Low earth orbit satellite thermal radiation pressure modeling method and system based on heat transfer mechanism

Abstract

The invention discloses a low earth orbit satellite thermal radiation pressure modeling method based on a heat transfer mechanism. The method comprises the steps that panel optical parameters and thermodynamic parameters are determined; calculating panel radiation absorption power according to the panel optical parameters; determining the initial temperature of the panel according to the optical parameters and the thermodynamic parameters of the panel in combination with the radiation absorption power of the panel; according to the panel initial temperature, the panel temperature of a subsequent epoch is updated through a numerical integration method; and determining the thermal radiation pressure of the low-orbit satellite according to the updated panel temperature. Based on a panel-satellite body heat transfer mechanism, low-orbit satellite thermal radiation pressure modeling including panel initial temperature determination and temperature dynamic updating is realized.

Description

Technical Field

[0001] This invention belongs to the field of aerospace, and in particular relates to a method and system for modeling thermal radiation pressure of low-orbit satellites based on heat transfer mechanisms. Background Technology

[0002] After absorbing solar and terrestrial radiation, low-Earth orbit (LEO) satellites experience a rise in surface temperature and release thermal radiation. The reaction force generated by these thermal radiation photons on the satellite surface is called thermal radiation pressure, or thermal re-radiation pressure. Among all radiation pressure-related perturbations for LEO satellites, thermal radiation pressure perturbation is second only to solar radiation pressure. Researching and refining the thermal radiation pressure model for LEO satellites is not only crucial for reducing non-conservative force errors in precise mechanical models and improving satellite orbit determination and prediction accuracy, but it will also provide significant assistance in areas such as upper-level lateral wind field inversion and LEO satellite onboard accelerometer calibration.

[0003] In existing analytical modeling of thermal radiation pressure for low-Earth orbit (LEO) satellites, it is typically assumed that the satellite surface is in a state of constant thermal equilibrium, and the absorbed visible and infrared radiation energy is instantly re-radiated in the form of thermal radiation (Hackel et al., 2017). A study by Vielberg et al. (2020) on nonconservative force modeling for the GRACE satellite showed that thermal radiation pressure is the second largest radiation pressure after light pressure and cannot be ignored in the accurate modeling of nonconservative forces for LEO satellites. Wang et al. (2023), in their simplified dynamic orbit determination study of GRACE-FO, found that the static radiation pressure model based on instantaneous re-radiation cannot accurately reflect the thermal state of the satellite surface, leading to biases in the estimation of the light pressure scale factor. Wöske et al. (2018), in their onboard accelerometer calibration study, considered the influence of the satellite panel's thermal inertia and used a transient thermal radiation pressure model instead of a static thermal radiation pressure model to improve calibration accuracy. This transient thermal radiation pressure model has also been adopted by the GROOPS software to mitigate the impact of nonconservative force modeling errors in gravity field recovery studies (Mayer-Gürr et al., 2021). Siemes et al. (2023) investigated the impact of heat transfer between the panel and satellite on thermal radiation pressure modeling and found that the initial temperature error could affect thermal radiation pressure modeling for up to 6 hours under the heat transfer mechanism.

[0004] In summary, existing analytical models of thermal radiation pressure mostly use static thermal radiation pressure modeling, while transient thermal radiation models often treat the satellite panel as thermal insulator. There is relatively little research on the impact of the thermal conduction mechanism between the satellite panel and the star on the transient thermal radiation pressure model. Summary of the Invention

[0005] To address the problems existing in the modeling of thermal radiation pressure of low-Earth orbit (LEO) satellites, this invention provides a method for modeling LEO satellite thermal radiation pressure based on a heat transfer mechanism. This method, based on the heat transfer mechanism between the panel and the satellite body, realizes the modeling of LEO satellite thermal radiation pressure, including the determination of the initial temperature of the panel and dynamic temperature updates.

[0006] According to one aspect of the present invention, a method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms is provided, comprising:

[0007] Determine the optical and thermodynamic parameters of the panel;

[0008] The panel's radiative absorption power is calculated based on the panel's optical parameters.

[0009] Based on the panel's optical and thermodynamic parameters, and in conjunction with the panel's radiative absorption power, the initial temperature of the panel is determined.

[0010] Based on the initial panel temperature, the panel temperature in subsequent epochs is updated using a numerical integration method.

[0011] The thermal radiation pressure of the low-orbit satellite is determined based on the updated panel temperature.

