Space radiation measurement reference satellite orbit design and optimization method and system

Abstract

The invention provides a space radiation measurement reference satellite orbit design and optimization method and system. Comprising the steps of determining an orbit height difference range of a reference satellite and a satellite to be subjected to reference transmission, a reference satellite orbit inclination angle range, a reference satellite revisit period and orbit height, an ascending node right ascension, an orbit eccentricity rate, a perigee argument, networking work requirements and orbit parameters, and completing space radiation measurement reference satellite orbit design optimization. The reference transfer model is established according to reference transfer load characteristics and transfer constraint conditions, the transfer frequency, single transfer duration and transfer scene space-time distribution of a reference satellite and a satellite to be subjected to reference transfer are simulated and analyzed according to the reference transfer model, the orbit parameters of the reference satellite are optimized, the networking efficiency of the reference satellite is considered, and the time consumption of the reference satellite is reduced. The method can serve in-orbit reference transmission of multi-type multi-satellite remote sensing satellites.

Description

Technical Field

[0001] The present invention belongs to the technical field of on-orbit absolute radiation calibration, and specifically relates to a method and system for designing and optimizing the orbit of a space radiation measurement reference satellite, and in particular to a method for designing the orbit of a space radiation measurement reference satellite for performing reference transfer on a remote sensing satellite. Background Art

[0002] After decades of development, the accuracy of remote sensing satellite radiometric calibration remains elusive, constrained by traditional payload calibration system design and the theoretical limitations of ground-based radiometric correction technology. To further enhance accuracy, applying ground-based metrology techniques to space presents significant challenges, including multiple difficulties with reference sources, payloads, platforms, and space-to-ground mission planning and scheduling. Furthermore, the cost is significantly higher than with traditional onboard calibration systems, making it unrealistic to carry a space radiometric reference payload on every operational satellite. Therefore, the concept of a Space Radiometric Reference Satellite program was proposed, aiming to launch an ultra-high-precision radiometric reference satellite with traceability capabilities. Establishing a space radiometric reference satellite, or reference satellite, is a fundamental means of studying long-term climate change and ensuring the integrity of multi-source remote sensing data and products.

[0003] The orbit design of the space radiation measurement reference satellite is affected by the characteristics of the payload to be transferred, the characteristics of the reference payload, the reference transfer constraints, the frequency and duration of the reference transfer, the temporal and spatial distribution of the reference transfer scenarios, etc., which will fundamentally affect the frequency and accuracy of the on-orbit reference transfer. At the same time, the setting needs to take into account the long-term stable observation of global climate factors.

[0004] At present, in the field of space radiation measurement reference satellites, the patent document "A method for land resource satellite calibration orbit planning and reference payload orbit optimization" (CN111523209A) proposes a method for land resource satellite calibration orbit planning and reference payload orbit optimization. It optimizes the orbit of the radiation reference satellite through the least squares method, but does not consider the reference transfer constraints and the flexible maneuverability of the reference satellite, and does not study the spatiotemporal distribution of the reference transfer scenario.

[0005] The patent document "A Radiation Calibration Method Based on an On-Orbit Reference Satellite" (CN106643796A) proposes an on-orbit absolute radiation calibration method based on a reference satellite, but does not demonstrate or set the orbit of the reference satellite; the patent document "Space Radiation Reference Transfer Calibration Method" (CN114674421A) focuses on using the observation data of the radiation reference satellite to correct the calibration coefficient of the satellite payload to be calibrated, but also does not demonstrate or set the orbit of the reference satellite.

[0006] The paper "Cross-calibration Benchmark Payload Orbit Optimization Method Based on Genetic Algorithm" proposes a benchmark payload orbit optimization method based on genetic algorithm that serves multi-target payload orbits. It also does not consider the benchmark transfer duration and the spatiotemporal distribution of the benchmark transfer scenario. The paper "China's Space Radiation Measurement Benchmark Technology" systematically introduces the development history, current status and planning of the benchmark satellite, but does not describe the detailed orbit design of the benchmark satellite.

[0007] Therefore, there is an urgent need for a space radiation measurement reference satellite orbit design method for benchmark transmission to remote sensing satellites, to clarify the orbit design and optimization of the reference satellite, and to solve the problem of difficulty in improving the absolute radiation calibration accuracy of remote sensing satellite payloads on orbit. Summary of the Invention

[0008] In view of the defects in the prior art, the purpose of the present invention is to provide a method and system for designing and optimizing the orbit of a space radiation measurement reference satellite.

[0009] A method for designing and optimizing a space radiation measurement reference satellite orbit according to the present invention includes:

[0010] Step 1: Determine the orbital height difference range between the reference satellite and the satellite to be transferred based on the orbital height of the remote sensing satellite to be transferred;

[0011] Step 2: Determine the orbital inclination range of the reference satellite according to the mission requirements of the reference satellite;

[0012] Step 3: Determine the revisit period and orbital altitude of the reference satellite based on the field of view angle, orbital altitude difference range, and climate observation requirements of the reference payload carried by the reference satellite.

[0013] Step 4: Simulate and analyze the transmission frequency, single transmission duration, and spatiotemporal distribution of transmission scenarios between the reference satellite and the satellite to be transmitted to the reference according to the reference transmission constraint conditions;

[0014] Step 5: Determine the orbital eccentricity and perigee argument according to the reference satellite orbit freezing requirements;

[0015] Step 6: Determine the reference satellite network requirements and orbital parameters based on the spatiotemporal distribution of the transmission scenario;

[0016] Step 7. Complete the orbit design of the space radiation measurement reference satellite using orbit altitude, revisit period, eccentricity, argument of perigee, orbit inclination, right ascension of ascending node and networking characteristics.

[0017] Preferably, in step 1, the reference satellite adjusts the yaw and roll two-dimensional attitudes to keep the reference payload consistent with the observation axis of the payload to be transferred to the reference. When the orbital plane of the reference satellite is perpendicular to the orbital plane of the satellite to be transferred to the reference, the demand for the roll angular rate of the reference satellite reaches a maximum value, and the roll angular rate of the reference satellite is calculated. Find the corresponding orbital height difference between the reference satellite and the satellite to be transferred to the reference in the reference satellite roll angular rate change curve;

[0018] in, is the orbital angular velocity of the reference star;

[0019] h J is the orbital height of the reference star;

[0020] R E is the radius of the Earth;

[0021] μ E =3.98603×10 5 km 3 / s 2 is the Earth's gravitational constant;

[0022] α is the reference transfer geocentric half angle,

[0023] θ is the semi-cone angle of the load to be transferred to the ground for observation;

[0024] h M Transfer the satellite orbit altitude to be referenced.

