A method, device and equipment for constructing a low-orbit satellite solar pressure model

Abstract

The application discloses a low-orbit satellite solar radiation pressure model construction method, device and equipment, and relates to the technical field of satellite navigation and positioning. In the method, a carrier phase and an accelerometer are used to calculate the perturbation force total acceleration of a low-orbit satellite. The environmental state information of the low-orbit satellite is collected, and the solar radiation pressure acceleration is obtained in combination with the perturbation force total acceleration. Based on the initial radiation coefficient used in the initial low-orbit satellite solar radiation pressure model, the corrected radiation coefficient is obtained in combination with the solar radiation pressure acceleration. The low-orbit satellite solar radiation pressure model of the first correction period is constructed by using the corrected radiation coefficient. In this way, the measured data collected by the on-board carrier phase and the accelerometer is used to calculate the perturbation force total acceleration of the low-orbit satellite, and the corrected radiation coefficient is obtained based on the least square estimation to finely construct the low-orbit satellite solar radiation pressure model, thereby avoiding the serious dependence of the light pressure model of the low-orbit satellite on the prior information parameters in the orbit operation.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of satellite navigation and positioning, and in particular to a method and device for constructing a low-orbit satellite solar radiation pressure model and equipment. BACKGROUND

[0002] Low-orbit satellites play a key role in many fields such as navigation and positioning, communication interconnection, and earth observation, due to their fast orbital period and strong earth sensing capability. The accurate determination of the three-dimensional position of a low-orbit satellite directly determines its application level. The development of space tracking technologies (such as global satellite positioning, satellite Doppler orbit measurement and radio positioning, satellite laser ranging, and very long baseline interferometry) provides high-precision geometric constraint information for low-orbit platforms.

[0003] Under the related art, a low-orbit satellite solar radiation pressure model is mainly constructed based on an analytical method, which specifically describes the light pressure received by a low-orbit satellite from three aspects of a radiation source, a propagation path, and the surface reflection and absorption of the low-orbit satellite. In addition, the low-orbit satellite solar radiation pressure model considers the geometric relationship among the sun, the earth, and the low-orbit satellite to more clearly describe the dynamic changes in the solar radiation pressure acceleration received by the low-orbit satellite during its orbital operation.

[0004] The above method constructs a low-orbit satellite solar radiation pressure model by using reflection, absorption, and diffuse reflection coefficients to describe the interaction between visible light and infrared radiation and the surface of a low-orbit satellite, which relies on prior information parameters of the low-orbit satellite. In addition, the above method assumes that the surface characteristics of the low-orbit satellite are uniform and stable, and considers the influence of solar activity and the rotation period on the solar radiation intensity.

[0005] However, due to the complex structure and dynamic changes of the actual low-orbit satellite surface, as well as the complex changes in solar radiation intensity, the precision of the low-orbit satellite solar radiation pressure model constructed in the above manner is limited, thereby affecting the precise orbit calculation of the low-orbit satellite.

[0006] For example, the reflection and absorption characteristics of the surface of a low-orbit satellite may change due to factors such as material aging, space environment electromagnetic interference, etc. In addition, the change in the solar activity period also causes fluctuations in the solar radiation intensity, which results in the low-orbit satellite solar radiation pressure model being unable to accurately describe the dynamic changes in the solar radiation pressure acceleration received by the low-orbit satellite during its orbital operation, causing significant deviations between the measured value and the actual value of the solar radiation pressure acceleration, thereby reducing the precision and reliability of the low-orbit satellite solar radiation pressure model, and ultimately affecting the precision of the low-orbit satellite precise orbit determination and the application level of the satellite platform. SUMMARY

[0007] The application provides a method, device and equipment for constructing a low-orbit satellite solar radiation pressure model, to avoid the severe dependence of a low-orbit satellite solar radiation pressure model on prior information parameters of a low-orbit satellite in orbit operation.

[0008] In a first aspect, the embodiments of the application provide a method for constructing a low-orbit satellite solar radiation pressure model, which comprises the following steps:

[0009] In a first correction period, total perturbation acceleration of a low-orbit satellite is calculated; the total perturbation acceleration is used to describe total acceleration of the low-orbit satellite based on perturbation forces; the perturbation forces are dynamic forces of the low-orbit satellite based on celestial bodies and space environment;

[0010] Environmental state information of the low-orbit satellite is collected; the environmental state information is used to describe orbit operation state of the low-orbit satellite and surrounding operation space environment;

[0011] Based on the total perturbation acceleration and the environmental state information, solar radiation pressure acceleration is obtained.

[0012] Based on an initial radiation coefficient used by an initial low-orbit satellite solar radiation pressure model, a corrected radiation coefficient is obtained in combination with the solar radiation pressure acceleration.

[0013] The low-orbit satellite solar radiation pressure model of the first correction period is constructed by using the corrected radiation coefficient.

[0014] In an optional embodiment, the method further comprises the following steps:

[0015] Satellite-borne GNSS data of the low-orbit satellite is obtained, and ionospheric errors of the satellite-borne GNSS data are eliminated to obtain ionosphere-free satellite-borne GNSS data; the ionosphere-free satellite-borne GNSS data is used to describe actual orbit operation parameters of the low-orbit satellite.

[0016] Based on an earth gravity field model and an initial operation orbit of the low-orbit satellite, theoretical orbit operation parameters of the low-orbit satellite are obtained.

[0017] An error mapping relationship between the actual orbit operation parameters and the theoretical orbit operation parameters is established.

[0018] Based on the error mapping relationship, a dynamics model of the low-orbit satellite is adjusted in a direction of error reduction; the dynamics model is a model used to describe force conditions of the low-orbit satellite in orbit operation, and the force conditions include various perturbation forces.

[0019] Based on the adjusted dynamics model, total perturbation acceleration of the low-orbit satellite is obtained.

[0020] In an alternative embodiment, the method further comprises:

[0021] obtaining an initial perturbation force total acceleration of the low-orbit satellite collected by the satellite-borne accelerometer;

[0022] constructing an attitude conversion matrix of the low-orbit satellite based on a mapping relationship between a satellite-fixed coordinate system of the low-orbit satellite and an orbit operation coordinate system of the low-orbit satellite;

[0023] converting the initial perturbation force total acceleration from the satellite-fixed coordinate system of the low-orbit satellite to the orbit operation coordinate system of the low-orbit satellite based on the attitude conversion matrix to obtain the perturbation force total acceleration of the low-orbit satellite.

[0024] In an alternative embodiment, the method further comprises:

[0025] The perturbation force total acceleration is generated based on at least one of the following accelerations: solar radiation pressure acceleration, earth radiation pressure acceleration, and atmospheric drag acceleration;

[0026] The solar radiation pressure acceleration is used to describe the acceleration of the low-orbit satellite caused by the force of solar radiation pressure;

[0027] The earth radiation pressure acceleration is used to describe the acceleration of the low-orbit satellite caused by the earth's own infrared radiation pressure and reflected solar radiation pressure;

[0028] The atmospheric drag acceleration is used to describe the acceleration of the low-orbit satellite caused by atmospheric drag.