[0012] As a further technical solution, the optical and thermodynamic parameters of the panel are determined, including:

[0013] Before the launch of the low-orbit satellite, the optical and thermodynamic parameters of each panel are obtained through calibration in a ground laboratory.

[0014] As a further technical solution, the panel radiation absorption power is calculated based on the panel's optical parameters, including:

[0015] The radiation absorption power of the panel is calculated by calculating the portion of solar and terrestrial radiation energy absorbed by the panel in the multi-panel radiation pressure model of a low-orbit satellite.

[0016] As a further technical solution, in the dynamic radiation pressure model, if the time span between two consecutive epochs is too long, the initial temperature of the panel is redefined.

[0017] As a further technical solution, when determining the initial temperature of the panel, it is also necessary to assume that the panel surface is in thermodynamic equilibrium, that is, the panel's radiation absorption power, thermal radiation release power and heat transfer power are in balance.

[0018] As a further technical solution, the panel temperature in subsequent epochs is updated using a numerical integration method, including: , in, This represents the current epoch satellite panel surface temperature; The panel temperature of the previous epoch. This represents the temperature change between the two integral quantization intervals before and after the panel. The number of subintervals into which the numerical integral is divided. The sub-interval index; For the first The radiative absorption power of the panel in each integrator sub-interval; For the first The thermal radiation power of the panel in each integral sub-interval; For the first The heat transfer power between the panel of each integrator interval and the satellite body; The specific heat capacity of the panel determines the temperature change of the panel caused by the absorption, release, or transfer of heat.

[0019] According to one aspect of the present invention, a low-Earth orbit satellite thermal radiation pressure modeling system based on heat transfer mechanisms is provided, comprising:

[0020] The first main module is used to determine the optical and thermodynamic parameters of the panel.

[0021] The second main module is used to calculate the panel's radiation absorption power based on the panel's optical parameters.

[0022] The third main module is used to determine the initial temperature of the panel based on the panel's optical and thermodynamic parameters, combined with the panel's radiation absorption power.

[0023] The fourth main module is used to update the panel temperature in subsequent epochs based on the initial panel temperature using a numerical integration method.

[0024] The fifth main module is used to determine the thermal radiation pressure of low-orbit satellites based on the updated panel temperature.

[0025] According to one aspect of the present invention, a low-Earth orbit satellite thermal radiation pressure modeling device based on a heat transfer mechanism is provided, comprising a memory and a processor, wherein the memory stores program instructions that are executed by the processor, and the processor invokes the program instructions to execute the low-Earth orbit satellite thermal radiation pressure modeling method based on a heat transfer mechanism.

[0026] According to one aspect of the present invention, a non-transitory computer-readable storage medium is provided, the non-transitory computer-readable storage medium storing computer instructions that cause the computer to execute the described method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms.

[0027] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0028] This invention proposes a method for modeling the thermal radiation pressure of low-Earth orbit (LEO) satellites, based on the heat transfer mechanism between the satellite panel and the satellite body. This method involves determining the initial temperature of the panel and dynamically updating the temperature. Compared with traditional static thermal radiation pressure modeling methods, the temperature change of the LEO satellite panel obtained by this method is more gradual, and the dynamic change of the panel temperature is more consistent with the physical characteristics of thermodynamic changes. Moreover, the improvement effect of this invention on the thermal radiation pressure modeling of LEO satellites is particularly significant during periods when LEO satellites frequently enter and exit the Earth's shadow. Attached Figure Description

[0029] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the accompanying drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the accompanying drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0030] Figure 1 This is a flowchart illustrating a method for modeling the thermal radiation pressure of low-orbit satellites based on a heat transfer mechanism, provided in an embodiment of the present invention.

[0031] Figure 2 (a)-(d) are schematic diagrams comparing the panel temperatures of each panel of the GRACE-C satellite in 2019 for 090 days and 180 days, obtained by the traditional static thermal radiation pressure model and the method provided in the embodiments of the present invention. Detailed Implementation

[0032] The terms “comprising” and “having”, and any variations thereof, in the specification, claims, and accompanying drawings of this invention are intended to cover a non-exclusive inclusion, such as a process, method, system, product, or apparatus that includes a series of steps or units, not necessarily limited to those explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0033] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention. In addition, the technical features of the various embodiments or individual embodiments provided by the present invention can be arbitrarily combined to form new technical solutions. Such combinations are not bound by the order of steps and / or structural composition patterns, but must be based on the ability of those skilled in the art to implement them. When the combination of technical solutions is contradictory or cannot be implemented, it should be considered that such a combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.