[0025] The mission requirements include global coverage, climate observations, and the spatial and temporal distribution of benchmark delivery scenarios.

[0026] The reference star orbit inclination in step 2 ranges from 50° to 100°.

[0027] Preferably, in step 3, the ground equatorial spacing between adjacent reference satellite orbits is greater than the ground width corresponding to the reference payload field of view angle, and the reference satellite regression period is determined;

[0028] The orbit height is determined based on the relationship between the regression orbit node period and the orbit semi-major axis and inclination.

[0029] The orbit semi-major axis is determined based on the regression coefficient and the orbit node period.

[0030] The benchmark transfer constraints include:

[0031] The time difference between the reference satellite and the satellite to be transferred to the reference observing the same target is less than 5 minutes;

[0032] The observation angles of the reference satellite and the satellite to be transferred to the reference are matched within 1°.

[0033] The observation angles include the solar zenith angle, the observation azimuth angle and the observation zenith angle.

[0034] Preferably, in step 4, a reference transfer model is established according to the reference transfer constraints and payload width, and the transfer frequency, single transfer duration and spatiotemporal distribution of transfer scenarios between the reference satellite and the satellite to be transferred to the reference are simulated and analyzed to optimize the orbital parameters.

[0035] The method for establishing the reference transfer model is to introduce a pyramid configuration, which is fixed to the satellite to be transferred and moves with the satellite to be transferred in the orbital plane. The top of the pyramid is the center of mass of the earth E ＊ The line connecting the pyramid vertex and the satellite to be transferred to the benchmark passes through the center of the base of the pyramid, and the other four vertices of the pyramid are located on the virtual sphere of the benchmark satellite's orbit.

[0036] Introduce the angle constants ψ and ε, ψ=n M T is located in the orbital plane of the satellite to be transferred to the reference, and the angle constant ε is perpendicular to the orbital plane of the satellite to be transferred to the reference. When the orbital altitude of the satellite to be transferred to the reference is not lower than that of the reference satellite, ε = a - θ; when the orbital altitude of the satellite to be transferred to the reference is lower than that of the reference satellite, ε = θ - a;

[0037]

[0038] Among them, n M Transfer the satellite orbital angular velocity to be referenced;

[0039] T is the benchmark transfer time constraint;

[0040] θ is the ground observation semi-cone angle of the load to be transferred by the benchmark;

[0041] R M Indicates the orbital height of the satellite to be transferred as the reference;

[0042] R J Indicates the orbital altitude of the reference star.

[0043] When the trajectory of the benchmark star is inside the pyramid, it is considered that a benchmark transfer event has occurred;

[0044] According to the benchmark transfer model and combined with mission requirements, the benchmark transfer efficiency between the benchmark satellite and the satellite to be benchmarked is simulated and analyzed, and the orbital inclination and ascending node right ascension parameters are traversed and optimized.

[0045] Preferably, in step 5, the reference satellite adopts a small eccentricity near-circular frozen orbit, with the perigee argument ω=90° and the orbit eccentricity

[0046] Where i is the orbital inclination;

[0047] a is the semi-major axis of the orbit;

[0048] R E is the radius of the Earth;

[0049] J2 and J3 are both high-order perturbations due to the flat Earth.

[0050] In step six, the benchmark satellite networking work requirements and orbital parameters are determined based on the payload characteristics and spatiotemporal distribution of the satellite to be benchmarked.

[0051] According to the present invention, a space radiation measurement reference satellite orbit design and optimization system is provided, comprising:

[0052] Module 1: determining the orbital height difference range between the reference satellite and the satellite to be transferred based on the orbital height of the remote sensing satellite to be transferred;

[0053] Module 2: Determine the orbital inclination range of the reference satellite based on its mission requirements;

[0054] Module 3: Determine the revisit period and orbital altitude of the reference satellite based on the field of view angle, orbital altitude difference range and climate observation requirements of the reference payload carried by the reference satellite;

[0055] Module 4: Simulate and analyze the transmission frequency, single transmission duration, and spatiotemporal distribution of transmission scenarios between the reference satellite and the satellite to be transmitted based on the reference transmission constraint conditions;

[0056] Module 5: Determine the orbital eccentricity and perigee argument according to the requirements of the reference satellite orbit freezing;

[0057] Module 6: Determine the benchmark satellite network requirements and orbital parameters based on the spatiotemporal distribution of the transmission scenario;

[0058] Module 7: Complete the orbit design of the space radiation measurement reference satellite using orbit altitude, revisit period, eccentricity, argument of perigee, orbit inclination, right ascension of ascending node and networking characteristics.

[0059] Preferably, the reference satellite in the module 1 adjusts the yaw and roll two-dimensional attitude to keep the reference payload consistent with the observation axis of the payload to be transferred to the reference. When the orbital plane of the reference satellite is perpendicular to the orbital plane of the satellite to be transferred to the reference, the demand for the reference satellite's roll angular rate reaches a maximum value, and the reference satellite's roll angular rate is calculated. Find the corresponding orbital height difference between the reference satellite and the satellite to be transferred to the reference in the reference satellite roll angular rate change curve;

[0060] in, is the orbital angular velocity of the reference star;

[0061] h J is the orbital height of the reference star;

[0062] R E is the radius of the Earth;

[0063] μ E =3.98603×105 km 3 / s 2 is the Earth's gravitational constant;

[0064] α is the reference transfer geocentric half angle,

[0065] θ is the semi-cone angle of the load to be transferred to the ground for observation;

[0066] h M Transfer the satellite orbit altitude to be referenced.

[0067] The mission requirements include global coverage, climate observations, and the spatial and temporal distribution of benchmark delivery scenarios.

[0068] The reference star orbit inclination in the module 2 ranges from 50° to 100°.

[0069] Preferably, the ground equatorial spacing between adjacent tracks of the reference satellite orbit in the module three is greater than the ground width corresponding to the reference payload field of view angle, and the reference satellite regression period is determined;

[0070] The orbit height is determined based on the relationship between the regression orbit node period and the orbit semi-major axis and inclination.

[0071] The orbit semi-major axis is determined based on the regression coefficient and the orbit node period.