[0029] In an alternative embodiment, the method further comprises:

[0030] The environment state information contains orbit operation state parameters and running space environment parameters;

[0031] The solar radiation pressure acceleration corresponding to the perturbation force total acceleration and the environment state information is obtained, including:

[0032] Based on the orbit operation state parameters and the running space environment parameters, the earth radiation pressure acceleration and the atmospheric drag acceleration of the low-orbit satellite are calculated respectively in a low-orbit satellite orbit operation environment in a first correction period;

[0033] The earth radiation pressure acceleration and the atmospheric drag acceleration are removed from the perturbation force total acceleration to obtain the solar radiation pressure acceleration of the low-orbit satellite.

[0034] In an alternative embodiment, the method further comprises:

[0035] obtaining a correction quantity of the radiation coefficient based on the error mapping relationship, in combination with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient;

[0036] obtaining an initial solar radiation pressure acceleration corresponding to the initial low-orbit satellite solar radiation pressure model, and establishing an error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration;

[0037] obtaining a correction quantity of the radiation coefficient based on the error mapping relationship, in combination with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient;

[0038] superimposing the radiation coefficient correction quantity and the initial radiation coefficient to obtain the corrected radiation coefficient.

[0039] In a second aspect, an embodiment of the present application provides a low-orbit satellite solar radiation pressure model construction device, comprising:

[0040] an acquisition module configured to calculate a perturbation force total acceleration of a low-orbit satellite in a first correction period, wherein the perturbation force total acceleration is used to describe a total acceleration of the low-orbit satellite based on a perturbation force, the perturbation force is a dynamic force of the low-orbit satellite based on celestial bodies and space environment, and environment state information of the low-orbit satellite is collected, wherein the environment state information is used to describe an orbit operation state of the low-orbit satellite and a surrounding operation space environment;

[0041] a processing module configured to obtain a solar radiation pressure acceleration based on the perturbation force total acceleration and the environment state information, obtain a corrected radiation coefficient based on an initial radiation coefficient used by an initial low-orbit satellite solar radiation pressure model in combination with the solar radiation pressure acceleration, and construct a low-orbit satellite solar radiation pressure model of the first correction period by using the corrected radiation coefficient.

[0042] In an optional embodiment, the acquisition module is specifically configured to:

[0043] acquire on-board GNSS data of the low-orbit satellite, eliminate ionospheric errors of the on-board GNSS data, and obtain ionosphere-free on-board GNSS data, wherein the ionosphere-free on-board GNSS data is used to describe actual orbit operation parameters of the low-orbit satellite;

[0044] obtain theoretical orbit operation parameters of the low-orbit satellite based on an earth gravity field model and an initial operation orbit of the low-orbit satellite;

[0045] establish an error mapping relationship between the actual orbit operation parameters and the theoretical orbit operation parameters;

[0046] Based on the error mapping relationship, the dynamic model of the low-orbit satellite is adjusted in the error reduction direction; the dynamic model is a model used to describe the force condition of the low-orbit satellite in the orbit operation process, and the force condition includes various perturbation forces;

[0047] Based on the adjusted dynamic model, the total perturbation acceleration of the low-orbit satellite is obtained.

[0048] In an optional embodiment, the acquisition module is further specifically used for:

[0049] The initial total perturbation acceleration of the low-orbit satellite collected by the on-board accelerometer is acquired;

[0050] Based on the mapping relationship between the star-fixed coordinate system of the low-orbit satellite and the orbit operation coordinate system of the low-orbit satellite, an attitude conversion matrix of the low-orbit satellite is constructed;

[0051] Based on the attitude conversion matrix, the initial total perturbation acceleration is converted from the star-fixed coordinate system of the low-orbit satellite to the orbit operation coordinate system of the low-orbit satellite, and the total perturbation acceleration of the low-orbit satellite is obtained.

[0052] In an optional embodiment, the processing module is specifically used for:

[0053] The total perturbation acceleration and the environment state information are used to obtain the corresponding solar radiation pressure acceleration, including:

[0054] Based on the orbit operation state parameters and the operation space environment parameters, the earth radiation pressure acceleration and the atmospheric resistance acceleration received by the low-orbit satellite in the low-orbit satellite orbit operation environment are calculated respectively;

[0055] The earth radiation pressure acceleration and the atmospheric resistance acceleration are removed from the total perturbation acceleration, and the solar radiation pressure acceleration of the low-orbit satellite is obtained.

[0056] In an optional embodiment, the processing module is further used for:

[0057] Based on the initial radiation coefficient used by the initial low-orbit satellite solar radiation pressure model, the corrected radiation coefficient is obtained in combination with the solar radiation pressure acceleration, including:

[0058] The initial solar radiation pressure acceleration corresponding to the initial low-orbit satellite solar radiation pressure model is acquired, and an error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration is established;

[0059] Based on the error mapping relationship, the correction amount of the radiation coefficient is obtained in combination with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient.

[0060] Superimpose the radiation coefficient correction quantity and the initial radiation coefficient to obtain the corrected radiation coefficient.

[0061] In a third aspect, the present application provides an electronic device, comprising:

[0062] a memory for storing a computer program;

[0063] a processor for executing the computer program stored on the memory to implement the steps of the data updating method.

[0064] In a fourth aspect, the present application provides a computer readable storage medium, wherein the computer readable storage medium stores a computer program, and the computer program is executed by a processor to implement the steps of the data updating method.

[0065] In the embodiments of the present application, the total perturbation force acceleration and the environmental state parameters collected in real time for the low-orbit satellite are taken as inputs, and the actual solar radiation pressure acceleration in the first correction period is obtained through decomposition calculation; secondly, the initial radiation coefficient of the initial low-orbit satellite solar radiation pressure model is used in combination with the solar radiation pressure acceleration in the first correction period to obtain the corrected radiation coefficient, so that the corrected radiation coefficient always tracks the changes of the latest measured data, and ensures that the low-orbit satellite solar radiation pressure model continuously reflects the actual state of the light pressure received by the low-orbit satellite, effectively avoiding the dependence of the traditional method on the prior parameters such as material reflectivity.

