Construction method, device and equipment for solar light pressure model of low earth orbit satellite

Abstract

The invention discloses a method, a device and equipment for constructing a solar light pressure model of a low-orbit satellite, and relates to the technical field of satellite navigation and positioning. In the method, a method based on a carrier phase and an accelerometer is considered to calculate the total perturbative force acceleration of the low-orbit satellite, and the total perturbative force acceleration of the low-orbit satellite is calculated; environment state information of the low-orbit satellite is collected and combined with the total perturbation force acceleration to obtain the sunlight pressure acceleration, and based on the initial radiation coefficient used by the initial low-orbit satellite sunlight pressure model and combined with the sunlight pressure acceleration, the corrected radiation coefficient is obtained; and constructing a low-orbit satellite solar light pressure model of the first correction period by adopting the corrected radiation coefficient. In this way, the total acceleration of the perturbation force of the low-orbit satellite is calculated by using the satellite-borne carrier phase and the actually measured data collected by the accelerometer, and the corrected radiation coefficient is obtained based on the least square estimation to finely construct the solar light pressure model of the low-orbit satellite. And serious dependence of the light pressure model of the low-orbit satellite in orbit operation on prior information parameters is avoided.

Description

Technical Field

[0001] This application relates to the field of satellite navigation and positioning technology, and in particular to a method, apparatus and equipment for constructing a low-orbit satellite solar radiation pressure model. Background Technology

[0002] Low Earth Orbit (LEO) satellites play a crucial role in numerous fields, including navigation, communication, and Earth observation, thanks to their rapid orbital periods and powerful Earth sensing capabilities. The accurate determination of the three-dimensional position of LEO satellites directly determines their application level. The development of space tracking technologies (such as Global Positioning System, Doppler orbit measurement and radio positioning, satellite laser ranging, and Very Long Baseline Interferometry) provides high-precision geometric constraint information for LEO platforms.

[0003] Under relevant technologies, the solar radiation pressure model for low-Earth orbit (LEO) satellites is mainly constructed based on analytical methods. Specifically, it describes the solar radiation pressure experienced by LEO satellites from three aspects: radiation source, propagation path, and surface reflection and absorption of the LEO satellite. Furthermore, this LEO satellite solar radiation pressure model considers the geometric relationship between the Sun, Earth, and the LEO satellite to more clearly describe the dynamic changes in solar radiation pressure acceleration experienced by the LEO satellite during its orbital operation.

[0004] The aforementioned method constructs a solar radiation pressure model for low-Earth orbit (LEO) satellites using reflection, absorption, and diffuse reflection coefficients to describe the interaction between visible and infrared radiation and the LEO satellite surface. This method relies on prior information parameters of the LEO satellite. Furthermore, the method assumes that the LEO satellite surface characteristics are uniform and stable, and considers the influence of solar activity and rotation period on solar radiation intensity.

[0005] However, due to the complex structure and dynamic changes of the actual low-Earth orbit satellite surface, as well as the complex changes in solar radiation intensity, the accuracy of the low-Earth orbit satellite solar radiation pressure model constructed using the above method is limited to a certain extent, thus affecting the precise orbit calculation of low-Earth orbit satellites.

[0006] For example, the reflection and absorption characteristics of the surface of a low-Earth orbit (LEO) satellite may change due to factors such as material aging and electromagnetic interference in the space environment. In addition, changes in the solar activity cycle can also cause fluctuations in solar radiation intensity. As a result, the LEO satellite solar radiation pressure model cannot accurately describe the dynamic changes in solar radiation pressure acceleration experienced by the LEO satellite during its orbital operation. This leads to a significant deviation between the measured and actual values ​​of solar radiation pressure acceleration, thereby reducing the accuracy and reliability of the LEO satellite solar radiation pressure model and ultimately affecting the precision orbit determination accuracy of the LEO satellite and the application level of the satellite platform. Summary of the Invention

[0007] This application provides a method, apparatus, and equipment for constructing a solar radiation pressure model for low-Earth orbit (LEO) satellites, in order to avoid the heavy dependence of the LEO satellite solar radiation pressure model on prior information parameters of the LEO satellite during its orbital operation.

[0008] In a first aspect, embodiments of this application provide a method for constructing a solar radiation pressure model for a low-Earth orbit satellite, the method comprising: In the first calibration period, the total perturbation acceleration of the low-Earth orbit satellite is calculated; the total perturbation acceleration is used to describe the total acceleration generated by the low-Earth orbit satellite based on the perturbation force; the perturbation force is the force generated by the low-Earth orbit satellite based on the celestial body and space environment; Collect environmental status information of the low-Earth orbit satellite; the environmental status information is used to describe the orbital operation status of the low-Earth orbit satellite and the surrounding operating space environment; Based on the total acceleration of the perturbation force and the environmental state information, the solar radiation pressure acceleration is obtained; Based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model, and combined with the solar radiation pressure acceleration, the corrected radiation coefficient is obtained; The corrected radiation coefficient is used to construct a low-orbit satellite solar radiation pressure model for the first correction period.

[0009] In an optional embodiment, the method further includes: 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; the ionospheric-free onboard GNSS data is used to describe the actual orbital operating parameters of the low-Earth orbit satellite. Based on the Earth's gravity field model and the initial orbit of the low-Earth orbit satellite, the theoretical orbital parameters of the low-Earth orbit satellite are obtained. Establish the error mapping relationship between the actual orbital operating parameters and the theoretical orbital operating parameters; Based on the error mapping relationship, the dynamic model of the low-Earth orbit satellite is adjusted in the direction of error reduction; the dynamic model is a model used to describe the force situation of the low-Earth orbit satellite during orbit operation, and the force situation includes various perturbation forces; Based on the adjusted dynamic model, the total perturbation acceleration of the low-orbit satellite is obtained.