[0072] The benchmark transfer constraints include:

[0073] The time difference between the reference satellite and the satellite to be transferred to the reference observing the same target is less than 5 minutes;

[0074] The observation angles of the reference satellite and the satellite to be transferred to the reference are matched within 1°.

[0075] The observation angles include the solar zenith angle, the observation azimuth angle and the observation zenith angle.

[0076] Preferably, in the module four, a reference transfer model is established according to the reference transfer constraints and the payload width, and the transfer frequency, single transfer duration and spatiotemporal distribution of the transfer scenario between the reference satellite and the satellite to be referenced are simulated and analyzed to optimize the orbital parameters.

[0077] The method for establishing the reference transfer model is to introduce a pyramid configuration, which is fixed to the satellite to be transferred and moves with the satellite to be transferred in the orbital plane. The top of the pyramid is the center of mass of the earth E ＊ The line connecting the pyramid vertex and the satellite to be transferred to the benchmark passes through the center of the base of the pyramid, and the other four vertices of the pyramid are located on the virtual sphere of the benchmark satellite's orbit.

[0078] Introduce the angle constants ψ and ε, ψ=n MT is located in the orbital plane of the satellite to be transferred to the reference, and the angle constant ε is perpendicular to the orbital plane of the satellite to be transferred to the reference. When the orbital altitude of the satellite to be transferred to the reference is not lower than that of the reference satellite, ε = α - θ; when the orbital altitude of the satellite to be transferred to the reference is lower than that of the reference satellite, ε = θ - a;

[0079]

[0080] Among them, n M Transfer the satellite orbital angular velocity to be referenced;

[0081] T is the benchmark transfer time constraint;

[0082] θ is the ground observation semi-cone angle of the load to be transferred by the benchmark;

[0083] R M Indicates the orbital height of the satellite to be transferred as the reference;

[0084] R J Indicates the orbital altitude of the reference star.

[0085] When the orbit of the benchmark star is within the pyramid, it is considered that a benchmark transfer event has occurred.

[0086] According to the benchmark transfer model and combined with mission requirements, the benchmark transfer efficiency between the benchmark satellite and the satellite to be benchmarked is simulated and analyzed, and the orbital inclination and ascending node right ascension parameters are traversed and optimized.

[0087] Preferably, the reference satellite in the module 5 adopts a small eccentricity near-circular frozen orbit, with the perigee argument ω=90° and the orbit eccentricity

[0088] Where i is the orbital inclination;

[0089] a is the semi-major axis of the orbit;

[0090] R E is the radius of the Earth;

[0091] J2 and J3 are both high-order perturbations due to the flat Earth.

[0092] In the module six, the benchmark satellite networking work requirements and orbital parameters are determined based on the payload characteristics and spatiotemporal distribution of the satellite to be benchmarked.

[0093] Compared with the prior art, the present invention has the following beneficial effects:

[0094] 1. The present invention establishes a reference transfer model based on the reference transfer payload characteristics and transfer constraints, takes into account the reference satellite networking efficiency, and can serve the on-orbit reference transfer of multiple types and multiple satellites of remote sensing satellites.

[0095] 2. The present invention simulates and analyzes the transmission frequency, single transmission duration and spatiotemporal distribution of transmission scenarios between the reference satellite and the satellite to be benchmarked based on the reference transmission model, optimizes the orbital parameters of the reference satellite, and takes into account the long-term stable observation of global climate factors.

[0096] 3. The present invention can clarify and optimize the orbit design of the reference satellite, solving the problem of difficulty in improving the accuracy of the on-orbit absolute radiation calibration of remote sensing satellite payloads. BRIEF DESCRIPTION OF THE DRAWINGS

[0097] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments with reference to the following drawings:

[0098] Figure 1 Schematic diagram of the process of designing and optimizing the orbit of a reference satellite for space radiation measurement;

[0099] Figure 2 This is a schematic diagram of the transfer when the orbital plane of the reference satellite is perpendicular to the orbital plane of the satellite to be transferred to the reference;

[0100] Figure 3 This is a schematic diagram of the benchmark star roll angular rate change curve;

[0101] Figure 4 This is a diagram showing the distribution of longitude and latitude of the subsatellite point when the right ascension of the ascending node is fixed at 0° and the inclination is 50°;

[0102] Figure 5 This is a diagram showing the distribution of longitude and latitude of the subsatellite point when the right ascension of the ascending node is fixed at 0° and the inclination is 90°;

[0103] Figure 6 This is a diagram showing the distribution of the starting latitude over time when the right ascension of the ascending node is fixed at 0° and the inclination is 50°;

[0104] Figure 7 This is a diagram showing the distribution of the starting latitude over time when the right ascension of the ascending node is fixed at 0° and the inclination is 90°;

[0105] Figure 8 This is a diagram showing the distribution of longitude and latitude of the subsatellite point when the inclination angle is fixed at 90° and the right ascension of the ascending node is 120°;

[0106] Figure 9 This is a diagram showing the distribution of the starting latitude over time when the right ascension of the ascending node is 120° when the inclination angle is fixed at 90°;

[0107] Figure 10 Schematic diagram of the space-time distribution when the three benchmark stars form a network and the right ascension of the ascending node differs by 120°. DETAILED DESCRIPTION

[0108] The present invention will be described in detail below with reference to specific embodiments. The following examples will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those skilled in the art, several changes and improvements can be made without departing from the scope of the present invention. These all fall within the scope of protection of the present invention.

[0109] According to the present invention, a method for designing and optimizing the orbit of a reference satellite for space radiation measurement is provided. Figure 1 For example, the following steps are included:

[0110] Step 1: Determine the orbital height difference range between the reference satellite and the remote sensing satellite to be transferred based on the orbital height of the remote sensing satellite to be transferred, so as to meet the flexible maneuvering pointing capability of the reference satellite;

[0111] Specifically, the specific method for determining the orbital height difference range between the reference satellite and the satellite to be referenced is:

[0112] The reference satellite makes the reference payload and the target reference transfer payload observe the same visual axis through agile attitude maneuvers, that is, the observation azimuth and observation zenith angle matching constraints are met, including the adjustment of the two-dimensional attitude in the yaw direction and the roll direction. When the orbital plane of the reference satellite is perpendicular to the orbital plane of the target reference transfer satellite, Figure 2 For example, the reference transmission time between the two satellites reaches a minimum, and the reference satellite roll angular rate requirement reaches a maximum, which is used to illustrate the rate requirement for roll maneuvers in extreme situations.