[0066] In addition, the solar radiation pressure acceleration is recalculated based on the latest measured perturbation force data and environmental parameters in each correction period, and the radiation coefficient is adjusted accordingly. Through this data-driven way, the model can adapt to the surface performance changes of the low-orbit satellite and environmental disturbances, and avoid the dependence on the pre-accurately set material parameters. BRIEF DESCRIPTION OF DRAWINGS

[0067] Figure 1 A structural schematic diagram of a system architecture provided by the embodiments of the present application;

[0068] Figure 2 A flowchart of a construction method of a low-orbit satellite solar radiation pressure model provided by the embodiments of the present application;

[0069] Figure 3 A logic schematic diagram of the acceleration of the perturbation force received by a low-orbit satellite provided by the embodiments of the present application;

[0070] Figure 4 A logic schematic diagram of obtaining the total perturbation force acceleration based on the on-board GNSS data provided by the embodiments of the present application;

[0071] Figure 5A logic diagram for obtaining perturbation total acceleration based on a satellite-borne accelerometer is provided for an embodiment of the present application.

[0072] Figure 6 A logic diagram for an environmental state parameter of a low earth orbit satellite is provided for an embodiment of the present application.

[0073] Figure 7 A flowchart of a calculation method for a corrected radiation coefficient is provided for an embodiment of the present application.

[0074] Figure 8 A structural diagram of a construction device for a low earth orbit satellite solar pressure model is provided for an embodiment of the present application.

[0075] Figure 9 A structural diagram of an electronic device is provided for an embodiment of the present application. DETAILED DESCRIPTION

[0076] In order to make the purpose, technical scheme and advantages of the present application clearer, the present application will be further described in detail below with reference to the drawings. The specific operation method in the method embodiment can also be applied to the device embodiment or the system embodiment. It should be noted that in the description of the present application, "multiple" is understood as "at least two". The association relationship of the associated objects is described, which means that there can be three relationships, for example, A and / or B can represent the following three cases: A exists alone, A and B exist together, and B exists alone. A and B are connected, which means that A and B are directly connected and A and B are connected through C. In addition, in the description of the present application, "first", "second", etc. are used only for the purpose of distinguishing the description, and cannot be understood as indicating or implying relative importance, nor can it be understood as indicating or implying order.

[0077] Low earth orbit satellites play a key role in many fields such as navigation and positioning, communication interconnection, and earth observation, due to their fast orbital period and strong earth sensing capability. The accurate determination of the three-dimensional position of a low earth orbit satellite directly determines its application level. The development of space tracking technologies (such as global satellite positioning, satellite Doppler orbit measurement and radio positioning, satellite laser ranging, and very long baseline interferometry) provides high-precision geometric constraint information for low earth orbit platforms.

[0078] Under the related art, a low earth orbit satellite solar pressure model is mainly constructed based on an analytical method, which specifically describes the solar pressure force received by the low earth orbit satellite from three aspects of radiation source, propagation path, and surface reflection and absorption of the low earth orbit satellite. In addition, the low earth orbit satellite solar pressure model considers the geometric relationship among the sun, the earth, and the low earth orbit satellite, so as to more clearly describe the dynamic change of the solar pressure acceleration received by the low earth orbit satellite in the process of orbital operation.

[0079] The method constructs a low-orbit satellite solar radiation pressure model of a low-orbit satellite through reflection, absorption and diffuse reflection coefficients to describe the interaction between visible light and infrared radiation and the surface of the low-orbit satellite, and the method relies on prior information parameters of the low-orbit satellite. In addition, the method assumes that the surface characteristics of the low-orbit satellite are uniform and stable, and considers the influence of solar activity and the rotation period on the solar radiation intensity.

[0080] However, due to the complex structure and dynamic changes of the actual low-orbit satellite surface and the complex changes of the solar radiation intensity, the accuracy of the low-orbit satellite solar radiation pressure model constructed in the above manner is limited to a certain extent, thereby affecting the normal operation of the low-orbit satellite.

[0081] To solve the above technical problems, in the embodiments of the present application, in the first correction period, the perturbation force total acceleration of the low-orbit satellite is calculated, the environmental state information of the low-orbit satellite is collected, and based on the perturbation force total acceleration and the environmental state information, the solar radiation pressure acceleration is obtained, based on the initial radiation coefficient used by the initial low-orbit satellite solar radiation pressure model, the corrected radiation coefficient is obtained in combination with the solar radiation pressure acceleration, and the first correction period low-orbit satellite solar radiation pressure model is constructed by using the corrected radiation coefficient. By this way, the problem that the radiation coefficient of the traditional low-orbit satellite solar radiation pressure model excessively relies on the prior parameters of the low-orbit satellite and cannot adapt to the changes of the space environment of the low-orbit satellite orbit operation can be avoided.

[0082] The system architecture diagram to which the technical solutions of the embodiments of the present application are applicable will be briefly introduced below, and it should be noted that the system architecture diagram introduced below is only used to illustrate the embodiments of the present application and is not limited.

[0083] For example, refer to Figure 1 Fig. 1 is a system architecture diagram to which the embodiments of the present application are applicable, and the system architecture at least includes a low-orbit satellite 101, a sun 102, an earth 103 and an atmosphere 104.

[0084] The low-orbit satellite 101 is a core carrier for space exploration, and the low-orbit satellite 101 is subjected to the perturbation force generated by the sun 102, the earth 103 and the atmosphere 104, thereby generating a perturbation force total acceleration;

[0085] The sun 102 generates a solar radiation pressure force on the low-orbit satellite 101, so that the low-orbit satellite generates a solar radiation pressure acceleration;

[0086] The earth 103 generates a radiation force on the low-orbit satellite 101, so that the low-orbit satellite generates an earth radiation acceleration;

[0087] The atmosphere 104 generates an atmospheric resistance on the low-orbit satellite 101, so that the low-orbit satellite generates an atmospheric resistance acceleration.

[0088] In addition, the low-orbit satellite 101 calculates the corresponding corrected radiation coefficient based on the total perturbation force acceleration generated by the sun 102, the earth 103 and the atmosphere 104 in real time statistics, in combination with the environmental state parameters, and updates to the initial low-orbit satellite solar pressure model, to realize the accurate construction of the low-orbit satellite solar pressure model.

[0089] For example, referring to Figure 2 In the embodiment of the application, the specific process of constructing the low-orbit satellite solar pressure model is as follows:

[0090] Step 200, in the first correction period, the total perturbation force acceleration of the low-orbit satellite is calculated.

[0091] Specifically, referring to Figure 3 In the embodiment of the application, the total perturbation force acceleration of the low-orbit satellite collected in the first correction period is mainly used to describe the total acceleration of the low-orbit satellite based on the perturbation force, and the perturbation force is the dynamic force generated by the low-orbit satellite based on celestial bodies and space environment.

[0092] In practical application, the perturbation force mainly includes the light pressure radiation from the sun to the satellite surface, the earth radiation pressure composed of the short-wave radiation reflected by the earth surface and the long-wave radiation generated by itself, and the atmospheric resistance generated by the high-altitude atmosphere of the earth. Correspondingly, the acceleration generated by the solar pressure on the surface of the low-orbit satellite is the solar pressure acceleration, the acceleration generated by the earth radiation pressure on the surface of the low-orbit satellite is the earth radiation pressure acceleration, and the acceleration generated by the atmospheric resistance on the surface of the low-orbit satellite is the atmospheric resistance acceleration.