[0034] Existing analytical models of thermal radiation pressure mostly employ static thermal radiation pressure modeling, while transient thermal radiation models often treat the satellite panel as a thermal insulator. Research on the impact of the heat conduction mechanism between the satellite panel and the satellite body on transient thermal radiation pressure models is limited. Based on this, this invention proposes a method for modeling the thermal radiation pressure of low-Earth orbit satellites, including initial panel temperature determination and dynamic temperature updates, based on the heat transfer mechanism between the panel and the satellite body. Its overall technical approach is as follows: Figure 1 As shown, the specific steps include the following:

[0035] Step 1: Determine the optical and thermodynamic parameters of the panel.

[0036] The optical and thermodynamic parameters of a low-Earth orbit (LEO) satellite panel are a set of parameters used to describe the fixed physical properties of the satellite panel, including optical parameters such as visible light absorptivity and infrared light absorptivity, and thermodynamic parameters such as specific heat capacity and thermal conductivity. Visible and infrared light absorptivity affect the panel's absorption of external radiation and release of thermal radiation power; specific heat capacity directly affects the panel's temperature change rate; and thermal conductivity affects the heat transfer power between the panel and the satellite body. When constructing a thermal radiation pressure model for a LEO satellite, the above parameters for each panel must first be determined. These parameters are usually calibrated by a ground laboratory before satellite launch.

[0037] Step 1: Calculate the panel's radiative absorption power.

[0038] The radiation absorption power of the panel directly affects the panel temperature, which in turn affects the thermal radiation power of the satellite panel and the resulting thermal radiation pressure. Radiation absorption power This can be obtained by calculating the portion of solar and terrestrial radiation energy absorbed by the panels in the multi-panel solar radiation pressure model of low-Earth orbit satellites:

[0039]

[0040] in, Power absorbed by solar radiation. Earth's radiation absorption power is divided into ground albedo absorption power. With the power absorbed by Earth's thermal radiation Two parts; At the speed of light, is the solar flux coefficient, and the product of the two is the total solar irradiance at a distance of 1 AU from the sun; This is the shaded function, with a value range of 0-1; The light pressure scale factor; This represents the distance from the sun to a low-Earth orbit satellite. and These are the absorption rates of the panel for visible light and infrared light, respectively. and These are the ground surface element albedo and thermal infrared emissivity, respectively, which can be obtained by interpolation using ocean and land radiative flux data provided by the Clouds and the Earth's Radiant Energy System (CERES). The area of ​​the panel that absorbs radiation. The angle between the normal to the panel and the direction from the satellite to the sun. It is the angle between the normal of the panel and the direction from the satellite to the ground element.

[0041] Step 3: Determine the initial temperature of the panel.

[0042] Considering the dynamic process of the satellite panel's temperature rising due to radiation absorption, and then changing due to the release of thermal radiation or heat exchange with the satellite, the temperature change between subsequent epochs is determined through numerical integration after the initial surface temperature is established. However, in the dynamic radiation pressure model, if the time span between two consecutive epochs is too long, the initial surface temperature of the panel needs to be re-determined. (Initial surface temperature of the panel) The determination of the panel surface generally assumes thermodynamic equilibrium, that is, the panel's radiation absorption power, thermal radiation release power and heat transfer power are in balance.

[0043] Panel heat radiation release power This can be determined using the Stefan-Boltzmann law:

[0044]

[0045] in, This represents the area of ​​the panel; This is the infrared radiation emission coefficient of the panel, which is numerically the same as the infrared light absorption rate of the panel. The Stefan-Boltzmann constant is 5.670373. , This refers to the panel temperature.

[0046] Panel heat transfer power It is determined by the panel's thermal conductivity and the temperature difference between the panel and the satellite body:

[0047]

[0048] in, The thermal conductivity between the panel and the satellite body; For panel temperature, The initial temperature is the temperature of the satellite body, which is usually taken as the normal operating temperature of the electronic components of the low-orbit satellite, i.e., room temperature 298K.