[0113] The constraints of the benchmark transfer include time, space, lighting conditions and other constraints. For example, the time constraint is that the time difference between the benchmark star and the target star for the same target does not exceed 5 minutes; the space constraint is that the spatial distribution of the observed target (geographic longitude and latitude); the lighting constraint is that the solar altitude angle when the benchmark star and the target star observe the same target is greater than or equal to the solar altitude angle when the benchmark star and the target star observe the same target.

[0114] When the orbital height difference between the reference satellite and the satellite to be transferred to the reference takes different values, the curve of the reference satellite rolling angular rate change is as follows: Figure 3 For example, the reference satellite rolling angular rate reflects the reference satellite's flexible maneuvering pointing capability, and the calculation formula is:

[0115]

[0116] in, is the orbital angular velocity of the reference star, h J is the orbital height of the reference star, R E is the radius of the Earth, μ E =3.98603×10 5 km 3 / s 2 is the earth's gravitational constant, a is the half angle of the base transfer center of the earth, and it can be obtained by the sine theorem:

[0117]

[0118] θ is the semi-cone angle of the load to be transferred to the ground for observation, h M Transfer the satellite orbit altitude to be referenced.

[0119] The range of the orbital height difference between the reference satellite and the satellite to be benchmarked can be determined based on the orbital height of the satellite to be benchmarked and the flexible maneuvering pointing capability of the reference satellite according to the rolling angular rate change curve of the reference satellite.

[0120] The observational data acquired by the remote sensing payload to be transferred by the benchmark is determined by its characteristics and observation requirements and is not used as a process parameter for orbit design. However, the payload's design parameters, such as the field of view (FOV), are used as orbit design parameters. The payload's FOV is determined by the benchmark payload's orbital altitude, spatial resolution, and optical characteristics and requires iterative configuration. The satellites to be transferred by the benchmark are quantitative remote sensing satellites used for meteorological, environmental, resource, ocean, and land observations, requiring in-orbit absolute radiometric calibration.

[0121] In more preferred examples, Fengyun-3F is used as the satellite to be transferred to the benchmark, and the medium-resolution spectral imager carried on it is used as the remote sensing payload to be transferred to the benchmark.

[0122] The orbit altitude is about 836km, and the medium-resolution spectroscopic imager's Earth observation semi-cone angle is 55°.

[0123] according to Figure 3 As shown in the variation curve, when the orbital height difference between the two satellites is 200 km, the roll angular rate of the reference satellite reaches 2.5° / s. Considering the agile maneuverability of the reference satellite and the attitude pointing stability during the reference transfer, the orbital height of the reference satellite is appropriately selected to be around 600 km.

[0124] Step 2: Determine the orbital inclination range of the reference satellite based on the requirements of global coverage of the reference satellite, climate observation, temporal and spatial distribution of the reference transmission scenario, etc.

[0125] Specifically, the orbital inclination determines the orbital precession period and the maximum latitude range covered by the satellite. The orbital inclination also affects the number of precession cycles of the orbital plane per year. Low-inclination orbits have a fast rate of change of the ascending node (i.e., the precession rate of the orbital plane), and thus have richer spatiotemporal sampling of ground scenes in mid- and low-latitude regions. However, their disadvantage is that they cannot cover the entire world. Considering that the orbital inclinations of most low-orbit remote sensing satellites are between 50° and 100°, the reference star orbital inclination range is selected to be 50° to 100°.

[0126] In more preferred examples, the Fengyun-3F satellite to be transferred as the benchmark is in a sun-synchronous orbit with an orbital inclination of 98.75°. If the benchmark satellite orbit also chooses a sun-synchronous orbit, then under specific constraints, the intersection scenes with the benchmark transfer satellite are all located near the polar regions, mainly sea ice reflections, and the transfer scenes are single, which cannot meet the full-spectrum benchmark transfer accuracy requirements. Therefore, the benchmark satellite orbital inclination range is selected to be 50° to 100°. When global coverage is required, the orbital inclination is recommended to be set to 90°. When the benchmark transfer spatiotemporal scene distribution only focuses on mid- and low-latitude scenes, the orbital inclination is set to 50°.

[0127] The choice of orbital inclination mainly considers global coverage and the spatial and temporal richness of ground scene sampling. If global coverage is required to meet the needs of long-term stable climate observation, a 90° inclination orbit is preferred.

[0128] Step 3: Determine the revisit period and orbital altitude of the reference satellite based on the reference payload's field of view and climate observation requirements;

[0129] Specifically, the reference satellite revisit period is determined by the reference transfer satellite orbit altitude, climate observation requirements and reference payload field of view angle determined in step one. The principle is that the ground equatorial spacing between adjacent tracks of the reference satellite orbit is greater than the ground width corresponding to the reference payload field of view angle.

[0130] In more preferred examples, the proposed benchmark satellite will carry a Earth-Moon spectroscopic imager and an infrared interferometer. At an orbital altitude of 600 km, the reflectance spectrum of the benchmark payload will be approximately 50 km wide. With a return period greater than 56 days, global coverage is possible. The orbital node period is the time it takes for a satellite to cross the ascending node twice in the same direction, corresponding to the number of return days.

[0131] Because the inclination has little effect on the orbital node period, the orbital inclination can take a fixed value when determining the orbital semi-major axis from the regression coefficient and the orbital node period. In the embodiment of the present invention, the orbital inclination is 90°. When the regression period is 57 to 61 days, the optional orbital parameters for different regression coefficients at an orbital altitude of approximately 600 km are shown in Table 1.

[0132] Table 1. Optional orbital parameters around 600 km with a return period of 57 to 61 days:

[0133]

[0134]

[0135] Considering the revisit performance of the benchmark payload, the benchmark satellite regression period is selected as 61 days, the regression coefficient is 14+38 / 61, and the orbit altitude is 601.98km.

[0136] Step 4: Based on the constraints of the benchmark transfer and the payload field of view, the transfer frequency, single transfer duration, and the spatiotemporal distribution of the transfer scenario between the benchmark satellite and the satellite to be transferred to the benchmark are simulated and analyzed to further optimize the orbital parameters such as the benchmark satellite inclination and the ascending node right ascension.

[0137] Specifically, a benchmark transfer model is established based on constraints such as benchmark transfer constraints and load width.