[0010] In an optional embodiment, the method further includes: The initial total perturbation acceleration of the low-orbit satellite is obtained from the onboard accelerometer. Based on the mapping relationship between the fixed coordinate system of the low-Earth orbit satellite and the orbital coordinate system of the low-Earth orbit satellite, the attitude transformation matrix of the low-Earth orbit satellite is constructed. Based on the attitude transformation matrix, the initial total perturbation acceleration is transformed from the fixed coordinate system of the low-Earth orbit satellite to the orbital coordinate system of the low-Earth orbit satellite to obtain the total perturbation acceleration of the low-Earth orbit satellite.

[0011] In an optional embodiment, the method further includes: The total perturbation acceleration is generated based on at least one of the following accelerations: solar radiation pressure acceleration, Earth's radiation pressure acceleration, and atmospheric drag acceleration. The solar radiation pressure acceleration is used to describe the acceleration generated by the low-orbit satellite under the force of solar radiation pressure. The Earth radiation pressure acceleration is used to describe the acceleration generated by the Earth's own infrared radiation pressure and reflected solar radiation pressure on the low-orbit satellite. Atmospheric drag acceleration: used to describe the acceleration caused by atmospheric drag experienced by the low-orbit satellite.

[0012] In an optional embodiment, the method further includes: The environmental status information includes track operation status parameters and operating space environment parameters; The process of obtaining the corresponding solar radiation pressure acceleration based on the total perturbation acceleration and the environmental state information includes: Based on the orbital operation status parameters and the operating space environment parameters, the Earth's radiation pressure acceleration and atmospheric drag acceleration experienced by the low-Earth orbit satellite under the orbital operation environment in the first correction period are calculated respectively. The solar radiation pressure acceleration of the low-orbit satellite is obtained by removing the Earth's radiation pressure acceleration and the atmospheric drag acceleration from the total perturbation acceleration.

[0013] In an optional embodiment, the method further includes: Based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model, and combined with the solar radiation pressure acceleration, the corrected radiation coefficient is obtained, including: Obtain the initial solar radiation pressure acceleration corresponding to the initial low-orbit satellite solar radiation pressure model, and establish the error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration; Based on the aforementioned error mapping relationship, and combined with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient, the correction amount of the radiation coefficient is obtained; The corrected radiation coefficient is obtained by superimposing the radiation coefficient correction amount with the initial radiation coefficient.

[0014] Secondly, embodiments of this application provide a device for constructing a low-orbit satellite solar radiation pressure model, comprising: Acquisition module: used to calculate the total acceleration of the perturbation force of the low-Earth orbit satellite during the first correction period; the total acceleration of the perturbation force is used to describe the total acceleration generated by the low-Earth orbit satellite based on the perturbation force; the perturbation force is the force generated by the low-Earth orbit satellite based on the celestial body and space environment; and to collect the environmental state information of the low-Earth orbit satellite; the environmental state information is used to describe the orbital operation state of the low-Earth orbit satellite and the surrounding operating space environment; Processing module: used to obtain solar radiation pressure acceleration based on the total acceleration of the perturbation force and the environmental state information; to obtain the corrected radiation coefficient based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model and the solar radiation pressure acceleration; and to construct the low-Earth orbit satellite solar radiation pressure model for the first correction period using the corrected radiation coefficient.

[0015] In one optional embodiment, the acquisition module is specifically used for: 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; the ionospheric-free onboard GNSS data is used to describe the actual orbital operating parameters of the low-Earth orbit satellite. Based on the Earth's gravity field model and the initial orbit of the low-Earth orbit satellite, the theoretical orbital parameters of the low-Earth orbit satellite are obtained. Establish the error mapping relationship between the actual orbital operating parameters and the theoretical orbital operating parameters; Based on the error mapping relationship, the dynamic model of the low-Earth orbit satellite is adjusted in the direction of error reduction; the dynamic model is a model used to describe the force situation of the low-Earth orbit satellite during orbit operation, and the force situation includes various perturbation forces; Based on the adjusted dynamic model, the total perturbation acceleration of the low-orbit satellite is obtained.

[0016] In an optional embodiment, the acquisition module is further specifically used for: The initial total perturbation acceleration of the low-orbit satellite is obtained from the onboard accelerometer. Based on the mapping relationship between the fixed coordinate system of the low-Earth orbit satellite and the orbital coordinate system of the low-Earth orbit satellite, the attitude transformation matrix of the low-Earth orbit satellite is constructed. Based on the attitude transformation matrix, the initial total perturbation acceleration is transformed from the fixed coordinate system of the low-Earth orbit satellite to the orbital coordinate system of the low-Earth orbit satellite to obtain the total perturbation acceleration of the low-Earth orbit satellite.

[0017] In one optional embodiment, the processing module is specifically used for: The process of obtaining the corresponding solar radiation pressure acceleration based on the total perturbation acceleration and the environmental state information includes: Based on the orbital operation state parameters and the operating space environment parameters, the Earth's radiation pressure acceleration and atmospheric drag acceleration experienced by the low-Earth orbit satellite under the low-Earth orbit satellite operating environment are calculated respectively. The solar radiation pressure acceleration of the low-orbit satellite is obtained by removing the Earth's radiation pressure acceleration and the atmospheric drag acceleration from the total perturbation acceleration.

[0018] In an optional embodiment, the processing module is further configured to: Based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model, and combined with the solar radiation pressure acceleration, the corrected radiation coefficient is obtained, including: Obtain the initial solar radiation pressure acceleration corresponding to the initial low-orbit satellite solar radiation pressure model, and establish the error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration; Based on the aforementioned error mapping relationship, and combined with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient, the correction amount of the radiation coefficient is obtained; The corrected radiation coefficient is obtained by superimposing the radiation coefficient correction amount with the initial radiation coefficient.