[0049] Panel initial temperature The following equation must be satisfied:

[0050]

[0051] By rearranging the above equation, we can obtain information about The quadratic equation of :

[0052]

[0053] Solving the above quartic equation will yield the initial temperature of the panel:

[0054]

[0055]

[0056]

[0057]

[0058] Step 4: Panel temperature is dynamically updated.

[0059] The impact of thermal changes on the panel of a low-Earth orbit satellite on the rate of temperature change is related to the specific heat capacity of the panel.

[0060]

[0061] in, The specific heat capacity of the panel represents the amount of heat absorbed, released, or transferred by the panel. The degree of temperature change caused Its definition is as follows:

[0062]

[0063] Determining the initial temperature of the satellite panel Then, the panel temperature for subsequent epochs is updated using numerical integration:

[0064]

[0065] in, This represents the current epoch satellite panel surface temperature; The panel temperature of the previous epoch. This represents the temperature change between the two integral quantization intervals before and after the panel. The number of subintervals into which the numerical integral is divided. The sub-interval index; For the first The radiative absorption power of the panel in each integrator sub-interval; For the first The thermal radiation power of the panel in each integral sub-interval; For the first The heat transfer power between the integrator interval panel and the satellite body.

[0066] radiation Determined by interpolation of the panel absorption power between two consecutive epochs:

[0067]

[0068] in, The radiation absorption power of the satellite panel at the current epoch is calculated using the radiation pressure model. This represents the panel's radiation absorption power in the previous epoch.

[0069] thermal radiation power Determined by the panel surface temperature of the previous sub-interval:

[0070]

[0071] in, For the first The panel surface temperature of each sub-region; is the index of the integral subinterval.

[0072] The heat transfer power is determined by the temperature difference between the panel and the satellite body in the previous sub-section:

[0073]

[0074] in, For the first The satellite body temperature for each sub-interval was also obtained using numerical methods:

[0075]

[0076] in, The current satellite body temperature in the integral sub-interval; This represents the temperature change between the two sub-intervals. The thermal power of the satellite body in the current integral sub-interval is usually the thermal power of the electronic components provided by the low-Earth orbit satellite launcher. This represents the sum of the heat transfer power between each panel and the satellite body in the current integral sub-interval.

[0077] Step 5: Calculation of thermal radiation pressure of low-orbit satellites.

[0078] Once the temperature of the low-Earth orbit satellite panel is determined, the panel's thermal radiation power and the resulting thermal radiation pressure can be determined.

[0079]

[0080] in, The thermal radiation pressure exerted by the flat panel on the low-Earth orbit satellite is N; N is the number of panels on the low-Earth orbit satellite. For panel serial number, For the first Each panel emits heat radiation; It is the Stefan-Boltzmann constant; For the first The thermal radiation release power of each panel is calculated by combining the panel surface temperature with the Stefan-Boltzmann law. The speed of light; For the first The unit normal vector of each panel. For the first The area of ​​each panel, For the first Infrared radiation emission coefficient of each panel For the first The surface temperature of each panel.

[0081] Given the mass of low-Earth orbit satellites In this case, the perturbation acceleration generated by thermal radiation pressure can be further determined by Newton's second law. size:

[0082]

[0083] As an example, this invention uses the GRACE-FO satellite for experimental verification. The optical and thermodynamic parameters of the GRACE-FO satellite are shown in Table 1 below:

[0084] Table 1 Optical and thermodynamic parameters of the GRACE-FO satellite

[0085] The specific heat capacity of the GRACE-FO satellite is The satellite's thermal power is set to 200W as published on the World Meteorological Organization website.

[0086] Using both the traditional static thermal radiation pressure model and this method, the panel temperatures of each panel of the GRACE-C satellite on days 090 and 180 of 2019 were obtained as follows: Figure 2 As shown in (a)-(d), compared with the traditional static thermal radiation pressure model, the panel temperature change obtained by this method is more gradual, and the influence of the satellite's frequent entry and exit from the Earth's shadow on the panel surface temperature is effectively weakened (as shown in panel 8 in the figure), making the thermal radiation pressure model more consistent with the actual situation during the satellite's on-orbit operation. The improvement effect is more obvious during the 090-day period when the solar altitude angle on the orbital plane is relatively low, and the satellite's entry and exit from the Earth's shadow is more frequent. However, during the 180-day period when the solar altitude angle on the orbital plane is relatively high, the low-orbit satellite's single-sided panel is backlit for a long time (as shown in the figure). Figure 2 (c)-(d) Panel 1), the temperature modeling of the backlight panel in the traditional model has been too low for a long time, which will cause the non-conservative force modeling error in the normal direction of the satellite orbital plane. This can be significantly improved by introducing the heat transfer mechanism from the satellite body to the backlight panel.