[0138] According to the simulation analysis of the benchmark transfer model, the transmission frequency, single transmission duration and spatiotemporal distribution of transmission scenarios between the benchmark satellite and the satellite to be benchmarked are analyzed. By traversing the inclination and right ascension of the ascending node with a certain step size, the orbital parameters such as the inclination of the benchmark satellite and the right ascension of the ascending node are further optimized.

[0139] The benchmark transfer constraints generally include:

[0140] The time difference between the reference satellite and the satellite to be transferred to the reference observing the same uniform target is less than 5 minutes;

[0141] The observation angles of the reference satellite and the satellite to be transferred to the reference are matched within 1°, including the solar zenith angle, observation azimuth angle and observation zenith angle.

[0142] The method for establishing the reference transfer model is as follows: a pyramid configuration is introduced to describe the occurrence of a reference transfer event. The pyramid is fixed to the satellite to be transferred and moves with the satellite to be transferred in the orbital plane. The top of the pyramid is the center of mass of the earth E. ＊ The line connecting the pyramid vertex and the satellite to be transferred to the benchmark passes through the center of the base of the pyramid, and the other four vertices of the pyramid are located on the virtual sphere of the benchmark satellite's orbit.

[0143] Two angle constants are introduced to define the shape of the pyramid. The first one is ψ, which is located in the orbital plane of the satellite to be transferred to the reference. It is determined by the reference transfer time constraint and the orbital angular velocity of the satellite to be transferred to the reference. ψ=n M T, where n M is the orbital angular velocity of the satellite to be transferred, and T is the time constraint of the transfer.

[0144] The second angle constant is ε, which is perpendicular to the orbital plane of the satellite to be transferred. When the orbital altitude of the satellite to be transferred is not lower than that of the reference satellite, ε = α - θ, where θ is the semi-cone angle of the earth observation of the payload to be transferred, and α can be obtained by the sine theorem:

[0145]

[0146] When the orbital altitude of the satellite to be transferred to the reference is lower than that of the reference satellite, ε=θ-a. M Indicates the orbital height of the satellite to be transferred, R J Indicates the orbital altitude of the reference star.

[0147] When the orbit of the benchmark star is within the pyramid, it is considered that a benchmark transfer event has occurred.

[0148] In further preferred embodiments, T is set to 5 minutes. The θ value of the medium-resolution spectroscopic imager on the Fengyun-3F satellite is 55°. The ascending node right ascension step size is set to 30°, and the inclination step size is set to 10°. The orbital parameters, such as the reference satellite inclination and ascending node right ascension, are further optimized. With the ascending node right ascension fixed at 0°, the inclination angle is traversed from 50° to 100° in 10° steps. The annual reference transfer frequency between the reference satellite and the Fengyun-3F satellite is shown in Table 2.

[0149] Table 2. Frequency of benchmark transfers in the next year for different inclination values ​​when the right ascension of the ascending node is fixed:

[0150] Inclination angle (°) Benchmark delivery frequency 50 927 60 817 70 910 80 1027 90 934 100 905

[0151] When the inclination angle is fixed at 90°, the frequency of benchmark transfer between the lower benchmark satellite and Fengyun-3F satellite in one year is shown in Table 3.

[0152] Table 3. Frequency of benchmark transfers in the next year for different values ​​of right ascension of the ascending node when the inclination is fixed:

[0153] Ascending node right ascension (°) Benchmark delivery frequency 0 934 30 921 60 946 90 929 120 949 150 930 180 932 210 937 240 912 270 938 300 935 330 932

[0154] Figure 4 The longitude and latitude distribution of the subsatellite point of the benchmark satellite and Fengyun-3F satellite in one year of benchmark transfer events is shown when the right ascension of the ascending node is fixed at 0° and the inclination is 50°.

[0155] Figure 5 The longitude and latitude distribution of the subsatellite point of the benchmark satellite and Fengyun-3F satellite in one year of benchmark transfer events is shown when the right ascension of the ascending node is fixed at 0° and the inclination is 90°.

[0156] Figure 6 The temporal distribution of the starting latitude of the one-year benchmark transfer event between the benchmark satellite and Fengyun-3F satellite is shown when the right ascension of the ascending node is fixed at 0° and the inclination is 50°.

[0157] Figure 7 The temporal distribution of the starting latitude of the one-year benchmark transfer event between the benchmark satellite and Fengyun-3F satellite is shown when the right ascension of the ascending node is fixed at 0° and the inclination is 90°.

[0158] Although the low-inclination orbit cannot cover the entire world, due to the orbital precession characteristics, there are more benchmark transfer events in low-latitude regions, and the benchmark transfer events are more evenly distributed over time throughout the year, covering different seasons.

[0159] Figure 8The longitude and latitude distribution of the subsatellite point of the benchmark transfer event between the benchmark satellite and Fengyun-3F satellite in one year when the inclination angle is fixed at 90° and the right ascension of the ascending node is 120° is shown.

[0160] Step 5: Determine the orbital eccentricity and perigee argument according to the reference satellite orbit freezing requirements;

[0161] Specifically, the ground pixel resolution will change with the satellite altitude. In order to make the satellite have approximately the same resolution in various latitudes, the ground altitude of the satellite should be kept constant as much as possible, and the reference satellite adopts a small eccentricity near-circular orbit.

[0162] The small eccentricity near-circular orbit is an orbit whose orbital shape in the orbital plane is close to a circle. If the eccentricity is 0, it is a circle, but the actual eccentricity cannot be 0, so it is called a small eccentricity near-circular orbit.

[0163] When it is necessary to fly over areas of equal latitude at the same time, the height above the ground remains unchanged, and the orbit is further designed as a frozen orbit.

[0164] When a frozen orbit satellite flies over areas of equal latitude, its altitude from the ground remains unchanged, and the spatial resolution of areas at the same latitude is the same, which is conducive to image stitching by remote sensing satellites. The reference star orbit selects a frozen orbit.

[0165] The eccentricity and argument of perigee of the frozen orbit satisfy the following formula:

[0166] ω=90°

[0167]

[0168] Where ω is the argument of perigee, e is the orbital eccentricity, i is the orbital inclination, a is the semi-major axis of the orbit, and R E is the radius of the Earth, J2 and J3 are both high-order perturbations due to the flat Earth.

[0169] In more preferred embodiments, the determined orbital inclination and semi-major axis yield an eccentricity of the reference star of 0.0011.

[0170] Step 6: Determine the requirements and orbital parameters for the reference satellite network.

