abstract

The invention provides a satellite-borne high-precision orbit extrapolation method, a satellite-borne high-precision orbit extrapolation device and electronic equipment. The satellite-borne high-precision orbit extrapolation method comprises the following steps: providing an orbit extrapolation input quantity of a satellite at an initial extrapolation moment; calculating an extrapolation duration based on the initial extrapolation moment and a preset threshold value, and determining a termination extrapolation moment of the satellite based on the extrapolation duration; calculating the non-spherical gravitational perturbation acceleration of the earth according to the non-spherical gravitational perturbation model of the earth; calculating a day-moon three-body gravitational acceleration according to a day-moon three-body gravitational model; determining a satellite thrust acceleration according to the current working state of the satellite; adding the earth non-spherical gravitational perturbation acceleration, the sun-moon three-body gravitational acceleration and the satellite thrust acceleration to obtain an extrapolation resultant acceleration; performing orbit extrapolation calculation on the basis of the extrapolation resultant acceleration by adopting a fourth-order Runge-Kutta method to obtain a target extrapolation position and a target extrapolation speed at an extrapolation termination moment; and according to the target extrapolation position and the target extrapolation speed, updating the orbit extrapolation input quantity at the next extrapolation moment.

Description

TECHNICAL FIELD

[0001] The present application mainly relates to the field of spacecraft orbit extrapolation technology, and particularly relates to a high-precision orbit extrapolation method and device implemented on a satellite and an electronic device. BACKGROUND

[0002] Orbit extrapolation constructs a satellite orbit dynamics model by analyzing the force acting on the satellite, and then obtains orbit prediction information of the satellite based on the satellite orbit dynamics model. However, the commonly used satellite orbit dynamics model only considers the action of two-body gravity and J2 perturbation force, and the orbit prediction accuracy is not high, especially for geolunar space satellites and deep space satellites beyond the moon, it is difficult to meet the autonomous orbit prediction requirements. SUMMARY

[0003] The purpose of the present application is to provide a high-precision orbit extrapolation method and device implemented on a satellite and an electronic device, to solve the problem of low orbit prediction accuracy of existing geolunar space satellites and deep space satellites beyond the moon.

[0004] In a first aspect, the present application provides a high-precision orbit extrapolation method implemented on a satellite, comprising:

[0005] providing an orbit extrapolation input quantity of a satellite at an initial extrapolation time, wherein the orbit extrapolation input quantity comprises an initial extrapolation position and an initial extrapolation velocity;

[0006] calculating an extrapolation duration based on the initial extrapolation time and a preset threshold, and determining a terminal extrapolation time of the satellite based on the extrapolation duration;

[0007] calculating an earth non-spherical gravity perturbation acceleration according to an earth non-spherical gravity perturbation model;

[0008] calculating a sun-moon three-body gravity acceleration according to a sun-moon three-body gravity model;

[0009] determining a satellite thrust acceleration according to a current working state of the satellite;

[0010] adding the earth non-spherical gravity perturbation acceleration, the sun-moon three-body gravity acceleration and the satellite thrust acceleration to obtain an extrapolation combined acceleration;

[0011] performing orbit extrapolation calculation based on the extrapolation combined acceleration by using a fourth-order Runge-Kutta method to obtain a target extrapolation position and a target extrapolation velocity at the terminal extrapolation time;

[0012] updating the orbit extrapolation input quantity at a next extrapolation time according to the target extrapolation position and the target extrapolation velocity.

[0013] In some embodiments, the calculation of the extrapolation duration based on the initial extrapolation time and the preset threshold comprises:

[0014] reading the extrapolation start point valid flag, if the extrapolation start point valid flag is valid, judging whether a time difference between a current time and the initial extrapolation time is greater than or equal to the preset threshold value;

[0015] if the time difference is greater than or equal to the preset threshold value, taking the preset threshold value as the extrapolation duration, otherwise, taking the time difference as the extrapolation duration.