[0019] Thirdly, this application provides an electronic device, comprising: Memory, used to store computer programs; When a processor executes a computer program stored in the memory, it implements the steps of the above-described data update method.

[0020] Fourthly, this application provides a computer-readable storage medium storing a computer program that, when executed by a processor, implements the steps of the above-described data update method.

[0021] In this embodiment, the total perturbation acceleration and environmental state parameters collected in real time for the low-Earth orbit satellite are used as inputs. The actual solar radiation pressure acceleration in the first correction period is obtained through decomposition calculation. Then, the initial radiation coefficient of the initial low-Earth orbit satellite solar radiation pressure model is used, combined with the solar radiation pressure acceleration in the first correction period, to obtain the corrected radiation coefficient. This ensures that the corrected radiation coefficient always tracks the changes in the latest measured data, ensuring that the low-Earth orbit satellite solar radiation pressure model continuously reflects the actual state of the solar radiation pressure experienced by the low-Earth orbit satellite. This effectively avoids the dependence of traditional methods on prior parameters such as material reflectivity.

[0022] In addition, the solar radiation pressure acceleration is recalculated based on the latest measured perturbation data and environmental parameters for each calibration cycle, and the radiation coefficient is adjusted accordingly. This data-driven approach enables the model to autonomously adapt to changes in the surface properties of low-Earth orbit satellites and environmental disturbances, avoiding reliance on pre-set material parameters. Attached Figure Description

[0023] Figure 1 A schematic diagram of a system architecture provided in an embodiment of this application; Figure 2 A flowchart illustrating a method for constructing a low-Earth orbit satellite solar radiation pressure model, provided in an embodiment of this application; Figure 3 A logic diagram illustrating the perturbation acceleration experienced by a low-Earth orbit satellite, provided as an embodiment of this application; Figure 4 A logic diagram illustrating the acquisition of total perturbation acceleration based on spaceborne GNSS data, provided in an embodiment of this application; Figure 5 A logic diagram illustrating the acquisition of total perturbation acceleration based on a spaceborne accelerometer, provided in an embodiment of this application; Figure 6 A logical diagram illustrating environmental state parameters of a low-Earth orbit satellite provided in an embodiment of this application; Figure 7 A flowchart illustrating a method for calculating the corrected radiation coefficient, provided in an embodiment of this application; Figure 8 A schematic diagram of a device for constructing a low-orbit satellite solar radiation pressure model provided in an embodiment of this application; Figure 9 This is a schematic diagram of the structure of an electronic device provided in an embodiment of this application. Detailed Implementation

[0024] To make the objectives, technical solutions, and advantages of this application clearer, the application will be further described in detail below with reference to the accompanying drawings. The specific operational methods in the method embodiments can also be applied to the device embodiments or system embodiments. It should be noted that in the description of this application, "multiple" is understood as "at least two". "And / or" describes the relationship between related objects, indicating that three relationships can exist. For example, A and / or B can represent: A existing alone, A and B existing simultaneously, and B existing alone. A connected to B can represent: A and B directly connected, and A and B connected through C. Furthermore, in the description of this application, terms such as "first" and "second" are used only for distinguishing the purpose of description and should not be construed as indicating or implying relative importance or order.

[0025] Low Earth Orbit (LEO) satellites play a crucial role in numerous fields, including navigation, communication, and Earth observation, thanks to their rapid orbital periods and powerful Earth sensing capabilities. The accurate determination of the three-dimensional position of LEO satellites directly determines their application level. The development of space tracking technologies (such as Global Positioning System, Doppler orbit measurement and radio positioning, satellite laser ranging, and Very Long Baseline Interferometry) provides high-precision geometric constraint information for LEO platforms.

[0026] Under relevant technologies, the solar radiation pressure model for low-Earth orbit (LEO) satellites is mainly constructed based on analytical methods. Specifically, it describes the solar radiation pressure experienced by LEO satellites from three aspects: radiation source, propagation path, and surface reflection and absorption of the LEO satellite. Furthermore, this LEO satellite solar radiation pressure model considers the geometric relationship between the Sun, Earth, and the LEO satellite to more clearly describe the dynamic changes in solar radiation pressure acceleration experienced by the LEO satellite during its orbital operation.

[0027] The aforementioned method constructs a solar radiation pressure model for low-Earth orbit (LEO) satellites using reflection, absorption, and diffuse reflection coefficients to describe the interaction between visible and infrared radiation and the LEO satellite surface. This method relies on prior information parameters of the LEO satellite. Furthermore, the method assumes that the LEO satellite surface characteristics are uniform and stable, and considers the influence of solar activity and rotation period on solar radiation intensity.

[0028] However, due to the complex structure and dynamic changes of the actual low-Earth orbit satellite surface, as well as the complex changes in solar radiation intensity, the accuracy of the low-Earth orbit satellite solar radiation pressure model constructed using the above method is limited to a certain extent, thus affecting the normal operation of low-Earth orbit satellites.

[0029] To address the aforementioned technical problems, in this embodiment, during the first correction period, the total perturbation acceleration of the low-Earth orbit (LEO) satellite is calculated, environmental state information of the LEO satellite is collected, and solar radiation pressure acceleration is obtained based on the total perturbation acceleration and environmental state information. Based on the initial radiation coefficient used in the initial LEO satellite solar radiation pressure model, and combined with the solar radiation pressure acceleration, a corrected radiation coefficient is obtained. The corrected radiation coefficient is then used to construct the LEO satellite solar radiation pressure model for the first correction period. This approach avoids the problem that the radiation coefficient of traditional LEO satellite solar radiation pressure models relies excessively on the prior parameters of the LEO satellite, thus failing to adapt to changes in the space environment during LEO satellite orbital operation.