[0171] Specifically, the benchmark satellite networking work requirements and orbital parameters are determined according to the payload characteristics of the satellite to be benchmarked and the spatiotemporal distribution of the benchmark satellite delivery scenario.

[0172] In more preferred examples, the networking requirements, that is, the networking characteristics, are determined by the payload characteristics of the satellite to be benchmarked and the spatiotemporal distribution of the benchmark satellite transmission scene. The spatiotemporal distribution characteristics of the single-satellite and three-satellite network benchmark transmission scenes are determined by the Figure 9 and Figure 10 For example. Figure 9The temporal distribution of the starting latitude of the one-year benchmark transfer event between the benchmark satellite and Fengyun-3F satellite is shown when the inclination angle is fixed at 90° and the right ascension of the ascending node is 120°. Figure 10 The figure shows the temporal and spatial distribution of the benchmark transfer events between the benchmark star and Fengyun-3F in one year when the three benchmark stars form a network and the right ascension of the ascending node differs by 120°.

[0173] The orbital inclination of the benchmark satellite is designed to be 90°, which makes the spatial distribution of the benchmark transmission scene uneven with the Fengyun-3F satellite. The transmission scene in mid- and low-latitudes only occurs in specific seasons, affecting the benchmark transmission efficiency. Multi-satellite networking can enrich the temporal and spatial distribution of the benchmark transmission scene in mid- and low-latitude regions.

[0174] The three-satellite network of benchmarks can not only increase the frequency of benchmark transmission, but also enrich the spatiotemporal distribution of benchmark transmission events. The benchmark transmission scenarios can cover different seasons and latitudes, which can further improve the accuracy of benchmark transmission of space radiation measurement benchmark satellites.

[0175] Step 7: Use the orbital altitude, revisit period, eccentricity, argument of perigee, orbital inclination, right ascension of ascending node, and networking characteristics determined above to complete the orbit design of the space radiation measurement reference satellite.

[0176] The present invention also provides a space radiation measurement reference satellite orbit design and optimization system, which can be implemented by executing the process steps of the space radiation measurement reference satellite orbit design and optimization method. That is, those skilled in the art can understand the space radiation measurement reference satellite orbit design and optimization method as a preferred implementation of the space radiation measurement reference satellite orbit design and optimization system.

[0177] According to the present invention, a space radiation measurement reference satellite orbit design and optimization system is provided, comprising:

[0178] Module 1: determining the orbital height difference range between the reference satellite and the satellite to be transferred based on the orbital height of the remote sensing satellite to be transferred;

[0179] Module 2: Determine the orbital inclination range of the reference satellite based on its mission requirements;

[0180] Module 3: Determine the revisit period and orbital altitude of the reference satellite based on the field of view angle, orbital altitude difference range and climate observation requirements of the reference payload carried by the reference satellite;

[0181] Module 4: Simulate and analyze the transmission frequency, single transmission duration, and spatiotemporal distribution of transmission scenarios between the reference satellite and the satellite to be transmitted based on the reference transmission constraint conditions;

[0182] Module 5: Determine the orbital eccentricity and perigee argument according to the requirements of the reference satellite orbit freezing;

[0183] Module 6: Determine the benchmark satellite network requirements and orbital parameters based on the spatiotemporal distribution of the transmission scenario;

[0184] Module 7: Complete the orbit design of the space radiation measurement reference satellite using orbit altitude, revisit period, eccentricity, argument of perigee, orbit inclination, right ascension of ascending node and networking characteristics.

[0185] In more preferred embodiments, the reference satellite in the module 1 adjusts the two-dimensional attitude in the yaw direction and the roll direction to keep the reference payload consistent with the observation axis of the payload to be transferred to the reference. When the orbital plane of the reference satellite is perpendicular to the orbital plane of the satellite to be transferred to the reference, the demand for the roll angular rate of the reference satellite reaches a maximum value. At this time, the roll angular rate of the reference satellite is calculated. Find the corresponding orbital height difference between the reference satellite and the satellite to be transferred to the reference in the reference satellite roll angular rate change curve;

[0186] in, is the orbital angular velocity of the reference star;

[0187] h J is the orbital height of the reference star;

[0188] R E is the radius of the Earth;

[0189] μ E =3.98603×10 5 km 3 / s 2 is the Earth's gravitational constant;

[0190] a is the half angle of the base transfer geocentric,

[0191] θ is the semi-cone angle of the load to be transferred to the ground for observation;

[0192] h M Transfer the satellite orbit altitude to be referenced.

[0193] The mission requirements include global coverage, climate observations, and the spatial and temporal distribution of benchmark delivery scenarios.

[0194] Considering that the orbital inclination of most low-orbit remote sensing satellites is between 50° and 100°, the orbital inclination of the reference satellite in the module 2 is in the range of 50° to 100°.

[0195] In more preferred embodiments, the ground equatorial spacing between adjacent tracks of the reference satellite orbit in module three is greater than the ground width corresponding to the reference payload field of view angle, and the reference satellite regression period is determined according to this principle;

[0196] The orbit height is determined based on the relationship between the regression orbit node period and the orbit semi-major axis and inclination.

[0197] The orbit semi-major axis is determined based on the regression coefficient and the orbit node period.

[0198] The benchmark transfer constraints include:

[0199] The time difference between the reference satellite and the satellite to be transferred to the reference observing the same target is less than 5 minutes;

[0200] The observation angles of the reference satellite and the satellite to be transferred to the reference are matched within 1°.

[0201] The observation angles include the solar zenith angle, the observation azimuth angle and the observation zenith angle.

[0202] In more preferred examples, the module four establishes a benchmark transfer model based on the benchmark transfer constraints and payload width, simulates and analyzes the transmission frequency, single transmission duration and spatiotemporal distribution of transmission scenarios between the benchmark satellite and the satellite to be benchmarked, and further optimizes orbital parameters such as the benchmark star inclination and ascending node right ascension.

[0203] The method for establishing the reference transfer model is to introduce a pyramid configuration, which is fixed to the satellite to be transferred and moves with the satellite to be transferred in the orbital plane. The top of the pyramid is the center of mass of the earth E ＊ The line connecting the pyramid vertex and the satellite to be transferred to the benchmark passes through the center of the base of the pyramid, and the other four vertices of the pyramid are located on the virtual sphere of the benchmark satellite's orbit.