[0016] In some embodiments, the calculating the earth non-spherical gravity perturbation acceleration according to the earth non-spherical gravity perturbation model comprises:

[0017] obtaining an inertial frame to earth fixed frame transfer matrix;

[0018] calculating the earth fixed frame position of the satellite according to the initial extrapolation position and the inertial frame to earth fixed frame transfer matrix;

[0019] calculating the non-spherical gravity potential coefficient corresponding to the earth fixed frame position according to the earth non-spherical gravity perturbation model;

[0020] calculating the non-spherical gravity acceleration of the satellite in the earth fixed frame using the non-spherical gravity potential coefficient; and,

[0021] obtaining the earth non-spherical gravity perturbation acceleration according to the non-spherical gravity acceleration of the satellite in the earth fixed frame and the transpose matrix of the inertial frame to earth fixed frame transfer matrix.

[0022] In some embodiments, the calculating the sun-earth-moon three-body gravity acceleration according to the sun-earth-moon three-body gravity model comprises:

[0023] obtaining a current Chebyshev standard time, a geolunar system Chebyshev coefficient, a solar system Chebyshev coefficient and a moon Chebyshev coefficient;

[0024] calculating a geolunar system center of mass position, a sun center of mass position and a moon center of mass position according to the current Chebyshev standard time, the geolunar system Chebyshev coefficient, the solar system Chebyshev coefficient and the moon Chebyshev coefficient;

[0025] calculating the sun gravity acceleration and the moon gravity acceleration corresponding to the initial extrapolation position according to the sun-earth-moon three-body gravity model;

[0026] adding the sun gravity acceleration and the moon gravity acceleration to obtain the sun-earth-moon three-body gravity acceleration.

[0027] In some embodiments, the determining the satellite thrust acceleration according to the current working state of the satellite comprises:

[0028] if the current working state of the satellite is the orbit control mode, the satellite thrust acceleration is measured by an on-board accelerometer;

[0029] if the current working state of the satellite is the non-orbit control mode, the satellite thrust acceleration is set to zero.

[0030] In some embodiments, the orbit extrapolation calculation based on the extrapolated combined acceleration is performed by using the fourth-order Runge-Kutta method, to obtain the target extrapolation position and the target extrapolation velocity at the termination extrapolation time, including:

[0031] Let Vel in be the initial extrapolation velocity, Pos in be the initial extrapolation position, T S be the extrapolation time length, then the formula of the orbit extrapolation calculation is:

[0032] ,

[0033] ,

[0034] wherein, ,

[0035] wherein Pos out is the target extrapolation position, and Vel out is the target extrapolation velocity.

[0036] In a second aspect, the application provides a high-precision orbit extrapolation device implemented on a satellite, including:

[0037] an orbit extrapolation time length calculation module, configured to provide orbit extrapolation input quantities of the satellite at an initial extrapolation time, wherein the orbit extrapolation input quantities include an initial extrapolation position and an initial extrapolation velocity; calculate an extrapolation time length based on the initial extrapolation time and a preset threshold; and determine a termination extrapolation time of the satellite based on the extrapolation time length;

[0038] an orbit perturbation force calculation module, configured to calculate an earth non-spherical gravity perturbation acceleration according to an earth non-spherical gravity perturbation model; calculate a sun-moon three-body gravity acceleration according to a sun-moon three-body gravity model; determine a satellite thrust acceleration according to a current working state of the satellite; and add the earth non-spherical gravity perturbation acceleration, the sun-moon three-body gravity acceleration and the satellite thrust acceleration to obtain an extrapolation combined acceleration;

[0039] a position and velocity information solving module, configured to perform orbit extrapolation calculation based on the extrapolation combined acceleration by using the fourth-order Runge-Kutta method, to obtain a target extrapolation position and a target extrapolation velocity at the termination extrapolation time;

[0040] A position and velocity information updating module is configured to update orbit extrapolation input at a next extrapolation time according to the target extrapolated position and the target extrapolated velocity.

[0041] In a third aspect, an electronic device is provided. The electronic device includes one or more processors and one or more memories coupled to the one or more processors and storing instructions thereon. When the instructions are executed by the one or more processors alone or in combination, the electronic device performs the method described above.

[0042] In a fourth aspect, a non-transitory computer-readable storage medium storing machine executable instructions is provided. The machine executable instructions, when executed by one or more processors of a machine, cause the machine to perform any of the methods described above.

[0043] In a fifth aspect, a computer program product including machine executable instructions is provided. The machine executable instructions, when executed by one or more processors of a machine, cause the machine to perform any of the methods described above.