[0087] The implementation of the various embodiments of the present invention is based on programmed processing through a device with processor functionality. Therefore, in practical engineering, the technical solutions and functions of the various embodiments of the present invention are encapsulated into various modules. Based on this reality, and building upon the above embodiments, the embodiments of the present invention provide a low-Earth orbit satellite thermal radiation pressure modeling system based on a heat transfer mechanism. This system is used to execute a low-Earth orbit satellite thermal radiation pressure modeling method based on a heat transfer mechanism from the above method embodiments.

[0088] The system includes: a first main module for determining the optical and thermodynamic parameters of the panel; a second main module for calculating the panel's radiative absorption power based on the panel's optical parameters; a third main module for determining the initial temperature of the panel based on the panel's optical and thermodynamic parameters, combined with the panel's radiative absorption power; a fourth main module for updating the panel temperature in subsequent epochs using a numerical integration method based on the initial panel temperature; and a fifth main module for determining the thermal radiation pressure of the low-Earth orbit satellite based on the updated panel temperature.

[0089] This invention provides a low-Earth orbit satellite thermal radiation pressure modeling system based on heat transfer mechanisms. Addressing the current situation where existing analytical models for thermal radiation pressure mostly employ static thermal radiation pressure modeling, and transient thermal radiation models often treat the satellite panel as thermal insulator, there is limited research on the impact of the heat conduction mechanism between the satellite panel and the satellite body on transient thermal radiation pressure models, this invention utilizes the aforementioned modules to achieve low-Earth orbit satellite thermal radiation pressure modeling based on the heat transfer mechanism between the panel and the satellite body, including determining the initial panel temperature and dynamically updating the temperature.

[0090] It should be noted that the system embodiments provided by the present invention are used not only to implement the methods in the above method embodiments, but also to implement the methods in other method embodiments provided by the present invention. The only difference is that corresponding functional modules are set. The principle is basically the same as that of the above system embodiments provided by the present invention. As long as those skilled in the art can improve the modules in the above system embodiments by referring to the specific technical solutions in other method embodiments and combining technical features to obtain corresponding technical means and technical solutions composed of these technical means, on the basis of the above system embodiments, and on the premise of ensuring the practicality of the technical solutions, they can obtain corresponding system-like embodiments for implementing the methods in other method-like embodiments.

[0091] Based on the same inventive concept as the foregoing embodiments, this embodiment of the invention also provides a low-Earth orbit satellite thermal radiation pressure modeling device based on a heat transfer mechanism, including a memory and a processor. The memory stores program instructions that are executed by the processor, and the processor calls the program instructions to execute the low-Earth orbit satellite thermal radiation pressure modeling method based on a heat transfer mechanism.

[0092] In embodiments of the present invention, the memory can be non-volatile memory, such as a hard disk drive (HDD) or a solid-state drive (SSD), or it can be volatile memory, such as random-access memory (RAM). Memory is any other medium capable of carrying or storing desired program code having an instruction or data structure form and accessible by a computer, but is not limited thereto. The memory in embodiments of the present invention can also be a circuit or any other device capable of implementing a storage function for storing program instructions and / or data.

[0093] In this embodiment of the invention, the processor may be a general-purpose processor, a digital signal processor, an application-specific integrated circuit (ASIC), a field-programmable gate array (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, or discrete hardware components, capable of implementing or executing the methods, steps, and logic block diagrams disclosed in this embodiment of the invention. The general-purpose processor may be a microprocessor or any conventional processor. The steps of the methods disclosed in this embodiment of the invention can be directly manifested as being executed by a hardware processor, or executed by a combination of hardware and software modules within the processor.

[0094] Based on the same inventive concept as the foregoing embodiments, this embodiment of the invention also provides a non-transitory computer-readable storage medium storing computer instructions. These computer instructions cause the computer to execute the following method for modeling the thermal radiation pressure of low-Earth orbit satellites based on a heat transfer mechanism:

[0095] Determine the optical and thermodynamic parameters of the panel;

[0096] The panel's radiative absorption power is calculated based on the panel's optical parameters.

[0097] Based on the panel's optical and thermodynamic parameters, and in conjunction with the panel's radiative absorption power, the initial temperature of the panel is determined.