[0204] Introduce the angle constants ψ and ε, ψ=n M T is located in the orbital plane of the satellite to be transferred to the reference, and the angle constant ε is perpendicular to the orbital plane of the satellite to be transferred to the reference. When the orbital altitude of the satellite to be transferred to the reference is not lower than that of the reference satellite, ε = a - θ; when the orbital altitude of the satellite to be transferred to the reference is lower than that of the reference satellite, ε = θ - a;

[0205]

[0206] Among them, n M Transfer the satellite orbital angular velocity to be referenced;

[0207] T is the benchmark transfer time constraint;

[0208] θ is the ground observation semi-cone angle of the load to be transferred by the benchmark;

[0209] R M Indicates the orbital height of the satellite to be transferred as the reference;

[0210] R J Indicates the orbital altitude of the reference star.

[0211] When the orbit of the benchmark star is within the pyramid, it is considered that a benchmark transfer event has occurred.

[0212] According to the benchmark transfer model and combined with mission requirements, the benchmark transfer efficiency between the benchmark satellite and the satellite to be benchmarked is simulated and analyzed, and parameters such as orbital inclination and ascending node right ascension are traversally optimized.

[0213] In more preferred embodiments, the reference satellite in the module 5 adopts a small eccentricity near-circular frozen orbit, with the perigee argument ω=90° and the orbit eccentricity

[0214] Where i is the orbital inclination;

[0215] a is the semi-major axis of the orbit;

[0216] R E is the radius of the Earth;

[0217] J2 and J3 are both high-order perturbations due to the flat Earth.

[0218] In the module six, the benchmark satellite networking work requirements and orbital parameters are determined based on the payload characteristics and spatiotemporal distribution of the satellite to be benchmarked.

[0219] Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, and units provided by the present invention in purely computer-readable program code, it is entirely possible to implement the same functions of the system and its various devices, modules, and units provided by the present invention in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; the devices, modules, and units for implementing various functions can also be considered as both software modules implementing the method and structures within the hardware component.

[0220] The above describes specific embodiments of the present invention. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art may make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. The embodiments of this application and the features in the embodiments may be combined with each other in any manner unless there is a conflict.

Claims

1. A method for designing and optimizing a reference satellite orbit for space radiation measurement, characterized in that: include: Step 1: Determine the orbital height difference range between the reference satellite and the satellite to be transferred based on the orbital height of the remote sensing satellite to be transferred; Step 2: Determine the orbital inclination range of the reference satellite according to the mission requirements of the reference satellite; Step 3: Determine the revisit period and orbital altitude of the reference satellite based on the field of view angle, orbital altitude difference range, and climate observation requirements of the reference payload carried by the reference satellite. Step 4: Simulate and analyze the transmission frequency, single transmission duration, and spatiotemporal distribution of transmission scenarios between the reference satellite and the satellite to be transmitted to the reference according to the reference transmission constraint conditions; Step 5: Determine the orbital eccentricity and perigee argument according to the reference satellite orbit freezing requirements; Step 6: Determine the reference satellite network requirements and orbital parameters based on the spatiotemporal distribution of the transmission scenario; Step 7. Complete the orbit design of the space radiation measurement reference satellite using orbit altitude, revisit period, eccentricity, argument of perigee, orbit inclination, right ascension of ascending node and networking characteristics.

2. The method for designing and optimizing the orbit of a reference satellite for space radiation measurement according to claim 1, wherein: In step 1, the reference satellite adjusts the yaw and roll two-dimensional attitudes to keep the reference payload consistent with the observation axis of the payload to be transferred to the reference. When the orbital plane of the reference satellite is perpendicular to the orbital plane of the satellite to be transferred to the reference, the demand for the reference satellite's roll angular rate reaches a maximum value. The reference satellite's roll angular rate is calculated. Find the corresponding orbital height difference between the reference satellite and the satellite to be transferred to the reference in the reference satellite roll angular rate change curve; in, is the orbital angular velocity of the reference star; h J is the orbital height of the reference star; R E is the radius of the Earth; μ E =3.98603×10 5 km 3 / s 2 is the Earth's gravitational constant; α is the reference transfer geocentric half angle, θ is the semi-cone angle of the load to be transferred to the ground for observation; h M Transfer the satellite orbital altitude to be referenced; The mission requirements include global coverage, climate observations, and the spatial and temporal distribution of benchmark delivery scenarios; The reference star orbit inclination in step 2 ranges from 50° to 100°.

3. The method for designing and optimizing the orbit of a reference satellite for space radiation measurement according to claim 1, wherein: In step 3, the ground equatorial spacing between adjacent reference satellite orbits is greater than the ground width corresponding to the reference payload field of view angle, and the reference satellite regression period is determined; The orbit height is determined based on the relationship between the regression orbit node period and the orbit semi-major axis and inclination; The orbital semi-major axis is determined according to the regression coefficient and the orbital node period; The benchmark transfer constraints include: The time difference between the reference satellite and the satellite to be transferred to the reference observing the same target is less than 5 minutes; The observation angles of the reference satellite and the satellite to be transferred to the reference satellite match within 1°; The observation angles include the solar zenith angle, the observation azimuth angle and the observation zenith angle.

4. The method for designing and optimizing the orbit of a reference satellite for space radiation measurement according to claim 1, wherein: In step 4, a reference transfer model is established based on the reference transfer constraints and payload width, and the transfer frequency, single transfer duration, and spatiotemporal distribution of transfer scenarios between the reference satellite and the satellite to be transferred to the reference are simulated and analyzed to optimize the orbital parameters. The method for establishing the reference transfer model is to introduce a pyramid configuration, which is fixed to the satellite to be transferred and moves with the satellite to be transferred in the orbital plane. The top of the pyramid is the center of mass of the earth E ＊ , the line connecting the pyramid vertex and the satellite to be transferred to the benchmark passes through the center of the pyramid base, and the other four vertices of the pyramid are located on the virtual sphere of the benchmark satellite orbit; Introduce the angle constants ψ and ε, ψ=n M T is located in the orbital plane of the satellite to be transferred to the reference, and the angle constant ε is perpendicular to the orbital plane of the satellite to be transferred to the reference. When the orbital altitude of the satellite to be transferred to the reference is not lower than that of the reference satellite, ε = α - θ; when the orbital altitude of the satellite to be transferred to the reference is lower than that of the reference satellite, ε = θ - α; Among them, n M Transfer the satellite orbital angular velocity to be referenced; T is the benchmark transfer time constraint; θ is the ground observation semi-cone angle of the load to be transferred by the benchmark; R M Indicates the orbital height of the satellite to be transferred as the reference; R J Indicates the orbital altitude of the reference star; When the trajectory of the benchmark star is inside the pyramid, it is considered that a benchmark transfer event has occurred; According to the benchmark transfer model and combined with mission requirements, the benchmark transfer efficiency between the benchmark satellite and the satellite to be benchmarked is simulated and analyzed, and the orbital inclination and ascending node right ascension parameters are traversed and optimized.