[0030] The following is a brief introduction to the system architecture diagram applicable to the technical solutions of the embodiments of this application. It should be noted that the system architecture diagram described below is only used to illustrate the embodiments of this application and is not intended to limit the scope of the application.

[0031] For example, see Figure 1As shown, it is a system architecture diagram applicable to the embodiments of this application. The system architecture includes at least a low-orbit satellite 101, a sun 102, an earth 103, and an atmosphere 104.

[0032] Low Earth Orbit Satellite 101: As the core carrier of space exploration, Low Earth Orbit Satellite 101 is subjected to perturbation forces from the Sun 102, Earth 103 and Atmosphere 104, resulting in a total perturbation acceleration. Sun 102: exerts solar radiation pressure on low-Earth orbit satellite 101, causing the low-Earth orbit satellite to experience solar radiation pressure acceleration; Earth 103: exerts a radiating force on low-Earth orbit satellite 101, causing the low-Earth orbit satellite to experience Earth radiation acceleration. Atmosphere 104: Generates atmospheric drag on low-Earth orbit satellite 101, causing atmospheric drag acceleration in the low-Earth orbit satellite.

[0033] In addition, based on the real-time statistical data of the total acceleration of the perturbation force generated by the Sun (102), Earth (103), and Atmosphere (104) on the low-Earth orbit satellite, and combined with environmental state parameters, the low-Earth orbit satellite 101 calculates the corresponding corrected radiation coefficient and updates it to the initial low-Earth orbit satellite solar radiation pressure model, thereby achieving accurate construction of the low-Earth orbit satellite solar radiation pressure model.

[0034] For example, see Figure 2 As shown in the embodiments of this application, the specific process for constructing a low-orbit satellite solar radiation pressure model is as follows: Step 200: During the first correction period, calculate the total perturbation acceleration of the low-orbit satellite.

[0035] For details, please refer to Figure 3 As shown in the embodiments of this application, the total acceleration of the perturbation force 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, which is the force generated by the low-orbit satellite based on the celestial body and space environment.

[0036] In practical applications, the perturbation force mainly includes the solar radiation pressure from the sun reaching the satellite surface, the Earth's radiation pressure consisting of short-wave radiation reflected from the Earth's surface and long-wave radiation generated by the satellite itself, and atmospheric drag generated by the Earth's upper atmosphere. Correspondingly, the acceleration generated by solar radiation pressure on the surface of a low-Earth orbit satellite is called solar radiation pressure acceleration, the acceleration generated by Earth's atmospheric radiation pressure on the surface of a low-Earth orbit satellite is called Earth's radiation pressure acceleration, and the acceleration generated by atmospheric drag on the surface of a low-Earth orbit satellite is called atmospheric drag acceleration.

[0037] Therefore, the so-called total perturbation acceleration is the result of the fusion of solar radiation pressure acceleration, Earth's radiation pressure acceleration, and atmospheric drag acceleration. It is a holistic descriptive parameter and does not specifically distinguish the magnitudes of the individual parameters of solar radiation pressure acceleration, Earth's radiation pressure acceleration, and atmospheric drag acceleration.

[0038] For example: let solar radiation pressure acceleration be denoted as The Earth's radiation pressure acceleration is denoted as... The atmospheric drag acceleration is denoted as ,So, , This refers to the total acceleration of the perturbation force collected. By analyzing and considering the perturbation forces generated by the Sun, Earth, and atmosphere on the orbital motion of low-Earth orbit (LEO) satellites, we can more accurately reflect the force situation of LEO satellites in complex space environments, providing reliable data support for the construction of LEO satellite solar radiation pressure models.

[0039] 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: 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.

[0040] 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.

[0041] 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:

[0042] 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.

[0043] 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.

[0044] 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.

[0045] 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.

[0046] 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.

[0047] Furthermore, in order to more accurately reflect the actual force situation of low-Earth orbit satellites during orbital operation, it is necessary to perform offset correction on the total perturbation acceleration of low-Earth orbit satellites.

[0048] Specifically, in this embodiment of the application, since the onboard accelerometer has its own system error, it is necessary to perform overall offset correction on the total acceleration of the perturbation force according to the system's preset correction parameter set to obtain the corrected total acceleration of the perturbation force.

[0049] In practical applications, the accelerometer calibration parameter set includes bias correction parameters and scale correction parameters, which are used to correct the total perturbation acceleration caused by the system's own errors in the spaceborne accelerometer. Existing deviations.

[0050] For example, the deviation correction parameter is denoted as The scale correction parameter is denoted as The corrected total acceleration of the perturbation force .

[0051] By removing the calculated Earth radiation pressure and atmospheric drag acceleration from the total perturbation acceleration, a precise solar radiation pressure acceleration can be obtained, providing key data support for calculating the radiation coefficient correction based on error mapping and constructing a precise low-orbit satellite solar radiation pressure model.

[0052] Step 201: Collect environmental status information of low-orbit satellites.

[0053] For details, please refer to Figure 6 As shown in this embodiment, environmental state information of a low-Earth orbit (LEO) satellite can be collected via satellite payload sensors. This environmental state information is used to describe the LEO satellite's orbital operating status and the surrounding operating space environment. The orbital operating parameters include at least one of the following: the LEO satellite's attitude parameters, structural parameters, and surface reflection parameters. The operating space environment parameters include at least one of the following: solar radiation intensity and atmospheric density.