[0093] Therefore, the so-called total perturbation force acceleration is the integration result of the solar pressure acceleration, the earth radiation pressure acceleration and the atmospheric resistance acceleration, which is a whole description parameter and does not specifically distinguish the parameter size of the solar pressure acceleration, the earth radiation pressure acceleration and the atmospheric resistance acceleration.

[0094] For example, the solar pressure acceleration is denoted as The earth radiation pressure acceleration is denoted as The atmospheric resistance acceleration is denoted as Therefore, , is the total perturbation force acceleration collected. In this way, by analyzing and considering the perturbation force generated by the sun, the earth and the atmosphere on the orbit operation of the low-orbit satellite, the force condition of the low-orbit satellite in the complex space environment can be more accurately reflected, and reliable data support is provided for the construction of the low-orbit satellite solar pressure model of the low-orbit satellite.

[0095] The total perturbation acceleration of a low-Earth orbit (LEO) satellite can be calculated using two methods: if the LEO satellite is not equipped with an onboard accelerometer, the total perturbation acceleration is calculated based on the satellite's onboard GNSS data; if the LEO satellite is equipped with an onboard accelerometer, the total perturbation acceleration is calculated based on the onboard accelerometer data. The steps for calculating the perturbation acceleration of an LEO satellite based on these two data sources are as follows:

[0096] Step 200-1: Calculate the total perturbation acceleration of the low-Earth orbit satellite based on the onboard GNSS data of the low-Earth orbit satellite.

[0097] For details, please refer to Figure 4 As shown in the embodiments of this application, firstly, the onboard GNSS data of the low-Earth orbit satellite is acquired, and the ionospheric error of the onboard GNSS data is eliminated to obtain ionospheric-free onboard GNSS data; then, based on the Earth's gravity field model and the initial operating orbit of the low-Earth orbit satellite, the theoretical orbital operating parameters of the low-Earth orbit satellite are obtained; next, an error mapping relationship is established between the actual orbital operating parameters and the theoretical orbital operating parameters; finally, based on the error mapping relationship, the dynamic model of the low-Earth orbit satellite is adjusted in the direction of error reduction, and based on the adjusted dynamic model, the total perturbation acceleration of the low-Earth orbit satellite is obtained.

[0098] For example, a low-Earth orbit (LEO) satellite's onboard GNSS receiver receives pseudorange and phase observations from navigation satellite S at different frequencies as onboard GNSS data. However, this onboard GNSS data is susceptible to ionospheric interference, leading to errors. Therefore, dual-frequency combination technology is needed to eliminate ionospheric errors, based on:

[0099] Obtain ionospheric-free spaceborne GNSS data. Where, and The pseudorange observations corresponding to the dual-frequency data, and The carrier phase observations corresponding to dual-frequency data, and The frequencies correspond to their respective frequency points. Specifically, for pseudorange and phase observations, accuracy indices of 1m and 0.01m are used as prior constraint weighting ratios for parameter estimation in precise orbit determination of low-Earth orbit satellites. The ionospheric pseudorange obtained based on the dual-frequency combination... and carrier phase By combining the prior constraint weight ratio of 1m pseudorange and 0.01m phase, the actual orbital parameters of the low-Earth orbit (LEO) satellite can be obtained through the least-squares precise orbit determination algorithm. Then, using the Earth's gravity field model and the LEO satellite's initial orbit, the theoretical orbital parameters are calculated through numerical integration. Comparing the actual and theoretical orbital parameters yields the position and velocity deviations, which represent the error mapping relationship. These deviations are caused by perturbations such as atmospheric drag and solar radiation pressure. Based on this error mapping relationship, the perturbation-related parameters in the dynamic model are adjusted until the error is minimized. Finally, the adjusted perturbation parameters are substituted into the model to calculate and vector synthesize the accelerations of each perturbation force, thus obtaining the total perturbation acceleration.

[0100] Step 200-2: Calculate the total perturbation acceleration of the low-Earth orbit satellite based on the onboard accelerometer data of the low-Earth orbit satellite.

[0101] For details, please refer to Figure 5 As shown in the embodiment of this application, firstly, the initial total perturbation acceleration of the low-Earth orbit satellite is acquired by the onboard accelerometer; then, based on the mapping relationship between the satellite's fixed coordinate system and its orbital coordinate system, the attitude transformation matrix of the low-Earth orbit satellite is constructed; finally, based on the attitude transformation matrix, the initial total perturbation acceleration is transformed from the satellite's fixed coordinate system to its orbital coordinate system to obtain the total perturbation acceleration of the low-Earth orbit satellite.

[0102] For example, the onboard accelerometer of a low-Earth orbit satellite collects the initial total acceleration caused by the perturbation force experienced by the low-Earth orbit satellite during this period. Since the spaceborne accelerometer is in a star-fixed coordinate system, therefore, The components are described based on the structural orientation of the low-Earth orbit (LEO) satellite. However, analyzing the impact of perturbations on the LEO satellite's orbit requires using the LEO satellite's orbital coordinate system (along the orbital plane, radial, and normal). Therefore, it is necessary to first construct the mapping relationship between the fixed coordinate system and the orbital coordinate system based on the measured pitch, yaw, and roll angles of the LEO satellite, establish the attitude transformation matrix R, and then... Substituting into matrix R, we obtain the total perturbation acceleration in the low-Earth orbit satellite's orbital coordinate system.

[0103] The total perturbation acceleration of low-Earth orbit (LEO) satellites is calculated by selecting an appropriate adaptation method based on whether the satellite has an onboard accelerometer. When the LEO satellite does not have an onboard accelerometer, the total perturbation acceleration is calculated based on the satellite's onboard GNSS data. When the LEO satellite has an onboard accelerometer, the total perturbation acceleration is calculated based on the onboard accelerometer data. This method can adapt to different hardware configurations, accurately obtain the total perturbation acceleration, and provide support for the construction of LEO satellite solar radiation pressure models.

[0104] Further, in order to more accurately reflect the real force situation of the low-orbit satellite in the orbit operation, it is necessary to perform bias correction on the perturbation total acceleration of the low-orbit satellite.

[0105] Specifically, in the embodiment of the application, since the system error of the on-board accelerometer exists, it is necessary to correct the overall offset of the perturbation total acceleration according to the system preset correction parameter set to obtain the corrected perturbation total acceleration.

[0106] In actual application, the accelerometer correction parameter set includes a bias correction parameter and a scale correction parameter, which are used to correct the system error of the on-board accelerometer to cause the perturbation total acceleration to exist bias.

[0107] For example, the bias correction parameter is denoted as , and the scale correction parameter is denoted as , and the corrected perturbation total acceleration is .