5. The method for designing and optimizing the orbit of a reference satellite for space radiation measurement according to claim 1, wherein: In step 5, the reference satellite adopts a small eccentricity near-circular frozen orbit, with the perigee argument ω=90° and the orbit eccentricity Where i is the orbital inclination; a is the semi-major axis of the orbit; R E is the radius of the Earth; J2 and J3 are both high-order perturbations due to the flat earth; In step six, the benchmark satellite networking work requirements and orbital parameters are determined based on the payload characteristics and spatiotemporal distribution of the satellite to be benchmarked.

6. A space radiation measurement reference satellite orbit design and optimization system, characterized in that: include: Module 1: determining the orbital height difference range between the reference satellite and the satellite to be transferred based on the orbital height of the remote sensing satellite to be transferred; Module 2: Determine the orbital inclination range of the reference satellite based on its mission requirements; Module 3: Determine the revisit period and orbital altitude of the reference satellite based on the field of view angle, orbital altitude difference range and climate observation requirements of the reference payload carried by the reference satellite; Module 4: Simulate and analyze the transmission frequency, single transmission duration, and spatiotemporal distribution of transmission scenarios between the reference satellite and the satellite to be transmitted based on the reference transmission constraint conditions; Module 5: Determine the orbital eccentricity and perigee argument according to the requirements of the reference satellite orbit freezing; Module 6: Determine the benchmark satellite network requirements and orbital parameters based on the spatiotemporal distribution of the transmission scenario; Module 7: Complete the orbit design of the space radiation measurement reference satellite using orbit altitude, revisit period, eccentricity, argument of perigee, orbit inclination, right ascension of ascending node and networking characteristics.

7. The space radiation measurement reference satellite orbit design and optimization system according to claim 6, characterized in that: In the module 1, the reference satellite adjusts the yaw and roll two-dimensional attitude to keep the reference payload consistent with the observation axis of the payload to be transferred to the reference. When the orbital plane of the reference satellite is perpendicular to the orbital plane of the satellite to be transferred to the reference, the demand for the reference satellite's roll angular rate reaches a maximum value. The reference satellite's roll angular rate is calculated. Find the corresponding orbital height difference between the reference satellite and the satellite to be transferred to the reference in the reference satellite roll angular rate change curve; in, is the orbital angular velocity of the reference star; h J is the orbital height of the reference star; R E is the radius of the Earth; μ E =3.98603×10 5 km 3 / s 2 is the Earth's gravitational constant; α is the reference transfer geocentric half angle, θ is the semi-cone angle of the load to be transferred to the ground for observation; h M Transfer the satellite orbital altitude to be referenced; The mission requirements include global coverage, climate observations, and the spatial and temporal distribution of benchmark delivery scenarios; The reference star orbit inclination in the module 2 ranges from 50° to 100°.

8. The space radiation measurement reference satellite orbit design and optimization system according to claim 6, characterized in that: The ground equatorial spacing between adjacent tracks of the reference satellite orbit in the module three is greater than the ground width corresponding to the reference payload field of view angle, and the reference satellite regression period is determined; The orbit height is set according to the relationship between the regression orbit node period and the orbit semi-major axis and inclination; The orbital semi-major axis is determined based on the regression coefficient and the orbital node period; The benchmark transfer constraints include: The time difference between the reference satellite and the satellite to be transferred to the reference observing the same target is less than 5 minutes; The observation angles of the reference satellite and the satellite to be transferred to the reference satellite match within 1°; The observation angles include the solar zenith angle, the observation azimuth angle and the observation zenith angle.

9. The space radiation measurement reference satellite orbit design and optimization system according to claim 6, characterized in that: In the fourth module, a reference transfer model is established according to the reference transfer constraints and payload width, and the transfer frequency, single transfer duration, and spatiotemporal distribution of transfer scenarios between the reference satellite and the satellite to be transferred to the reference are simulated and analyzed to optimize the orbital parameters. The method for establishing the reference transfer model is to introduce a pyramid configuration, which is fixed to the satellite to be transferred and moves with the satellite to be transferred in the orbital plane. The top of the pyramid is the center of mass of the earth E ＊ , the line connecting the pyramid vertex and the satellite to be transferred to the benchmark passes through the center of the pyramid base, and the other four vertices of the pyramid are located on the virtual sphere of the benchmark satellite orbit; Introduce the angle constants ψ and ε, ψ=n M T is located in the orbital plane of the satellite to be transferred to the reference, and the angle constant ε is perpendicular to the orbital plane of the satellite to be transferred to the reference. When the orbital altitude of the satellite to be transferred to the reference is not lower than that of the reference satellite, ε = a - θ; when the orbital altitude of the satellite to be transferred to the reference is lower than that of the reference satellite, ε = θ - α; Among them, n M Transfer the satellite orbital angular velocity to be referenced; T is the benchmark transfer time constraint; θ is the ground observation semi-cone angle of the load to be transferred by the benchmark; R M Indicates the orbital height of the satellite to be transferred as the reference; R J Indicates the orbital altitude of the reference star; When the orbit of the benchmark star is within the pyramid, it is considered that a benchmark transfer event has occurred. According to the benchmark transfer model and combined with mission requirements, the benchmark transfer efficiency between the benchmark satellite and the satellite to be benchmarked is simulated and analyzed, and the orbital inclination and ascending node right ascension parameters are traversed and optimized.

10. The space radiation measurement reference satellite orbit design and optimization system according to claim 6, characterized in that: The reference satellite in module 5 adopts a small eccentricity near-circular frozen orbit with perigee argument ω=90° and orbit eccentricity Where i is the orbital inclination; a is the semi-major axis of the orbit; R E is the radius of the Earth; J2 and J3 are both high-order perturbations due to the flat earth; In the module six, the benchmark satellite networking work requirements and orbital parameters are determined based on the payload characteristics and spatiotemporal distribution of the satellite to be benchmarked.

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