[0044] The beneficial effects of the present application are as follows:

[0045] 1. The present application realizes high-precision orbit prediction of geolunar space satellites and deep space satellites beyond the moon by comprehensively considering the influence of the non-spherical gravity perturbation of the earth, the three-body gravity perturbation of the sun and the moon, and the satellite thrust. The process can be realized autonomously on the satellite, thereby improving the autonomous orbit extrapolation performance of geolunar space satellites and deep space satellites beyond the moon.

[0046] 2. The present application combines the fourth-order Runge-Kutta method and extrapolation with acceleration to perform orbit extrapolation calculation. This method has low complexity and is helpful for stable application of geolunar space satellites and deep space satellites beyond the moon.

[0047] It should be understood that the summary section is not intended to identify key or essential features of embodiments of the disclosure or to limit the scope of the disclosure. Other features of the disclosure will be readily apparent from the following description. BRIEF DESCRIPTION OF DRAWINGS

[0048] The above and other objects, features and advantages of the disclosure will become more apparent from the following description when taken in conjunction with the accompanying drawings, in which:

[0049] FIG. 1 shows a flowchart of a high-precision orbit extrapolation method implemented on a satellite;

[0050] FIG. 2 shows a schematic diagram of a high-precision orbit extrapolation device implemented on a satellite;

[0051] FIG. 3 shows a schematic diagram of an exemplary electronic device. DETAILED DESCRIPTION

[0052] The principles of the present disclosure will now be described with reference to some embodiments. It is to be understood that the description of these embodiments is merely intended to illustrate the present disclosure and to help the skilled person understand and implement the present disclosure, and does not in any way pose any limitation on the scope of the present disclosure. The disclosure described herein can be implemented in a manner different from that described below.

[0053] In the following description and claims, unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this disclosure belongs.

[0054] Reference in the present disclosure to "one embodiment", "an embodiment", "exemplary embodiment", etc., indicates that a described embodiment can include a particular feature, structure, or characteristic, but every embodiment can not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Furthermore, when a particular feature, structure, or characteristic is described in connection with an example embodiment, it is submitted that it is within the knowledge of those skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described.

[0055] It should be understood that although the terms "first" and "second" and the like can be used herein to describe various elements, these elements should not be limited by these terms. These terms are only used to distinguish one element from another. For example, a first element could be termed a second element, and, similarly, a second element could be termed a first element, without departing from the scope of example embodiments. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed terms.

[0056] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of example embodiments. As used herein, the singular forms "a", "an" and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise. The terms "comprises", "comprising", "including", "having", "containing", "involving", and / or "including" as used herein, specifies the presence of stated features, elements, and / or components, but do not preclude the presence or addition of one or more other features, elements, components, and / or groups thereof.

[0057] FIG. 1 shows a flowchart of a high-precision orbit extrapolation method implemented on a satellite according to the present application. Referring to FIG. 1, the method comprises:

[0058] S101, providing an orbit extrapolation input quantity of the satellite at an initial extrapolation time.

[0059] The orbit extrapolation input quantity includes an initial extrapolation position and an initial extrapolation velocity.

[0060] S102, calculate the extrapolation duration based on the initial extrapolation time and the preset threshold, and determine the terminal extrapolation time of the satellite based on the extrapolation duration.

[0061] In some embodiments, calculating the extrapolation duration based on the initial extrapolation time and the preset threshold comprises:

[0062] reading the extrapolation start point valid flag, and if the extrapolation start point valid flag is valid, determining whether the time difference between the current time and the initial extrapolation time is greater than or equal to the preset threshold;

[0063] If the time difference is greater than or equal to the preset threshold, the preset threshold is taken as the extrapolation duration, otherwise the time difference is taken as the extrapolation duration.

[0064] It will be understood that if the extrapolation start point valid flag is invalid, the satellite can directly exit the extrapolation calculation and the current step is terminated.

[0065] Let the current time be t k , the initial extrapolation time be t0, the preset threshold be D tmax , the extrapolation duration be T S , and the interpolation be t k1 . When t k -t0≥D tmax , then t k1 =D tmax +t0, set the current extrapolation valid flag to invalid, otherwise t k1 =t k , set the current extrapolation valid flag to valid, and the expression of the extrapolation duration T S is shown in the following formula (1):

[0066] T S =t k1 -t0 (1)

[0067] S103, calculate the earth non-spherical gravitational perturbation acceleration according to the earth non-spherical gravitational perturbation model.