[0108] By removing the calculated earth radiation pressure and atmospheric resistance acceleration from the perturbation total acceleration, the accurate solar pressure acceleration can be obtained, which provides key data support for error mapping to correct the radiation coefficient and constructing an accurate low-orbit satellite solar pressure model.

[0109] Step 201, collecting the environmental state information of the low-orbit satellite.

[0110] Specifically, as shown in Figure 6 , in the embodiment of the application, the environmental state information of the low-orbit satellite can be collected by a satellite load sensor, and the environmental state information is used to describe the orbit operation state of the low-orbit satellite and the surrounding operation space environment. Among them, the orbit operation parameters include at least one of the following: the operation attitude parameter of the low-orbit satellite, the structure parameter of the low-orbit satellite and the surface reflection parameter of the low-orbit satellite. And, the operation space environment parameter includes at least one of the following: the solar radiation intensity and the atmospheric density.

[0111] For example, the direction cosine of the angle between the solar incident light and the illuminated panel of the low-orbit satellite is taken as the operation attitude parameter of the low-orbit satellite, the area A and the mass m of the surface panel of the low-orbit satellite are taken as the structure parameter of the low-orbit satellite, and the mirror reflection parameter , the diffuse reflection parameter and the absorption parameter of the low-orbit satellite are taken as the surface reflection parameter of the low-orbit satellite, which together constitute the orbit operation parameters of the low-orbit satellite.

[0112] In addition, the solar flux of the sun reaching the surface of the low-orbit satellite As the intensity of solar radiation, the atmospheric density of the low-orbit satellite operation space is denoted as These parameters jointly constitute the operation space environment parameters of the low-orbit satellite. The orbit operation state parameters and the operation space environment parameters are taken as the environment state information of the low-orbit satellite. The main content of the environment state information of the low-orbit satellite in the first correction period and the related parameters are created, as shown in Table 1:

[0113] Table 1

[0114]

[0115] By obtaining the environment state parameters of the low-orbit satellite, the orbit operation state of the low-orbit satellite and the surrounding space environment can be comprehensively mastered, further data support is provided for the construction of the solar radiation pressure model of the low-orbit satellite, and the construction of the solar radiation pressure model of the low-orbit satellite based on the environment state parameters and the total perturbation acceleration data has strong adaptability in complex space environment.

[0116] Step 202, obtaining the corresponding solar radiation pressure acceleration based on the total perturbation acceleration and the environment state information.

[0117] Specifically, in the embodiment of the application, since the total perturbation acceleration collected is an overall description parameter and does not specifically distinguish the parameter size of the solar radiation pressure acceleration, the earth radiation pressure acceleration and the atmospheric drag acceleration, based on the orbit operation state parameters and the operation space environment parameters, the earth radiation pressure acceleration and the atmospheric drag acceleration received by the low-orbit satellite in the low-orbit satellite orbit operation environment are calculated respectively, and then the earth radiation pressure acceleration and the atmospheric drag acceleration are removed from the total perturbation acceleration to obtain the solar radiation pressure acceleration of the low-orbit satellite.

[0118] For example, the earth radiation pressure acceleration received by the low-orbit satellite in the low-orbit satellite orbit operation environment in the first correction period is denoted as , and the atmospheric drag acceleration is denoted as , then the solar radiation pressure acceleration received by the low-orbit satellite in the low-orbit satellite orbit operation environment in the first correction period is .

[0119] By calculating the earth radiation pressure acceleration and the atmospheric drag acceleration by using the orbit operation state parameters and the operation space parameters respectively, it is ensured that the solar radiation pressure acceleration can be accurately separated from the overall description of the total perturbation acceleration, and the accuracy of the obtained solar radiation pressure acceleration is ensured.

[0120] Step 203, obtaining the corrected radiation coefficient based on the initial radiation coefficient used by the initial low-orbit satellite solar radiation pressure model and combining the solar radiation pressure acceleration.

[0121] Specifically, referring toFigure 7 As shown in the execution of step 203, the following operations can be included:

[0122] Step 203-1, obtaining an initial solar radiation pressure acceleration corresponding to an initial low-orbit satellite solar radiation pressure model, and establishing an error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration.

[0123] Specifically, in the embodiments of the present application, the change of the radiation coefficient will cause the change of the solar radiation pressure acceleration, and the change can be approximated by the partial derivative (first-order Taylor expansion approximation). Therefore, an error mapping relationship between the actual solar radiation pressure acceleration collected by the on-board accelerometer and the initial solar radiation pressure acceleration can be established. Before constructing the error function relationship, the partial derivative expression of the radiation coefficient needs to be derived.

[0124] For example, on the basis of the known low-orbit satellite flight attitude, low-orbit satellite macro model (geometric structure) and optical properties (reflection and absorption), the solar radiation pressure acceleration received by a single panel on the surface of the low-orbit satellite can be obtained by using formula (1):

[0125] (1)

[0126] In the formula, is the ratio of the area (unit: square meter) of the panel material to the mass (unit: kilogram), is the value of the speed of light, which is 299792458 m / s, is the light pressure scale parameter, is the angle between the incident light and the normal of the panel, and are the radiation coefficients of the visible light and infrared wave bands, is the earth shadow function. is the radiation intensity of the sun reaching the surface of the low-orbit satellite (solar flux), which can be characterized by the solar constant to represent the total solar irradiance of the low-orbit satellite in the near-earth orbit, and can be obtained by using formula (2).

[0127] (2)

[0128] In the formula, is an astronomical unit distance, representing the average distance from the sun to the earth, and the value is 1.495978707 meters, is the instantaneous distance from the low-orbit satellite to the sun. The solar radiation pressure received by the surface of the low-orbit satellite can be modeled as three parts of specular reflection, diffuse reflection and absorption when quantitatively described, and the radiation coefficients of the visible light and infrared wave bands can be calculated according to formula (3).

[0129] (3)

[0130] In the formula, , and These represent the specular reflectivity, diffuse reflectivity, and absorptivity of the surface material of a low-Earth orbit satellite, respectively. It is the normal of the low-orbit satellite panel. This is the vector direction from the Sun to the low-Earth orbit satellite. Absorbed solar photons radiate energy outward as heat (secondary radiation), which, according to Newton's third law of motion, creates a reaction force, known as the celestial body's thermal re-radiation force.

[0131] Using the reflection, diffuse reflection, and absorption coefficients of the low-orbit satellite radiation panel in the dual-band as parameters to be estimated, the partial derivative of the radiation coefficient of the physical analytical type of light pressure in the visible light band can be obtained using formula (4).

[0132]

[0133] (4)

[0134]

[0135] Formula (4) above describes the partial derivative of the panel radiation coefficient of the low-orbit satellite body and the solar sail.