[0054] For example, the angle between the incident sunlight and the illuminated panel of a low-orbit satellite. The direction cosine is used as the attitude parameter of the low-Earth orbit satellite, the area A and mass m of the low-Earth orbit satellite's surface panel are used as the structural parameters of the low-Earth orbit satellite, and the specular reflection parameters experienced by the low-Earth orbit satellite are used as the parameters. Diffuse reflection parameters and absorption parameters These parameters, which are surface reflection parameters of low-Earth orbit satellites, together constitute the orbital operation parameters of low-Earth orbit satellites.

[0055] In addition, the solar flux reaching the surface of low-Earth orbit satellites. As a measure of solar radiation intensity, the atmospheric density of the low-Earth orbit satellite operating space is denoted as... These parameters collectively constitute the operational space environment parameters of a low-Earth orbit (LEO) satellite. The orbital operational status parameters and operational space environment parameters are used as the environmental status information of the LEO satellite. Table 1 illustrates the main contents and relationships of the environmental status information and related parameters of the LEO satellite during the first calibration cycle. Table 1

[0056] By acquiring the environmental state parameters of low-Earth orbit (LEO) satellites, we can gain a comprehensive understanding of their orbital operation and the surrounding space environment. This provides further data support for the construction of LEO satellite solar radiation pressure models and enables the construction of LEO satellite solar radiation pressure models based on environmental state parameters and total perturbation acceleration data to have strong adaptability in complex space environments.

[0057] Step 202: Based on the total acceleration of the perturbation force and the environmental state information, obtain the corresponding solar radiation pressure acceleration.

[0058] Specifically, in this embodiment, since the total perturbation acceleration is a holistic descriptive parameter and does not specifically distinguish the magnitudes of solar radiation pressure acceleration, Earth radiation pressure acceleration, and atmospheric drag acceleration, the Earth radiation pressure acceleration and atmospheric drag acceleration experienced by the low-Earth orbit satellite are calculated separately under the low-Earth orbit satellite orbital operating environment based on the orbital operation state parameters and the operating space environment parameters. Then, the Earth radiation pressure acceleration and atmospheric drag acceleration are removed from the total perturbation acceleration to obtain the solar radiation pressure acceleration of the low-Earth orbit satellite.

[0059] For example, the Earth's radiation pressure acceleration experienced by a low-Earth orbit satellite under the operating environment during the first correction cycle is denoted as... Atmospheric drag acceleration is denoted as Then, the solar radiation pressure acceleration experienced by the low-Earth orbit satellite under the low-Earth orbit satellite operating environment during the first correction period is: .

[0060] By using orbital operating state parameters and operating space parameters to calculate Earth's radiation pressure acceleration and atmospheric drag acceleration respectively, it is ensured that solar radiation pressure acceleration can be accurately separated from the total perturbation acceleration described holistically, thus ensuring the accuracy of the obtained solar radiation pressure acceleration.

[0061] Step 203: Based on the initial radiation coefficient used in the initial low-orbit satellite solar radiation pressure model, and combined with solar radiation pressure acceleration, obtain the corrected radiation coefficient.

[0062] For details, please refer to Figure 7 As shown, step 203 may include the following operations: Step 203-1: Obtain the initial solar radiation pressure acceleration corresponding to the initial low-orbit satellite solar radiation pressure model, and establish the error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration.

[0063] Specifically, in this embodiment, since changes in the radiation coefficient lead to changes in solar radiation pressure acceleration, and these changes can be approximated by partial derivatives (first-order Taylor expansion), an error mapping relationship can be established between the actual solar radiation pressure acceleration collected by the spaceborne accelerometer and the initial solar radiation pressure acceleration. Before constructing the error function relationship, it is necessary to derive the partial derivative expression of the radiation coefficient.

[0064] For example, based on the known flight attitude, macroscopic model (geometry), and optical properties (reflection and absorption) of a low-Earth orbit satellite, the solar radiation pressure acceleration experienced by a single panel on the surface of the low-Earth orbit satellite can be obtained using formula (1): (1) In the formula, It is the ratio of the area (in square meters) to the mass (in kilograms) of the panel material. The speed of light is 299,792,458 m / s. It is a parameter of the light pressure scale. The angle between the incident ray and the normal of the panel. and It is the emissivity in the visible and infrared bands. It is the ground shadow function. It is the radiation intensity (solar flux) of the sun reaching the surface of the low-Earth orbit satellite. The total solar irradiance of the low-Earth orbit satellite in near-Earth orbit can be characterized by the solar constant, which can be obtained by formula (2).

[0065] (2)

[0066] In the formula, One astronomical unit (AU) represents the average distance from the Sun to the Earth, with a value of 1.495978707. rice, The instantaneous distance from the low-orbit satellite to the sun. The solar radiation pressure on the surface of the low-orbit satellite can be quantitatively described by modeling it into three parts: specular reflection, diffuse reflection, and absorption. The radiation coefficients of visible light and infrared bands can be calculated according to formula (3).

[0067] (3)

[0068] 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.

[0069] 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).

[0070] (4)

[0071] Formula (4) above describes the partial derivative of the panel radiation coefficient of the low-orbit satellite body and the solar sail.

[0072] 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.

[0073] 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 accelerations.

[0074] Onboard accelerometer observation data at i epochs Define residual The theoretical total acceleration of the perturbation force, including corrections, and the actual observed values. The difference, i.e. , The physical meaning is the sum of the a priori theoretical perturbation acceleration and the actual observed value without correction terms. The deviation. Using , and By simplifying the partial derivative symbols, we finally obtain the error equation for the i-th epoch. When the onboard accelerometer has n sets of observation data ( When ), n error equations with consistent form can be obtained.