[0068] In some embodiments, calculating the earth non-spherical gravitational perturbation acceleration according to the earth non-spherical gravitational perturbation model comprises the following steps:

[0069] Step one, obtain the inertial system to the earth fixed system transfer matrix.

[0070] Let the inertial system to the earth fixed system transfer matrix be E wi , and its expression is shown in the following formula (2):

[0071] (2)

[0072] wherein, is the rotation matrix of true sidereal time Greenwich hour angle, is the nutation matrix, is the precession matrix.

[0073] Step two, calculate the satellite's geodetic position according to the initial extrapolation position and the inertial system to the geodetic system transfer matrix.

[0074] Let the initial extrapolation position be Pos in , and the satellite's geodetic position be Pos w , whose expression is shown in the following formula (3):

[0075] Pos w = E wi Pos in (3)

[0076] Step three, calculate the geodetic system position corresponding to the non-spherical gravity potential coefficient according to the earth's non-spherical gravity perturbation model.

[0077] The expression of the earth's non-spherical gravity perturbation model is shown in the following formula (4):

[0078]

[0079] (4)

[0080] Where, the initial value and , m and l are the order and degree respectively, , is the non-spherical gravity potential coefficient.

[0081] Step four, calculate the satellite's non-spherical gravity acceleration in the geodetic system using the non-spherical gravity potential coefficient.

[0082] Let the satellite's non-spherical gravity acceleration in the geodetic system be , which is shown in the following formula (5):

[0083] (5)

[0084] Where, , is the non-spherical gravity coefficient of 4 orders.

[0085] Step five, obtain the earth's non-spherical gravity perturbation acceleration according to the satellite's non-spherical gravity acceleration in the geodetic system and the transpose matrix of the inertial system to the geodetic system transfer matrix.

[0086] Let the earth's non-spherical gravity perturbation acceleration be , whose expression is shown in the following formula (6):

[0087] (6)

[0088] S104, calculating the Sun-Earth-Moon gravitational acceleration according to the Sun-Earth-Moon gravitational model.

[0089] In some embodiments, the step of calculating the Sun-Earth-Moon gravitational acceleration according to the Sun-Earth-Moon gravitational model comprises the following steps:

[0090] Step one, obtaining the current Chebyshev standard time, the Earth-Moon system Chebyshev coefficient, the solar system Chebyshev coefficient and the Moon system Chebyshev coefficient.

[0091] Let the current Chebyshev standard time be T au , and its expression under the Earth-Moon system, the solar system and the Moon system is shown in the three formulas from top to bottom in equation (7) as follows:

[0092] T au = 2*(2*(T k -T_ staEarth ) / (T_ endEarth -T_ staEarth )-1)

[0093] T au = 2*(2*(T k -T_ staSun ) / (T_ endSun -T_ staSun )-1)

[0094] T au = 2*(2*(T k -T_ staMoon ) / (T_ endMoon -T_ staMoon )-1)(7)

[0095] Wherein, T staEarth , T endEarth are the Earth-Moon system Julian day start and end time, T staSun , T endSun are the solar system Julian day start and end time, T staMoon , T endMoon are the Moon system Julian day start and end time.

[0096] Let the Earth-Moon system Chebyshev coefficient be Pc_ Earth [], and its expression is shown in equation (8) as follows:

[0097] Pc_ Earth [0] = 1;

[0098] Pc_ Earth [1]= T au ;

[0099] Pc_ Earth [i]= T au *Pc_ Earth [i-1] - Pc_ Earth [i-2];i = 2....8(8)

[0100] Let the Chebyshev coefficient of the solar system be Pc_ Sun [], whose expression is shown in the following formula (9) :

[0101] Pc_ Sun [0] = 1;

[0102] Pc_ Sun [1]= T au ;

[0103] Pc_ Sun [i]= T au *Pc_ Sun [i-1] - Pc_ Sun [i-2];i = 2....8(9)

[0104] Let the Chebyshev coefficient of the lunar system be Pc_ Moon [], whose expression is shown in the following formula (10) :

[0105] Pc_ Moon [0] = 1;

[0106] Pc_ Moon [1]= T au ;

[0107] Pc_ Moon [i]= T au *Pc_ Moon [i-1] - Pc_ Moon [i-2];i = 2....8(10)

[0108] Step two, according to the current Chebyshev standard time, the Chebyshev coefficient of the Earth-moon system, the Chebyshev coefficient of the solar system and the Chebyshev coefficient of the moon, calculate the Earth-moon system barycenter position, the solar barycenter position and the moon barycenter position.