[0136] Considering the accelerometer data of low-orbit satellites The relationship between the radiation coefficient and the solar radiation coefficient is nonlinear, making direct calculation of the correction factor for the radiation coefficient quite difficult. Therefore, based on the principle that the effect of a small change in the radiation coefficient on solar radiation pressure acceleration can be approximated using the first-order partial derivative, the following approach can be used: Initial radiation coefficient around low-Earth orbit satellites Perform a Taylor series expansion and retain the first-order terms to express the solar radiation pressure acceleration. The nonlinear relationship between the radiation coefficient and the radiation coefficient is converted into a linear relationship between the correction factor of the radiation coefficient and the correction factor of the solar radiation pressure acceleration.

[0137] Based on the total acceleration of low-orbit satellite perturbation Combining the Taylor expansion results, this equation can be rewritten as a linear observation equation. .in, It is the area of ​​panel k. It is the initial low-Earth orbit satellite panel k-band radiation coefficient. , and Let be the partial derivative of solar radiation pressure acceleration with respect to each radiation coefficient. , and This is the correction factor between the initial radiation coefficient and the true radiation coefficient (parameter to be estimated). and for other known perturbation forces acceleration.

[0138] the i-th epoch of the onboard accelerometer observation data , define the residual as the difference between the theoretical total perturbation force acceleration containing the correction number and the actual observation value , that is , The physical meaning of is the deviation of the prior theoretical total perturbation force acceleration without the correction term from the actual observation value , and Simplify the partial derivative symbol , and finally obtain the error equation of the i-th epoch When the onboard accelerometer has n sets of observation data

[0139] Step 203-2, based on the error mapping relationship, combining the initial low-orbit satellite solar pressure model and the corresponding initial radiation coefficient, the correction amount of the radiation coefficient is obtained.

[0140] Specifically, in the embodiments of the application, the core method for solving the to-be-estimated parameter (the correction number of the radiation coefficient) is the least square method, and the optimal solution is found by minimizing the sum of squares of residuals to obtain the to-be-estimated parameter.

[0141] For example, when the onboard accelerometer has n sets of observation data , the error equations of all epochs can be integrated into a matrix form. Define

[0142]

[0143] The matrix form of the overall error equation is . The core of the least square method is to minimize the sum of squares of residual vectors (i.e., to minimize the overall deviation between the calculated value and the observed value). Define the objective function as , and into the objective function to obtain . In order to minimize , the derivative of is taken and set to 0. Through matrix derivative operation, the normal equation (the necessary condition for the least square optimal solution) is derived as . Then, solving the equation can obtain the optimal to-be-estimated parameter vector = .

[0144] Step 203-3, superimpose the radiation coefficient correction amount and the initial radiation coefficient to obtain the corrected radiation coefficient.

[0145] Specifically, in the embodiments of the present application, the correction amount of the radiation coefficient is obtained, and the initial radiation coefficient is known, so that the correction amount of the radiation coefficient is superimposed on the initial radiation coefficient to obtain the optimal radiation coefficient, that is, the corrected radiation coefficient.

[0146] For example, when the correction amount of the radiation coefficient obtained by solving is = After that, assuming that the initial radiation coefficient is , and , the corrected radiation coefficient can be obtained as , and . The essence of this process is to back-propagate the parameter correction amount through the observation data, so that the solar pressure acceleration calculated by the model is maximally fitted to the actual observation value, thereby realizing accurate estimation of the radiation coefficient.

[0147] By linearizing the nonlinear relationship around the initial radiation coefficient through Taylor expansion, establishing an error mapping and solving the radiation coefficient correction amount by the least square method, and then superimposing the correction amount on the initial radiation coefficient, the nonlinear relationship can be converted into a solvable linear problem, the parameter correction is back-propagated through the observation data, the solar pressure acceleration calculated by the model is fitted to the actual observation, and accurate estimation of the radiation coefficient is realized, thereby providing reliable parameter support for the solar pressure model correction of the low-orbit satellite.

[0148] Step 204, constructing a low-orbit satellite solar pressure model of a first correction period by using the corrected radiation coefficient.

[0149] Specifically, in the embodiments of the present application, the correction amount of the radiation coefficient has been solved by the least square method, and based on the correction amount of the radiation coefficient and the initial radiation coefficient, the corrected radiation coefficient is obtained, so that the accurate low-orbit satellite solar pressure model can be constructed based on the corrected radiation coefficient.

[0150] By using the corrected radiation coefficient estimated by the least square method to estimate the actual force of the low-orbit satellite, a low-orbit satellite solar pressure model suitable for the orbit environment of the first correction period can be constructed, thereby providing accurate dynamic model support for precise orbit calculation and high-precision prediction of the low-orbit satellite.

[0151] Based on the same inventive concept, the embodiments of the present application also provide a low-orbit satellite solar pressure model construction device, for example, as shown in Figure 8 The data security transmission device 800 includes:

[0152] The acquisition module 8000 is configured to: calculate total perturbation force acceleration of the low-orbit satellite in a first correction period, the total perturbation force acceleration is used to describe total acceleration of the low-orbit satellite based on perturbation force, the perturbation force is a dynamic force of the low-orbit satellite based on celestial bodies and space environment, and collect environment state information of the low-orbit satellite, the environment state information is used to describe orbit operation state of the low-orbit satellite and surrounding operation space environment.

[0153] The processing module 8001 is configured to: obtain solar radiation pressure acceleration based on the total perturbation force acceleration and the environment state information, obtain corrected radiation coefficient based on an initial radiation coefficient used by an initial solar radiation pressure model of the low-orbit satellite, and construct a low-orbit satellite solar radiation pressure model of the first correction period by using the corrected radiation coefficient.

[0154] In an optional embodiment, the acquisition module 8000 is specifically configured to:

[0155] The total perturbation force acceleration of the low-orbit satellite is calculated, including:

[0156] The on-board GNSS data of the low-orbit satellite is acquired, and ionospheric error of the on-board GNSS data is eliminated to obtain ionosphere-free on-board GNSS data, the ionosphere-free on-board GNSS data is used to describe actual orbit operation parameters of the low-orbit satellite.

[0157] Theoretical orbit operation parameters of the low-orbit satellite are obtained based on an earth gravity field model and an initial operation orbit of the low-orbit satellite.

[0158] An error mapping relationship between the actual orbit operation parameters and the theoretical orbit operation parameters is established.

[0159] Based on the error mapping relationship, a dynamic model of the low-orbit satellite is adjusted in a direction of error reduction, the dynamic model is used to describe force conditions of the low-orbit satellite in an orbit operation process, and the force conditions include various perturbation forces.

[0160] The total perturbation force acceleration of the low-orbit satellite is obtained based on the adjusted dynamic model.

[0161] In an optional embodiment, the acquisition module 8000 is further specifically configured to:

[0162] The total perturbation force acceleration of the low-orbit satellite is calculated, including:

[0163] The initial total perturbation force acceleration of the low-orbit satellite detected by the on-board accelerometer is acquired.

