Method, device and equipment for determining satellite astronomical observation task parameters and medium

Abstract

This application discloses a method, apparatus, equipment, and medium for determining satellite astronomical observation mission parameters. Based on the satellite's orbital elements, the right ascension and declination of the target celestial body, and the right ascension and declination of the Sun, a first and second included angle corresponding to the satellite's current orbital period are calculated. Based on the first and second included angles, it is determined whether the satellite's current orbital period meets the observation conditions for the target celestial body. If the current orbital period meets the observation conditions for the target celestial body, the predicted observation mission parameters corresponding to the current orbital period are determined based on the satellite's operational parameters at the predicted start time of the current orbital period. Based on the predicted observation mission parameters and the dynamic satellite operational parameters of the current orbital period, the target observation mission parameters corresponding to the current orbital period are determined, effectively reducing the computational load while ensuring calculation accuracy.

Description

Technical Field

[0001] This disclosure generally relates to the field of satellite astronomical observation mission parameter calculation, and specifically relates to a method, apparatus, equipment and medium for determining satellite astronomical observation mission parameters. Background Technology

[0002] With the continuous development of satellite technology, small satellite constellations are being used more and more widely in astronomical observation. Small satellite constellations can achieve continuous monitoring of specific celestial bodies, rapid response to transient celestial bodies, and synchronous observation of large areas of the sky. As the number of satellites in a constellation continues to increase, the computational burden on the constellation is also constantly increasing. Many astronomical observation mission planning tasks that were originally conducted on the ground need to be gradually completed autonomously on-board. To solve the problem of the huge workload and inability to respond quickly on the ground, it is necessary to achieve autonomous on-board astronomical observation mission planning, thereby reducing manpower pressure while ensuring the smooth execution of missions.

[0003] For onboard autonomous astronomical observation missions, mission parameter calculation involves determining the start and end times of the astronomical observation mission based on the right ascension and declination of the target celestial body, and planning the mission configuration parameters. Due to limitations in the computing power of the onboard machine, the parameter calculation for onboard autonomous astronomical observation missions needs to be as efficient as possible while ensuring the required accuracy. Summary of the Invention

[0004] In view of the above-mentioned defects or deficiencies in the existing technology, it is desirable to provide a method, apparatus, equipment and medium for determining satellite astronomical observation mission parameters, which can effectively reduce the amount of computation while ensuring the accuracy of calculation.

[0005] In a first aspect, embodiments of this application provide a method for determining parameters of a satellite astronomical observation mission, including:

[0006] Based on the satellite's orbital elements, the right ascension and declination of the target celestial body, and the right ascension and declination of the Sun, calculate the first and second included angles corresponding to the satellite's current orbital period. The first included angle is the angle between the satellite's position vector and the target celestial body's position vector, and the second included angle is the angle between the target celestial body's position vector and the Sun's position vector.

[0007] Based on the first included angle and the second included angle, determine whether the current operating cycle of the satellite meets the observation conditions for the target celestial body;

[0008] When the current operating cycle of the satellite meets the observation conditions for the target celestial body, the estimated observation task parameters corresponding to the current operating cycle of the satellite are determined based on the satellite operating parameters at the predicted start time of the current operating cycle; the estimated observation task parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval;

[0009] Based on the estimated observation mission parameters and the dynamic satellite operation parameters of the current satellite operation cycle, the target observation mission parameters corresponding to the current satellite operation cycle are determined.

[0010] In some embodiments, when the current satellite operating cycle meets the observation conditions for the target celestial body, determining the estimated observation mission parameters corresponding to the current satellite operating cycle based on the satellite operating parameters at the predicted start time of the current satellite operating cycle includes:

[0011] When the current operating cycle of the satellite meets the observation conditions for the target celestial body, the satellite operating parameters at the predicted start time of the current operating cycle are used to determine the first latitude argument when the satellite enters the observation interval and the second latitude argument when it leaves the observation interval.

[0012] Based on the first latitude argument and the second latitude argument, determine the first time difference from the satellite's starting point of the prediction calculation to its entry into the observation interval and the second time difference from its exit from the observation interval.