[0109] Let the Earth-moon system barycenter position be EMBP, whose expression is shown in the following formula (11) :

[0110] EMBP= Ecof_ earth *Pc_ Earth ’*1000(11)

[0111] Wherein, Ecof_ earthis the sun coefficient.

[0112] Let the sun center position be SunP, which is expressed as shown in equation (12) below:

[0113] SunP= Ecof_ Sun *Pc_ Sun ’*1000 (12)

[0114] wherein Ecof_ Sun is the sun coefficient.

[0115] Let the moon center position be MoonP, which is expressed as shown in equation (13) below:

[0116] MoonP= Ecof_ Moon *Pc_ Moon ’*1000 (13)

[0117] wherein Ecof_ Moon is the moon coefficient.

[0118] Step three, calculate the sun gravitational acceleration and the moon gravitational acceleration corresponding to the initial extrapolation position according to the sun-moon-earth three-body gravitational model.

[0119] Let the sun gravitational acceleration be , which is expressed as shown in equation (14) below:

[0120] (14)

[0121] wherein P sunI is the sun center position.

[0122] Let the moon gravitational acceleration be , which is expressed as shown in equation (15) below:

[0123] (15)

[0124] wherein P moonI is the moon center position.

[0125] Step four, add the sun gravitational acceleration and the moon gravitational acceleration to obtain the sun-moon-earth three-body gravitational acceleration.

[0126] Let the sun-moon-earth three-body gravitational acceleration be , which is expressed as shown in equation (16) below:

[0127] (16)

[0128] S105, determine the satellite thrust acceleration according to the current working state of the satellite.

[0129] In some embodiments, the satellite thrust acceleration is determined according to the current working state of the satellite, comprising:

[0130] If the current working state of the satellite is the orbit control mode, the satellite thrust acceleration is measured by the on-board accelerometer;

[0131] If the current working state of the satellite is the non-orbit control mode, the satellite thrust acceleration is set to zero.

[0132] Let the satellite thrust acceleration be If the current working state is the non-orbit control mode, then = 0; otherwise, is the accelerometer calibration reading.

[0133] S106, the earth non-spherical gravity perturbation acceleration, the sun and moon three-body gravitational acceleration and the satellite thrust acceleration are added to obtain the extrapolation combined acceleration.

[0134] Let the extrapolation combined acceleration be Its expression is shown in the following formula (17):

[0135] (17)

[0136] S107, the fourth-order Runge-Kutta method is used to perform orbit extrapolation calculation based on the extrapolation combined acceleration to obtain the target extrapolation position and the target extrapolation velocity at the termination time of extrapolation.

[0137] In some embodiments, the fourth-order Runge-Kutta method is used to perform orbit extrapolation calculation based on the extrapolation combined acceleration to obtain the target extrapolation position and the target extrapolation velocity at the termination time of extrapolation, comprising:

[0138] Let Vel in be the initial extrapolation velocity, Pos in be the initial extrapolation position, and T S be the extrapolation time length, then the formula of the orbit extrapolation calculation is shown in the following formula (18):

[0139] ,

[0140] ,

[0141] wherein, , (18)

[0142] wherein Pos out is the target extrapolation position and Vel out is the target extrapolation velocity.

[0143] S108, according to the target extrapolation position and the target extrapolation velocity, the orbit extrapolation input quantity at the next extrapolation time is updated.

[0144] The initial extrapolation position is Posin' and the initial extrapolation velocity is Velin' in the orbit extrapolation input quantity at the next extrapolation time point, and they satisfy the following formula (19): in

[0145] Posin' = Pos out

[0146] Vel in out (19)

[0147] In summary, the target extrapolation position and the target extrapolation velocity at the termination extrapolation time point can be calculated through the above process, high-precision orbit prediction of the earth-moon space satellite and the deep space satellite beyond the moon is realized, the process can be realized on the satellite, and then the autonomous orbit extrapolation performance of the earth-moon space satellite and the deep space satellite beyond the moon is improved.