[0164] An attitude conversion matrix of the low-orbit satellite is constructed based on a mapping relationship between a star-fixed coordinate system of the low-orbit satellite and an orbit operation coordinate system of the low-orbit satellite.

[0165] Based on the attitude conversion matrix, the initial perturbation force total acceleration is converted from the star-fixed coordinate system of the low-orbit satellite to the orbit running coordinate system of the low-orbit satellite, and the perturbation force total acceleration of the low-orbit satellite is obtained.

[0166] In an optional embodiment, the processing module 8001 is specifically configured to:

[0167] The environmental state information includes the orbit running state parameter and the running space environment parameter.

[0168] Based on the perturbation force total acceleration and the environmental state information, the corresponding solar pressure acceleration is obtained, including:

[0169] Based on the orbit running state parameter and the running space environment parameter, the earth radiation pressure acceleration and the atmospheric drag acceleration received by the low-orbit satellite in the orbit running environment of the low-orbit satellite are calculated respectively.

[0170] The earth radiation pressure acceleration and the atmospheric drag acceleration are removed from the perturbation force total acceleration, and the solar pressure acceleration of the low-orbit satellite is obtained.

[0171] In an optional embodiment, the processing module 8001 is further configured to:

[0172] Based on the initial radiation coefficient used by the initial low-orbit satellite solar pressure model, the corrected radiation coefficient is obtained in combination with the solar pressure acceleration, including:

[0173] The initial solar pressure acceleration corresponding to the initial low-orbit satellite solar pressure model is obtained, and an error mapping relationship between the solar pressure acceleration and the initial solar pressure acceleration is established.

[0174] Based on the error mapping relationship, the correction amount of the radiation coefficient is obtained in combination with the initial low-orbit satellite solar pressure model and the corresponding initial radiation coefficient.

[0175] The correction amount of the radiation coefficient is superimposed with the initial radiation coefficient to obtain the corrected radiation coefficient.

[0176] The perturbation force total acceleration of the low-orbit satellite and the environmental state information of the low-orbit satellite are obtained by the obtaining module, the earth radiation pressure and the atmospheric drag acceleration are separated from the perturbation force total acceleration by the processing module to obtain the precise solar pressure acceleration, and the corrected radiation coefficient is obtained by establishing the error mapping, which can adapt to different hardware configurations of the low-orbit satellite, guarantee the accuracy of the solar pressure related data, and finally build the low-orbit satellite solar pressure model that fits the actual orbit environment.

[0177] Based on the same inventive concept, the embodiments of the present application provide a computer device, which can realize the functions of the foregoing low-orbit satellite solar pressure model construction device, and the computer device includes: Figure 9 , the electronic device includes:

[0178] The at least one processor 9001 and the memory 9002 connected with the at least one processor 9001 are not limited to the specific connection medium between the processor 9001 and the memory 9002 in the embodiments of the present application, Figure 9 In the specific connection medium between the processor 9001 and the memory 9002 in the embodiments of the present application, Figure 9 In the specific connection medium between the processor 9001 and the memory 9002 in the embodiments of the present application, Figure 9 In the specific connection medium between the processor 9001 and the memory 9002 in the embodiments of the present application,

[0179] In the embodiments of the present application, the memory 9002 stores instructions executable by the at least one processor 9001, and the at least one processor 9001 can execute the data updating method discussed above by executing the instructions stored in the memory 9002. The processor 9001 can realize Figure 8 The functions of each module in the device shown.

[0180] The processor 9001 is the control center of the device, and can connect each part of the entire control device through various interfaces and lines, and process data and various functions of the device by running or executing the instructions stored in the memory 9002 and calling the data stored in the memory 9002, thereby monitoring the entire device.

[0181] In a possible design, the processor 9001 can include one or more processing units, and the processor 9001 can integrate an application processor and a modem processor, where the application processor mainly processes an operating system, a user interface, and an application program, etc., and the modem processor mainly processes wireless communication. It can be understood that the above-mentioned modem processor can also not be integrated into the processor 9001. In some embodiments, the processor 9001 and the memory 9002 can be implemented on the same chip, and in some embodiments, they can also be implemented on independent chips respectively.

[0182] The processor 9001 can be a general processor, such as a central processing unit (CPU), a digital signal processor, an application-specific integrated circuit, a field programmable gate array or other programmable logic device, a discrete gate or transistor logic component, a discrete hardware component, and can implement or execute the methods, steps and logic block diagrams disclosed in the embodiments of the present application. The general processor can be a microprocessor or any conventional processor. The steps of the method for constructing a low-orbit satellite solar pressure model disclosed in the embodiments of the present application can be directly embodied as execution completed by a hardware processor, or executed by a combination of hardware and software modules in the processor.

[0183] The memory 9002 is a non-volatile computer readable storage medium, which can be used to store non-volatile software programs, non-volatile computer executable programs and modules. The memory 9002 can include at least one type of storage medium, such as flash memory, hard disk, multimedia card, card-type memory, random access memory (RAM), static random access memory (SRAM), programmable read-only memory (PROM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), magnetic storage, magnetic disk, optical disk, etc. The memory 9002 is any other medium capable of carrying or storing desired program code in the form of instructions or data structures and capable of being accessed by a computer, but is not limited thereto. The memory 9002 in the embodiments of the present application can also be a circuit or any other device capable of realizing a storage function, used for storing program instructions and / or data.

[0184] By designing and programming the processor 9001, the code corresponding to the method for constructing a low-orbit satellite solar pressure model introduced in the foregoing embodiments can be fixed into the chip, so that the chip can execute the steps of the method for constructing a low-orbit satellite solar pressure model of the embodiments shown in the figure at runtime. Figure 2 How to design and program the processor 9001 is a technology known to those skilled in the art, which will not be described here.

[0185] Based on the same inventive concept, the embodiments of the present application provide a computer readable storage medium storing a computer program executable by a computer device, which causes the computer device to execute the steps of the road network data detection method when the program runs on the computer device.

[0186] Those skilled in the art will appreciate that embodiments of the present application can be devised for a method, a system, or a computer program product. Accordingly, the present application can take the form of an entirely hardware embodiment, an entirely software embodiment or an embodiment combining software and hardware aspects. Furthermore, the present application can take the form of a computer program product on one or more computer readable storage media (including, but not limited to, disk storage, CD-ROMs, optical storage devices, etc.) embodying computer readable program code.