[0013] In some embodiments, determining the first time difference from the satellite's entry into the observation interval to its departure from the prediction calculation starting point and the second time difference based on the first latitude argument and the second latitude argument includes:

[0014] Based on the first latitude argument and the second latitude argument, determine the first true anomaly angle when the satellite enters the observation range and the second true anomaly angle when it leaves the observation range;

[0015] The first true anomaly angle and the second true anomaly angle are converted into the first mean anomaly angle when the satellite enters the observation range and the second mean anomaly angle when it leaves the observation range;

[0016] Based on the mean aperimeter angle at the predicted calculation start time, calculate the difference in mean aperimeter angle from the predicted calculation start time to entering the observation interval and leaving the observation interval;

[0017] Based on the mean apogee angle difference and the mean angular velocity, the first time difference from the start of the prediction calculation to entering the observation interval and the second time difference to leaving the observation interval are calculated.

[0018] In some embodiments, it also includes:

[0019] If the first time difference between the satellite's start point of prediction calculation and its entry into the observation interval is greater than the second time difference between the satellite's start point of prediction calculation and its departure from the observation interval, the second time difference is updated based on the length of the satellite's current operating cycle.

[0020] In some embodiments, determining the target observation mission parameters corresponding to the current satellite operating cycle based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle includes:

[0021] Based on the first time difference and the second time difference, as well as the rate of change of right ascension of the ascending node and the rate of change of argument of perigee, calculate the right ascension of the first ascending node and the argument of the first perigee when entering the observation interval, and the right ascension of the second ascending node and the argument of the second perigee when leaving the observation interval;

[0022] Correction calculations are performed based on the right ascension of the first and second ascending nodes to obtain the first corrected amplitude and the first corrected phase when entering the observation interval, and the second corrected amplitude and the second corrected phase when leaving the observation interval.

[0023] Based on the first correction amplitude and the first correction phase, as well as the second correction amplitude and the second correction phase, the first correction latitude argument when entering the observation range and the second correction latitude argument when leaving the observation range are calculated.

[0024] Based on the first corrected latitude argument, the first perigee argument, the second corrected latitude argument, and the second perigee argument, the first corrected time difference from the prediction calculation start point to entering the observation interval and the second corrected time difference to leaving the observation interval are calculated.

[0025] Based on the predicted start time, the first correction time difference, and the second correction time difference, the first target time for entering the observation interval and the second target time for leaving the observation interval are determined.

[0026] In some embodiments, it also includes:

[0027] If the first correction time difference from the start of the prediction calculation to the entry into the observation interval is greater than the second correction time difference from the departure from the observation interval, the second correction time difference is updated according to the length of the current operating cycle of the satellite.

[0028] Secondly, embodiments of this application provide a device for determining parameters of a satellite astronomical observation mission, comprising:

[0029] The angle calculation module is used to calculate the first angle and the second angle corresponding to the current orbital period of the satellite based on the satellite orbital elements, the right ascension and declination of the target celestial body and the right ascension and declination of the sun. The first angle is the angle between the satellite position vector and the target celestial body position vector, and the second angle is the angle between the target celestial body position vector and the sun position vector.

[0030] The observation condition judgment module is used to determine whether the current operating cycle of the satellite meets the observation conditions for the target celestial body based on the first included angle and the second included angle.

[0031] The prediction parameter determination module is used to determine the predicted observation task parameters corresponding to the current satellite operating cycle when the current satellite operating cycle meets the observation conditions for the target celestial body, based on the satellite operating parameters at the predicted start time of the current satellite operating cycle; the predicted observation task parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval;

[0032] The target parameter determination module is used to determine the target observation mission parameters corresponding to the current satellite operating cycle based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle.

[0033] Thirdly, embodiments of this application provide an electronic device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the method described in embodiments of this application.

[0034] Fourthly, embodiments of this application provide a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the method described in embodiments of this application.

[0035] Fifthly, embodiments of this application provide a computer program product, including a computer program, characterized in that, when the computer program is executed by a processor, it implements the method described in embodiments of this application.

[0036] The method, apparatus, equipment, and medium for determining satellite astronomical observation mission parameters provided in this application calculate the first and second included angles corresponding to the current satellite operating cycle based on the satellite's orbital elements, the target celestial body, and the right ascension and declination of the Sun. Based on these first and second included angles, it is determined whether the current satellite operating cycle meets the observation conditions for the target celestial body. If the current satellite operating cycle meets the observation conditions, the estimated observation mission parameters corresponding to the current satellite operating cycle are determined based on the satellite operating parameters at the predicted start time of the current satellite operating cycle. The estimated observation mission parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval. Based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle, the target observation mission parameters corresponding to the current satellite operating cycle are determined. This reduces the computational load to less than one ten-thousandth of traditional methods while ensuring computational accuracy, making the planning of onboard autonomous astronomical observation missions theoretically feasible and engineering feasible.