[0075] Step 203-2: Based on the error mapping relationship, and combining the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient, obtain the correction amount of the radiation coefficient.

[0076] Specifically, in the embodiments of this application, the core method for solving the parameter to be estimated (the correction number of the radiation coefficient) is the least squares method, which finds the optimal solution by minimizing the sum of squared residuals to obtain the parameter to be estimated.

[0077] For example, when the onboard accelerometer has n sets of observation data ( When this is the case, the error equations for all epochs can be integrated into matrix form. Definition

[0078] The matrix form of the overall error equation is: The core of the least squares method is to minimize the sum of squares of the residual vectors (i.e., to minimize the overall deviation between the calculated and observed values). The objective function is defined as follows: ,Will Substituting into the objective function yields In order to make Minimum, needs to be Take the derivative and set it to zero. Through matrix differentiation, the normal equation (a necessary condition for the optimal solution of least squares) can be derived as follows: Then, solving this equation yields the optimal vector of parameters to be estimated. = .

[0079] Step 203-3: Add the correction amount of the radiation coefficient to the initial radiation coefficient to obtain the corrected radiation coefficient.

[0080] Specifically, in the embodiments of this application, the correction amount of the radiation coefficient is obtained, and the initial radiation coefficient is known. Then, by superimposing the correction amount of the radiation coefficient with the initial radiation coefficient, the optimal radiation coefficient can be obtained, that is, the corrected radiation coefficient.

[0081] For example, when the correction amount of the solved radiation coefficient is = Then, assuming the initial radiation coefficient is , and Then the corrected radiation coefficient can be obtained as follows: , and This process essentially involves using observational data to deduce parameter corrections, ensuring that the solar radiation pressure acceleration calculated by the model closely matches the actual observed value, thereby achieving an accurate estimate of the radiation coefficient.

[0082] By linearizing the nonlinear relationship of solar radiation pressure acceleration around the initial radiation coefficient using Taylor expansion, an error mapping is established, and the radiation coefficient correction is solved using the least squares method. Then, the correction is superimposed with the initial radiation coefficient, which transforms the nonlinear relationship into a solvable linear problem. By back-calculating the parameter correction using observational data, the solar radiation pressure acceleration calculated by the model closely matches actual observations, achieving accurate estimation of the radiation coefficient and providing reliable parameter support for the calibration of low-orbit satellite solar radiation pressure models.

[0083] Step 204: Construct a low-orbit satellite solar radiation pressure model for the first correction period using the corrected radiation coefficient.

[0084] Specifically, in this embodiment, the correction amount of the radiation coefficient has been solved by the least squares method, and based on the correction amount of the radiation coefficient and the initial radiation coefficient, the corrected radiation coefficient is obtained. Then, an accurate low-orbit satellite solar radiation pressure model can be constructed based on the corrected radiation coefficient.

[0085] By estimating the corrected radiation coefficient of the low-Earth orbit satellite using the least squares method to determine the actual forces acting on the satellite, a solar radiation pressure model for the low-Earth orbit satellite adapted to the orbital environment of the first correction cycle can be constructed, providing accurate dynamic model support for precise orbit calculation and high-precision forecasting of low-Earth orbit satellites.

[0086] Based on the same inventive concept, this application also provides a device for constructing a low-orbit satellite solar radiation pressure model, for example, see [reference]. Figure 8 As shown, the data security transmission device 800 includes: Acquisition Module 8000: Used to calculate the total acceleration of the perturbation force of the low-Earth orbit satellite in the first correction period; the total acceleration of the perturbation force is used to describe the total acceleration of the low-Earth orbit satellite based on the perturbation force, which is the force generated by the low-Earth orbit satellite based on the celestial body and space environment; and to collect the environmental state information of the low-Earth orbit satellite, which is used to describe the orbital operation status of the low-Earth orbit satellite and the surrounding operating space environment. Processing module 8001: It is used to obtain solar radiation pressure acceleration based on the total acceleration of perturbation force and environmental state information, obtain the corrected radiation coefficient based on the initial radiation coefficient used in the initial low-orbit satellite solar radiation pressure model and the solar radiation pressure acceleration, and construct the low-orbit satellite solar radiation pressure model for the first correction period using the corrected radiation coefficient.

[0087] In one optional embodiment, the acquisition module 8000 is specifically used for: Calculating the total perturbation acceleration of a low-Earth orbit satellite includes: Acquire onboard GNSS data from low-Earth orbit satellites and eliminate ionospheric errors in the onboard GNSS data to obtain ionospheric-free onboard GNSS data; ionospheric-free onboard GNSS data is used to describe the actual orbital operating parameters of low-Earth orbit satellites. 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. Establish the error mapping relationship between actual orbital operating parameters and theoretical orbital operating parameters; Based on the error mapping relationship, the dynamic model of the low-Earth orbit satellite is adjusted in the direction of error reduction; the dynamic model is a model used to describe the force situation of the low-Earth orbit satellite during its orbital operation, and the force situation includes various perturbations. Based on the adjusted dynamic model, the total perturbation acceleration of the low-orbit satellite is obtained.

[0088] In an optional embodiment, the acquisition module 8000 is further specifically used for: Calculating the total perturbation acceleration of a low-Earth orbit satellite includes: Obtain the initial total perturbation acceleration of the low-Earth orbit satellite detected by the onboard accelerometer; Based on the mapping relationship between the fixed coordinate system and the orbital coordinate system of the low-Earth orbit satellite, the attitude transformation matrix of the low-Earth orbit satellite is constructed. Based on the attitude transformation matrix, the initial total perturbation acceleration is transformed from the fixed coordinate system of the low-Earth orbit satellite to the orbital coordinate system of the low-Earth orbit satellite to obtain the total perturbation acceleration of the low-Earth orbit satellite.