[0098] Based on the initial panel temperature, the panel temperature in subsequent epochs is updated using a numerical integration method.

[0099] The thermal radiation pressure of the low-orbit satellite is determined based on the updated panel temperature.

[0100] In summary, based on the heat transfer mechanism between the satellite panel and the satellite body, this invention proposes a method for modeling the thermal radiation pressure of low-Earth orbit (LEO) satellites, including initial panel temperature determination and dynamic temperature updates. The LEO satellite panel temperature obtained by this method exhibits a more gradual temperature change compared to traditional static thermal radiation pressure modeling methods. The dynamic changes in panel temperature better conform to the physical characteristics of thermodynamic changes, and the improvement effect of this invention on LEO satellite thermal radiation pressure modeling is particularly significant during periods when LEO satellites frequently enter and exit the Earth's shadow.

[0101] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features therein; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the technical solutions of the embodiments of the present invention.

Claims

1. A method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms, characterized in that, include: Determine the optical and thermodynamic parameters of the panel; The panel's radiative absorption power is calculated based on the panel's optical parameters. Based on the panel's optical and thermodynamic parameters, and in conjunction with the panel's radiative absorption power, the initial temperature of the panel is determined. Based on the initial panel temperature, the panel temperature in subsequent epochs is updated using a numerical integration method. The thermal radiation pressure of the low-orbit satellite is determined based on the updated panel temperature.

2. The method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms according to claim 1, characterized in that, Determine the optical and thermodynamic parameters of the panel, including: Before the launch of the low-orbit satellite, the optical and thermodynamic parameters of each panel are obtained through calibration in a ground laboratory.

3. The method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms according to claim 1, characterized in that, Based on the optical parameters of the panel, the panel's radiative absorption power is calculated, including: The radiation absorption power of the panel is calculated by calculating the portion of solar and terrestrial radiation energy absorbed by the panel in the multi-panel radiation pressure model of a low-orbit satellite.

4. The method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms according to claim 1, characterized in that, In the dynamic radiation pressure model, if the time span between two consecutive epochs is too long, the initial temperature of the panel is redefined.

5. The method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms according to claim 1, characterized in that, Determining the initial temperature of the panel also includes assuming that the panel surface is in thermodynamic equilibrium, that is, the panel's radiation absorption power, thermal radiation release power, and heat transfer power are in balance.

6. The method for modeling the thermal radiation pressure of low-Earth orbit satellites based on heat transfer mechanisms according to claim 1, characterized in that, The panel temperature for subsequent epochs is updated using numerical integration methods, including: ， in, This represents the current epoch satellite panel surface temperature; The panel temperature of the previous epoch. This represents the temperature change between the two integral quantization intervals before and after the panel. The number of subintervals into which the numerical integral is divided. The sub-interval index; For the first The radiative absorption power of the panel in each integrator sub-interval; For the first The thermal radiation power of the panel in each integral sub-interval; For the first The heat transfer power between the panel of each integrator interval and the satellite body; The specific heat capacity of the panel determines the temperature change of the panel caused by the absorption, release, or transfer of heat.

7. A low-Earth orbit satellite thermal radiation pressure modeling system based on heat transfer mechanisms, characterized in that, include: The first main module is used to determine the optical and thermodynamic parameters of the panel. The second main module is used to calculate the panel's radiation absorption power based on the panel's optical parameters. The third main module is used to determine the initial temperature of the panel based on the panel's optical and thermodynamic parameters, combined with the panel's radiation absorption power. The fourth main module is used to update the panel temperature in subsequent epochs based on the initial panel temperature using a numerical integration method. The fifth main module is used to determine the thermal radiation pressure of low-orbit satellites based on the updated panel temperature.

8. A low-orbit satellite thermal radiation pressure modeling device based on heat transfer mechanism, characterized in that, The system includes a memory and a processor, wherein the memory stores program instructions that are executed by the processor, and the processor invokes the program instructions to execute the low-orbit satellite thermal radiation pressure modeling method based on the heat transfer mechanism as described in any one of claims 1 to 6.

9. A non-transitory computer-readable storage medium, characterized in that, The non-transitory computer-readable storage medium stores computer instructions that cause the computer to execute the low-orbit satellite thermal radiation pressure modeling method based on heat transfer mechanism as described in any one of claims 1 to 6.

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