[0037] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0038] Other features, objects, and advantages of this application will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0039] Figure 1 A flowchart illustrating a method for determining satellite astronomical observation mission parameters according to an embodiment of this application is shown.

[0040] Figure 2 A flowchart illustrating a method for determining satellite astronomical observation mission parameters according to another embodiment of this application is shown;

[0041] Figure 3 A schematic diagram of the structure of a device for determining satellite astronomical observation mission parameters according to an embodiment of this application is shown;

[0042] Figure 4 A schematic diagram of the structure of a computer system suitable for implementing an electronic device or server according to embodiments of this application is shown. Detailed Implementation

[0043] The present application will now be described in further detail with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and not intended to limit it. Furthermore, it should be noted that, for ease of description, only the parts relevant to the invention are shown in the accompanying drawings.

[0044] It should be noted that, unless otherwise specified, the embodiments and features described in this application can be combined with each other. This application will now be described in detail with reference to the accompanying drawings and embodiments.

[0045] To further illustrate the technical solutions provided in the embodiments of this application, a detailed description is provided below in conjunction with the accompanying drawings and specific implementation methods. Although the embodiments of this application provide method operation instruction steps as shown in the following embodiments or drawings, the method may include more or fewer operation instruction steps based on conventional or non-creative effort. In steps where there is no logically necessary causal relationship, the execution order of these steps is not limited to the execution order provided in the embodiments of this application. In actual processing or when the device executes the method, it may be executed sequentially or in parallel according to the method shown in the embodiments or drawings.

[0046] Please refer to Figure 1 and Figure 2 , Figure 1A flowchart illustrating a method for determining satellite astronomical observation mission parameters according to an embodiment of this application is shown. Figure 1 As shown, the method includes:

[0047] Step 101: Based on the satellite orbital elements, the right ascension and declination of the target celestial body, and the right ascension and declination of the Sun, calculate the first and second included angles corresponding to the current satellite orbital period. The first included angle is the angle between the satellite position vector and the target celestial body position vector, and the second included angle is the angle between the target celestial body position vector and the Sun position vector.

[0048] It should be noted that since satellite orbital information is usually expressed in the form of position, velocity, and orbital elements rather than right ascension and declination, it is necessary to convert the satellite's orbital elements to right ascension and declination. Specifically, the satellite's right ascension... and declination It can be obtained by transforming a spherical triangle:

[0049]

[0050] in, The orbital inclination is the orbital element of a satellite. Right ascension of the ascending node, Argument of latitude , The perigee argument, It is the true near point angle.

[0051] For any two celestial bodies in space, their right ascension and declination are determined by... and The angle between their position vectors It can be calculated using the cosine formula from spherical trigonometry:

[0052]

[0053] The satellite's right ascension and declination Right ascension and declination of the target celestial body The Sun's right ascension and declination Substituting into the above formula, we get:

[0054]

[0055]

[0056] in, The first angle formed by the satellite's position vector and the target celestial body's position vector. The second angle formed by the target celestial body's position vector and the sun's position vector.

[0057] Step 102: Determine whether the satellite's current operating cycle meets the observation conditions for the target celestial body based on the first included angle and the second included angle.

[0058] It should be noted that, due to the angular limitations of astronomical observation payloads on the Earth and the Sun, such as the observation target needing to be at an angle greater than 60° with the Sun, it is necessary to determine whether the observation conditions are met before the autonomous mission planning calculation.

[0059] Regarding the first included angle, although the right ascension and declination of the target celestial body remain essentially constant within one satellite orbital period, the right ascension and declination change rapidly during the satellite's motion within one orbital period. Therefore, it is necessary to calculate the first included angle. The minimum value within a period is used to determine whether there is an observation window that meets the conditions.

[0060] To calculate the minimum value, first substitute the right ascension and declination of the satellite and the target celestial body into the aforementioned formula for calculating the included angle. Through trigonometric identity expansion and combining like terms, a first-order harmonic equation relating only to u is obtained:

[0061]

[0062] in: , , .