[0148] FIG. 2 shows a schematic diagram of a high-precision orbit extrapolation device provided by the application, which comprises:

[0149] An orbit extrapolation time length calculation module 201 is configured to provide orbit extrapolation input quantity of a satellite at an initial extrapolation time point, wherein the orbit extrapolation input quantity comprises an initial extrapolation position and an initial extrapolation velocity; an extrapolation time length is calculated based on the initial extrapolation time point and a preset threshold value, and a termination extrapolation time point of the satellite is determined based on the extrapolation time length;

[0150] An orbit perturbation force calculation module 202 is configured to calculate the earth non-spherical gravity perturbation acceleration according to an earth non-spherical gravity perturbation model; calculate the sun-moon three-body gravity acceleration according to a sun-moon three-body gravity model; determine the satellite thrust acceleration according to the current working state of the satellite; and add the earth non-spherical gravity perturbation acceleration, the sun-moon three-body gravity acceleration and the satellite thrust acceleration to obtain the extrapolation combined acceleration;

[0151] A position and velocity information solving module 203 is configured to perform orbit extrapolation calculation based on the extrapolation combined acceleration by using the fourth-order Runge-Kutta method to obtain the target extrapolation position and the target extrapolation velocity at the termination extrapolation time point;

[0152] A position and velocity information updating module 204 is configured to update the orbit extrapolation input quantity at the next extrapolation time point according to the target extrapolation position and the target extrapolation velocity.

[0153] ​​Further, as shown in Fig. 3, the exemplary embodiments of the present application also provide an electronic device, comprising one or more memories 301, and one or more processors 302, wherein the one or more memories 301 are coupled to the one or more processors 302 and store instructions thereon which are executable by the one or more processors 302 alone or in combination to cause the electronic device to perform the method according to any one of the first aspect.

[0154] It should be understood that the processor mentioned in the embodiments of the present application can be a CPU, and can also be other general-purpose processors, DSPs, ASICs, FPGAs or other programmable logic devices, discrete gates or transistor logic devices, discrete hardware components, etc. The general-purpose processor can be a microprocessor or the processor can also be any conventional processor.

[0155] It should also be understood that the memory mentioned in the embodiments of the present application can be a volatile memory or a non-volatile memory, or can include both volatile and non-volatile memories. Among them, the non-volatile memory can be a read-only memory (ROM), a programmable read-only memory, an erasable programmable read-only memory, an electrically erasable programmable read-only memory or a flash memory. The volatile memory can be a random access memory (RAM) used as an external cache. By way of example but not limitation, many forms of RAM are available, such as static random access memory, dynamic random access memory, synchronous dynamic random access memory, double data rate synchronous dynamic random access memory, enhanced synchronous dynamic random access memory, synchronous link dynamic random access memory and direct memory bus random access memory.

[0156] The present application also provides a non-transitory computer-readable storage medium storing machine-executable instructions, which can be executed by one or more processors of a machine. The machine can include an electronic device as mentioned above, etc. When the computer-executable instructions are executed by the one or more processors, the machine performs any one of the methods as mentioned above.

[0157] The computer-readable storage medium can contain a propagated data signal with computer program code embodied therein, for example, in baseband or as part of a carrier wave. Such propagated signals can take a wide variety of forms, including, but not limited to, electro-magnetic, optical, or any suitable combination thereof. Computer-readable storage medium can be accessed via connection to a processor, device, or device, to implement communication, propagation, or transmission of the program code. Program code embodied on the computer-readable storage medium can be transmitted using any appropriate medium, including, but not limited to, wireless, wireline, optical fiber cable, RF, or any suitable combination thereof.

[0158] This application also provides a computer program product including machine-executable instructions. When executed by one or more processors of a machine, the machine-executable instructions cause the machine to perform any of the methods described above.

[0159] The basic concepts have been described above. Obviously, for those skilled in the art, the above disclosure is merely illustrative and does not constitute a limitation of this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and corrections to this application. Such modifications, improvements, and corrections are suggested in this application, and therefore remain within the spirit and scope of the exemplary embodiments of this application.