[0089] In an optional embodiment, the processing module 8001 is specifically used for: The environmental status information includes track operation status parameters and operating space environment parameters; Based on the total perturbation acceleration and environmental state information, the corresponding solar radiation pressure acceleration is obtained, including: Based on orbital operation parameters and operational space environment parameters, the Earth's radiation pressure acceleration and atmospheric drag acceleration experienced by the low-Earth orbit satellite under the low-Earth orbit operational environment are calculated respectively. By removing the Earth's radiation pressure acceleration and atmospheric drag acceleration from the total perturbation acceleration, the solar radiation pressure acceleration of the low-orbit satellite can be obtained.

[0090] In an optional embodiment, the processing module 8001 is further configured to: Based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model, and combined with solar radiation pressure acceleration, the corrected radiation coefficient is obtained, including: Obtain the initial solar radiation pressure acceleration corresponding to the initial low-Earth orbit satellite solar radiation pressure model, and establish the error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration; Based on the error mapping relationship, and combined with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient, the correction amount of the radiation coefficient is obtained; The corrected emissivity is obtained by superimposing the emissivity correction factor on the initial emissivity.

[0091] The acquisition module obtains the total perturbation acceleration and environmental status information of the low-Earth orbit satellite. The processing module separates the Earth's radiation pressure and atmospheric drag acceleration from the total perturbation acceleration to obtain the accurate solar radiation pressure acceleration. Then, by establishing an error mapping, the corrected radiation coefficient is obtained. This model can be adapted to different hardware configurations of low-Earth orbit satellites, ensuring the accuracy of solar radiation pressure-related data. Finally, a low-Earth orbit satellite solar radiation pressure model that fits the actual orbital environment is constructed.

[0092] Based on the same inventive concept, this application provides a computer device that can realize the function of the aforementioned low-orbit satellite solar radiation pressure model construction device. (Refer to...) Figure 9 Electronic devices include: At least one processor 9001 and a memory 9002 connected to at least one processor 9001. In this embodiment, the specific connection medium between the processor 9001 and the memory 9002 is not limited. Figure 9 The example shown is the connection between processor 9001 and memory 9002 via bus 9000. Bus 9000... Figure 9The connections between other components are shown in thick lines only and are not intended to be limiting. The Bus 9000 can be divided into address bus, data bus, control bus, etc., for ease of representation. Figure 9 The character is represented by a single thick line, but this does not indicate that there is only one bus or one type of bus. Alternatively, the processor 9001 can also be called a controller; there is no restriction on the name.

[0093] In this embodiment, the memory 9002 stores instructions executable by at least one processor 9001. By executing the instructions stored in the memory 9002, the at least one processor 9001 can perform the data update method described above. The processor 9001 can implement... Figure 8 The functions of each module in the device shown.

[0094] The processor 9001 is the control center of the device. It can connect to various parts of the control device through various interfaces and lines. By running or executing instructions stored in memory 9002 and calling data stored in memory 9002, the processor can perform various functions and process data, thereby monitoring the device as a whole.

[0095] In one possible design, processor 9001 may include one or more processing units. Processor 9001 may integrate an application processor and a modem processor, wherein the application processor mainly handles the operating system, user interface, and applications, and the modem processor mainly handles wireless communication. It is understood that the modem processor may also not be integrated into processor 9001. In some embodiments, processor 9001 and memory 9002 may be implemented on the same chip; in some embodiments, they may also be implemented on separate chips.

[0096] The processor 9001 can be a general-purpose processor, such as a central processing unit (CPU), digital signal processor, application-specific integrated circuit, field-programmable gate array or other programmable logic device, discrete gate or transistor logic device, or discrete hardware component, capable of implementing or executing the methods, steps, and logic block diagrams disclosed in the embodiments of this application. The general-purpose processor can be a microprocessor or any conventional processor. The steps of the method for constructing a low-Earth orbit satellite solar radiation pressure model disclosed in the embodiments of this application can be directly manifested as execution by a hardware processor, or execution by a combination of hardware and software modules within the processor.

[0097] Memory 9002, as a non-volatile computer-readable storage medium, can be used to store non-volatile software programs, non-volatile computer-executable programs, and modules. Memory 9002 may 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. Memory 9002 can be any other medium capable of carrying or storing desired program code in the form of instructions or data structures that can be accessed by a computer, but is not limited thereto. In the embodiments of this application, memory 9002 can also be a circuit or any other device capable of implementing storage functions for storing program instructions and / or data.

[0098] By designing and programming the processor 9001, the code corresponding to the method for constructing the low-orbit satellite solar radiation pressure model described in the aforementioned embodiments can be embedded into the chip, thereby enabling the chip to execute the code during operation. Figure 2 The steps of constructing a low-Earth orbit satellite solar radiation pressure model according to the illustrated embodiment are as follows. How to design and program the processor 9001 is a technique well-known to those skilled in the art and will not be described further here.

[0099] Based on the same inventive concept, embodiments of this application provide a computer-readable storage medium storing a computer program executable by a computer device. When the program is run on the computer device, it causes the computer device to perform the steps of the above-described road network data detection method.

[0100] Those skilled in the art will understand that embodiments of this application can be provided as methods, systems, or computer program products. Therefore, this application can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, this application can take the form of a computer program product embodied on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0101] This application is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to this application. It should be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations 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 processor, 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, generate instructions for implementing the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0102] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0103] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0104] Obviously, those skilled in the art can make various modifications and variations to this application without departing from the spirit and scope of this application. Therefore, if such modifications and variations fall within the scope of the claims of this application and their equivalents, this application also intends to include such modifications and variations.