[0187] The present application is described in reference to the flow diagrams and / or block diagrams of methods, apparatus (systems) and computer program products according to this application. It will be understood that each block of the flow diagrams and / or block diagrams, and combinations of blocks in the flow diagrams and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general purpose computer, special purpose computer, embedded processing machine, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, create means for implementing the functions specified in the flow diagrams and / or block diagrams block or blocks. Figure 1 one or more functions specified in the flow diagram and / or block diagram block or blocks. Figure 1 means for carrying out each of the functionality

[0188] These computer program instructions can also be stored in a computer readable memory that can direct a computer or other programmable data processing apparatus to function in a particular manner, such that the instructions stored in the computer readable memory produce an article of manufacture including instructions which implement the flow diagrams and / or block diagrams block or blocks. Figure 1 one or more functions specified in the flow diagram and / or block diagram block or blocks. Figure 1 means for carrying out each of the functionality

[0189] These computer program instructions can also be loaded onto a computer or other programmable data processing apparatus to cause a series of operational steps to be performed on the computer or other programmable apparatus to produce a computer implemented process such that the instructions that execute on the computer or other programmable apparatus provide steps for implementing the flow diagrams and / or block diagrams block or blocks. Figure 1 one or more functions specified in the flow diagram and / or block diagram block or blocks. Figure 1 means for carrying out each of the functionality

[0190] Obviously, numerous modifications and variations of the present application are possible in light of the above teachings. It is therefore to be understood that within the scope of the appended claims and their equivalents, the application can be practiced otherwise than as specifically described.

Claims

1. A method for constructing a low earth orbit satellite solar radiation pressure model, characterized in that, The method comprises the following steps: In a first correction period, a perturbation force total acceleration of a low-orbit satellite is calculated; The perturbation force total acceleration is used to describe a total acceleration of the low-orbit satellite generated based on a perturbation force; the perturbation force comprises a dynamic force generated by the low-orbit satellite based on celestial bodies and a space environment; Environment state information of the low-orbit satellite is collected; The environment state information is used to describe an orbit operation state of the low-orbit satellite and an operation space environment around the low-orbit satellite; Based on the perturbation force total acceleration and the environment state information, a solar radiation pressure acceleration is obtained; An initial solar radiation pressure acceleration corresponding to an initial low-orbit satellite solar radiation pressure model is acquired, and an error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration is established; Based on the error mapping relationship, a correction amount of a radiation coefficient is obtained in combination with the initial low-orbit satellite solar radiation pressure model and a corresponding initial radiation coefficient; The correction amount of the radiation coefficient is superimposed with the initial radiation coefficient to obtain a corrected radiation coefficient; The low-orbit satellite solar radiation pressure model of the first correction period is constructed by using the corrected radiation coefficient.

2. The method of claim 1, wherein, The calculation of the perturbation force total acceleration of the low-orbit satellite comprises the following steps: Satellite-borne GNSS data of the low-orbit satellite is acquired, and ionospheric errors of the satellite-borne GNSS data are eliminated to obtain ionosphere-free satellite-borne GNSS data; the ionosphere-free satellite-borne GNSS data is used to describe actual orbit operation parameters of the low-orbit satellite; Based on an earth gravity field model and an initial operation orbit of the low-orbit satellite, theoretical orbit operation parameters of the low-orbit satellite are obtained; An error mapping relationship between the actual orbit operation parameters and the theoretical orbit operation parameters is established; Based on the error mapping relationship, a dynamic model of the low-orbit satellite is adjusted in a direction of error reduction; the dynamic model is a model used to describe force conditions of the low-orbit satellite in an orbit operation process, and the force conditions comprise various perturbation forces; Based on the adjusted dynamic model, the perturbation force total acceleration of the low-orbit satellite is obtained.

3. The method of claim 1, wherein, The calculation of the perturbation force total acceleration of the low-orbit satellite comprises the following steps: An initial perturbation force total acceleration of the low-orbit satellite collected by a satellite-borne accelerometer is acquired; Based on a mapping relationship between a satellite-fixed coordinate system of the low-orbit satellite and an orbit operation coordinate system of the low-orbit satellite, an attitude conversion matrix of the low-orbit satellite is constructed; Based on the attitude conversion matrix, the initial perturbation force total acceleration is converted from the satellite-fixed coordinate system of the low-orbit satellite to the orbit operation coordinate system of the low-orbit satellite to obtain the perturbation force total acceleration of the low-orbit satellite.

4. The method of claim 1, wherein, The perturbation force total acceleration is generated based on at least one of the following accelerations: a solar radiation pressure acceleration, an earth radiation pressure acceleration and an atmospheric resistance acceleration; The solar radiation pressure acceleration is used to describe an acceleration generated by the low-orbit satellite due to an action force of solar radiation pressure; The earth radiation pressure acceleration is used to describe an acceleration generated by the low-orbit satellite due to earth's own infrared radiation pressure and reflected solar radiation pressure; The atmospheric resistance acceleration is used to describe an acceleration generated by the low-orbit satellite due to atmospheric resistance.

5. The method according to any one of claims 1 to 4, characterized in that, The environment state information contains orbit operation state parameters and operation space environment parameters; The solar pressure acceleration is obtained based on the perturbation total acceleration and the environment state information, and the solar pressure acceleration comprises: The earth radiation pressure acceleration and the atmospheric resistance acceleration of the low-orbit satellite are calculated based on the orbit operation state parameters and the operation space environment parameters. The solar pressure acceleration of the low-orbit satellite is obtained by removing the earth radiation pressure acceleration and the atmospheric resistance acceleration from the perturbation total acceleration. 6.A device for constructing a low earth orbit satellite solar radiation pressure model, characterized in that, The device comprises: An acquisition module is configured to calculate a perturbation total acceleration of a low-orbit satellite in a first correction period, the perturbation total acceleration being used to describe a total acceleration of the low-orbit satellite based on a perturbation force, the perturbation force being a dynamic force of the low-orbit satellite based on celestial bodies and space environment, and collect environment state information of the low-orbit satellite, the environment state information being used to describe an orbit operation state of the low-orbit satellite and a surrounding operation space environment. A processing module is configured to obtain a corresponding solar pressure acceleration based on the perturbation total acceleration and the environment state information, obtain an initial solar pressure acceleration corresponding to an initial low-orbit satellite solar pressure model, establish an error mapping relationship between the solar pressure acceleration and the initial solar pressure acceleration, obtain a correction amount of a radiation coefficient based on the error mapping relationship, the initial low-orbit satellite solar pressure model and a corresponding initial radiation coefficient, superimpose the correction amount of the radiation coefficient and the initial radiation coefficient to obtain a corrected radiation coefficient, and construct a low-orbit satellite solar pressure model of the first correction period by using the corrected radiation coefficient.

7. A computer device, comprising: The device comprises: A memory is configured to store a computer program. A processor is configured to execute the computer program stored in the memory to implement the method steps in any one of claims 1-5.

8. A computer-readable storage medium, characterized in that, The computer program stored in the memory is executable by a computer device, and when the computer program runs on the computer device, the computer device is caused to execute the steps of the method in any one of claims 1-5.

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