[0160] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic related to at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0161] Some aspects of this application can be executed entirely by hardware, entirely by software (including firmware, resident software, microcode, etc.), or by a combination of hardware and software. The aforementioned hardware or software may be referred to as a "data block," "module," "engine," "unit," "component," or "system." The processor may be one or more application-specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DAPDs), programmable logic devices (PLDs), field-programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors, or combinations thereof. Furthermore, aspects of this application may manifest as computer products residing in one or more computer-readable media, including computer-readable program code. For example, computer-readable media may include, but are not limited to, magnetic storage devices (e.g., hard disks, floppy disks, magnetic tapes, etc.), optical discs (e.g., compressed CDs, digital multifunction DVDs, etc.), smart cards, and flash memory devices (e.g., cards, sticks, key drives, etc.).

[0162] Computer readable media can include a propagated data signal with a program code embodied therein, for example, in baseband or as part of a carrier wave. Such a propagated signal can take any of a variety of forms, including, but not limited to, electro-magnetic, optical, or any suitable combination thereof. Computer readable media can be any media that can be accessed by a computer. By way of example, and not limitation, such computer readable media can comprise RAM, ROM, EEPROM, CD-ROM or any other optical disk storage, magnetic disk storage or any other storage device, transmission or physical medium from which a computer can read. In this specification, the term "computer readable medium" is entirely

[0163] It should also be noted that, as used in the specification and the claims, the singular form "a", "an" and "the" include plural references unless the context clearly dictates otherwise. As used herein, the expression "and / or" includes combinations of all of the associated terms.

[0164] Some embodiments use numerical descriptors of components, attributes, and quantities. It should be understood that such numerical descriptors used in the description of embodiments are, in some examples, modified by the terms "about", "approximately", or "substantially". Unless otherwise stated, "about", "approximately", or "substantially" indicate that the described dimension is permitted to vary by ±20%. Accordingly, numerical values used in the specification and claims are approximations which can vary depending upon the desired properties sought to be obtained by the individual embodiment. In some embodiments, numerical values in the specification and claims have been presented in a range format. It is to be understood that such range format is used for convenience and generally describes and covers one or more of the values falling within the range. Unless otherwise stated, the range format also contemplates subdividing the range into multiple ranges, each having an upper and lower value.

[0165] Although the present application has been described with reference to the current embodiments, persons having ordinary skill in the art will recognize that changes can be made in form and detail without departing from the spirit and the scope of the application. Therefore, the disclosed embodiments should be considered in all respects as illustrative and not restrictive, the scope of the application to be indicated by the appended claims and their equivalents rather than the foregoing description.

Claims

1. A space-borne implemented high-precision orbit extrapolation method, characterized in that, The method comprises the following steps: providing an orbit extrapolation input quantity of a satellite at an initial extrapolation time, wherein the orbit extrapolation input quantity comprises an initial extrapolation position and an initial extrapolation velocity; calculating an extrapolation duration based on the initial extrapolation time and a preset threshold, and determining a terminal extrapolation time of the satellite based on the extrapolation duration; calculating an earth non-spherical gravity perturbation acceleration according to an earth non-spherical gravity perturbation model; calculating a sun-moon three-body gravity acceleration according to a sun-moon three-body gravity model; determining a satellite thrust acceleration according to a current working state of the satellite; adding the earth non-spherical gravity perturbation acceleration, the sun-moon three-body gravity acceleration and the satellite thrust acceleration to obtain an extrapolation combined acceleration; performing orbit extrapolation calculation based on the extrapolation combined acceleration by using a fourth-order Runge-Kutta method to obtain a target extrapolation position and a target extrapolation velocity at the terminal extrapolation time; updating an orbit extrapolation input quantity at a next extrapolation time according to the target extrapolation position and the target extrapolation velocity.

2. The method of claim 1, wherein, The step of calculating the extrapolation duration based on the initial extrapolation time and the preset threshold comprises the following steps: reading an extrapolation starting point valid flag, and judging whether a time difference between a current time and the initial extrapolation time is greater than or equal to the preset threshold when the extrapolation starting point valid flag is valid; if the time difference is greater than or equal to the preset threshold, taking the preset threshold as the extrapolation duration, otherwise, taking the time difference as the extrapolation duration.