Claims

1. A method for constructing a solar radiation pressure model for a low-Earth orbit satellite, characterized in that, include: During the first correction period, calculate the total perturbation acceleration of the low-Earth orbit satellite; The total acceleration due to perturbation is used to describe the total acceleration generated by the low-Earth orbit satellite based on perturbation; the perturbation includes the dynamics generated by the low-Earth orbit satellite based on celestial bodies and the space environment; Collect environmental status information of the low-orbit satellite; The environmental status information is used to describe the orbital operation status of the low-orbit satellite and the surrounding operating space environment; Based on the total acceleration of the perturbation force and the environmental state information, the solar radiation pressure acceleration is obtained; Based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model, and combined with the solar radiation pressure acceleration, the corrected radiation coefficient is obtained; The corrected radiation coefficient is used to construct a low-orbit satellite solar radiation pressure model for the first correction period.

2. The method as described in claim 1, characterized in that, The calculation of the total perturbation acceleration of low-Earth orbit satellites includes: 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; the ionospheric-free onboard GNSS data is used to describe the actual orbital operating parameters of the low-Earth orbit satellite. Based on the Earth's gravity field model and the initial orbit of the low-Earth orbit satellite, the theoretical orbital parameters of the low-Earth orbit satellite are obtained. Establish the error mapping relationship between the actual orbital operating parameters and the theoretical orbital operating parameters; Based on the error mapping relationship, the dynamic model of the low-Earth orbit satellite is adjusted in the direction of error reduction; the dynamic model is a model used to describe the force situation of the low-Earth orbit satellite during orbit operation, and the force situation includes various perturbation forces; Based on the adjusted dynamic model, the total perturbation acceleration of the low-orbit satellite is obtained.

3. The method as described in claim 1, characterized in that, The calculation of the total perturbation acceleration of low-Earth orbit satellites includes: Obtain the initial total perturbation acceleration of the low-orbit satellite as collected by the onboard accelerometer; Based on the mapping relationship between the fixed coordinate system of the low-Earth orbit satellite and the orbital coordinate system of the low-Earth orbit satellite, the attitude transformation matrix of the low-Earth orbit satellite is constructed. Based on the attitude transformation matrix, the initial total perturbation acceleration is transformed from the fixed coordinate system of the low-Earth orbit satellite to the orbital coordinate system of the low-Earth orbit satellite to obtain the total perturbation acceleration of the low-Earth orbit satellite.

4. The method as described in claim 1, characterized in that, The total perturbation acceleration is generated based on at least one of the following accelerations: solar radiation pressure acceleration, Earth's radiation pressure acceleration, and atmospheric drag acceleration. The solar radiation pressure acceleration is used to describe the acceleration generated by the low-orbit satellite under the force of solar radiation pressure. The Earth radiation pressure acceleration is used to describe the acceleration generated by the Earth's own infrared radiation pressure and reflected solar radiation pressure on the low-orbit satellite. Atmospheric drag acceleration: used to describe the acceleration caused by atmospheric drag experienced by the low-orbit satellite.

5. The method according to any one of claims 1-4, characterized in that, The environmental status information includes track operation status parameters and operating space environment parameters; The process of obtaining the corresponding solar radiation pressure acceleration based on the total perturbation acceleration and the environmental state information includes: Based on the orbital operation status parameters and the operating space environment parameters, the Earth's radiation pressure acceleration and atmospheric drag acceleration experienced by the low-Earth orbit satellite under the orbital operation environment in the first correction period are calculated respectively. The solar radiation pressure acceleration of the low-orbit satellite is obtained by removing the Earth's radiation pressure acceleration and the atmospheric drag acceleration from the total perturbation acceleration.

6. The method according to any one of claims 1-4, characterized in that, Based on the initial radiation coefficient used in the initial low-Earth orbit satellite solar radiation pressure model, and combined with the solar radiation pressure acceleration, the corrected radiation coefficient is obtained, including: Obtain the initial solar radiation pressure acceleration corresponding to the initial low-orbit satellite solar radiation pressure model, and establish the error mapping relationship between the solar radiation pressure acceleration and the initial solar radiation pressure acceleration; Based on the aforementioned error mapping relationship, and combined with the initial low-orbit satellite solar radiation pressure model and the corresponding initial radiation coefficient, the correction amount of the radiation coefficient is obtained; The corrected radiation coefficient is obtained by superimposing the radiation coefficient correction amount with the initial radiation coefficient.

7. A device for constructing a low-orbit satellite solar radiation pressure model, characterized in that, The device includes: Acquisition module: used to acquire the total acceleration of the perturbation force of the low-orbit satellite during the first correction period. The total acceleration of the perturbation force is used to describe the total acceleration generated by the low-orbit satellite based on the perturbation force, which is the force generated by the low-orbit satellite based on the celestial body and space environment. Also, to acquire the environmental state information of the low-orbit satellite. The environmental state information is used to describe the orbital operation state of the low-orbit satellite and the surrounding operating space environment. Processing module: used to obtain the corresponding solar radiation pressure acceleration based on the total acceleration of the perturbation force and the environmental state information; to obtain the corrected radiation coefficient based on the initial radiation coefficient used in the initial low-orbit satellite solar radiation pressure model and the solar radiation pressure acceleration; and to construct the low-orbit satellite solar radiation pressure model for the first correction period using the corrected radiation coefficient.

8. A computer device, characterized in that, include: Memory, used to store computer programs; The processor, when executing the computer program stored in the memory, implements the method steps of any one of claims 1 to 6.

9. A computer-readable storage medium, characterized in that, It stores a computer program executable by a computer device, which, when run on the computer device, causes the computer device to perform the steps of the method according to any one of claims 1 to 6.

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