3. The method of claim 2, wherein, The step of calculating the earth non-spherical gravity perturbation acceleration according to the earth non-spherical gravity perturbation model comprises the following steps: obtaining an inertial system to earth fixed system transfer matrix; calculating an earth fixed system position of the satellite according to the initial extrapolation position and the inertial system to earth fixed system transfer matrix; calculating non-spherical gravity potential coefficients corresponding to the earth fixed system position according to the earth non-spherical gravity perturbation model; calculating non-spherical gravity acceleration of the satellite in the earth fixed system by using the non-spherical gravity potential coefficients; and obtaining the earth non-spherical gravity perturbation acceleration according to the non-spherical gravity acceleration of the satellite in the earth fixed system and a transpose matrix of the inertial system to earth fixed system transfer matrix.

4. The method of claim 2, wherein, The step of calculating the sun-moon three-body gravity acceleration according to the sun-moon three-body gravity model comprises the following steps: obtaining current Chebyshev standard time, earth-moon system Chebyshev coefficients, solar system Chebyshev coefficients and moon Chebyshev coefficients; calculating earth-moon system center of mass position, sun center of mass position and moon center of mass position according to the current Chebyshev standard time, the earth-moon system Chebyshev coefficients, the solar system Chebyshev coefficients and the moon Chebyshev coefficients; calculating sun gravity acceleration and moon gravity acceleration corresponding to the initial extrapolation position according to the sun-moon three-body gravity model; adding the sun gravity acceleration and the moon gravity acceleration to obtain the sun-moon three-body gravity acceleration.

5. The method of claim 2, wherein, The step of determining the satellite thrust acceleration according to the current working state of the satellite comprises the following steps: if the current working state of the satellite is a orbit control mode, measuring the satellite thrust acceleration by using an onboard accelerometer; if the current working state of the satellite is a non-orbit control mode, setting the satellite thrust acceleration to zero.

6. The method according to any one of claims 1 to 5, wherein, The fourth-order Runge-Kutta method is used to calculate the orbit extrapolation based on the extrapolation combined acceleration to obtain the target extrapolation position and target extrapolation velocity at the terminal extrapolation time. Let Vel in be the initial extrapolated velocity, Pos in be the initial extrapolated position, T S be the extrapolation time, then the formula for the orbit extrapolation calculation is: ， ， wherein , where Pos out is the target extrapolation position, Vel out is the target extrapolation velocity.

7. A space-borne implemented high-precision orbit extrapolation device, characterized in that, The method comprises the following steps: The orbit extrapolation time length calculation module is configured to provide orbit extrapolation input quantities of the satellite at the initial extrapolation time, wherein the orbit extrapolation input quantities comprise an initial extrapolation position and an initial extrapolation velocity. The extrapolation time length is calculated based on the initial extrapolation time and a preset threshold, and the terminal extrapolation time of the satellite is determined based on the extrapolation time length. The orbit perturbation force calculation module is configured to calculate the earth non-spherical gravity perturbation acceleration according to an earth non-spherical gravity perturbation model; calculate the sun and moon three-body gravity acceleration according to a sun and moon three-body gravity model; determine the satellite thrust acceleration according to the current working state of the satellite; and add the earth non-spherical gravity perturbation acceleration, the sun and moon three-body gravity acceleration and the satellite thrust acceleration to obtain the extrapolation combined acceleration. The position and velocity information solving module is configured to calculate the orbit extrapolation based on the extrapolation combined acceleration by using the fourth-order Runge-Kutta method to obtain the target extrapolation position and target extrapolation velocity at the terminal extrapolation time. The position and velocity information updating module is configured to update the orbit extrapolation input quantities at the next extrapolation time according to the target extrapolation position and target extrapolation velocity.

8. An electronic device, comprising: The method comprises the following steps: One or more processors; And One or more memories coupled to the one or more processors and storing instructions thereon, which, when executed by the one or more processors alone or collectively, cause the electronic device to perform the method of any one of claims 1-7.

9. A non-transitory computer-readable storage medium storing machine executable instructions, the method comprising: The machine-executable instructions, when executed by one or more processors of a machine, cause the machine to perform the method of any one of claims 1-7.

10. A computer program product comprising machine executable instructions for: The machine-executable instructions, when executed by one or more processors of a machine, cause the machine to perform the method of any one of claims 1-7.

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