[0063] Furthermore, amplitude R and phase The expression is as follows:

[0064]

[0065] It should be understood that the amplitude R is... The minimum value.

[0066] Astronomical observations require that the first angle between the satellite's position vector and the target celestial body's position vector be satisfied. (Right now ), then: if This indicates that throughout the entire cycle The maximum value is still less than the threshold, i.e., the whole cycle. All greater than If the current period does not meet the Earth's angle limitation, astronomical observations of the target celestial body are not possible. Conversely, if the current period does meet the conditions for observing the target celestial body, then the conditions for observation are met.

[0067] In some embodiments, it is also necessary to determine the amplitude. Does it approach 0? If so... If the value approaches 0, then the target celestial body's position vector is perpendicular to the satellite's orbital plane during this period. In this case, the target celestial body remains within the satellite's astronomical observation range throughout the current cycle and can be observed continuously without further calculations.

[0068] For the second included angle, since the Sun's right ascension and declination change over a period of one year and at an extremely slow rate, the calculation for each period is directly used. You can make a judgment.

[0069] Assume that astronomical observations need to meet the following requirements. (Right now ), then: if That is, the entire cycle All less than or equal to If the solar angle is not met during this period, astronomical observations cannot be conducted. Conversely, if the solar angle is met, the conditions for observing the target celestial body are satisfied.

[0070] Step 103: When the satellite's current operating cycle meets the observation conditions for the target celestial body, determine the estimated observation mission parameters corresponding to the current operating cycle of the satellite based on the predicted start time of the satellite's current operating cycle calculation. The estimated observation mission parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval.

[0071] In other words, when the satellite's current operating cycle meets the observation conditions for the target celestial body, the observation mission parameters of the current operating cycle can be further analyzed to facilitate observation and control of the satellite within the current cycle. These observation mission parameters include the time when the satellite enters and leaves the observation interval during its current operating cycle. The predicted calculation start time can be a preset start time used for calculating the observation mission parameters, such as the time when the satellite enters its current operating cycle; this application does not specifically limit this.

[0072] In a feasible embodiment, when the current operating cycle of the satellite meets the observation conditions for the target celestial body, the first latitude argument when the satellite enters the observation interval and the second latitude argument when it leaves the observation interval are determined based on the satellite operating parameters at the predicted calculation start time of the current operating cycle; based on the first latitude argument and the second latitude argument, the first time difference from the predicted calculation start time to entering the observation interval and the second time difference to leaving the observation interval are determined.

[0073] In other words, in this embodiment of the application, the determined angle restriction information is first converted into the latitude amplitude interval boundary, and then the estimated first time difference for the satellite to enter the observation interval and the second time difference to leave the observation interval are determined by the latitude amplitude interval boundary.

[0074] Specifically, for a satellite orbit that meets the observation conditions, R ≥ |cos(θ) t_limit|, the first latitude argument when the satellite enters the observation range and the second latitude argument when leaving the observation interval It can be represented as follows:

[0075]

[0076] Where γ is the half-chord angle of the observation interval, .

[0077] Further, based on the first and second latitude argument angles, the first time difference from the satellite's entry into the observation interval and the second time difference from its departure from the observation interval are determined, including: determining the first true anomaly angle when the satellite enters the observation interval and the second true anomaly angle when it leaves the observation interval based on the first and second latitude argument angles; converting the first and second true anomaly angles into the first mean anomaly angle when the satellite enters the observation interval and the second mean anomaly angle when it leaves the observation interval; calculating the mean anomaly angle difference from the prediction calculation start point to the entry and departure from the observation interval based on the mean anomaly angle at the prediction calculation start point; and calculating the first time difference from the satellite's entry into the observation interval and the second time difference from its departure from the observation interval based on the mean anomaly angle difference and the average angular velocity.

[0078] Specifically, based on the first latitude argument and the second latitude argument, and combined with the perigee argument, the first true anomaly angle when entering the observation range and the second true anomaly angle when leaving the observation range are obtained:

[0079]

[0080] in, The first true anomaly angle when entering the observation range. It is the second true anterior angle when leaving the observation interval.

[0081] Convert the first true anomaly and the second true anomaly into the first mean anomaly when entering the observation interval and the second mean anomaly when leaving the observation interval:

[0082]

[0083]

[0084] in, The first approximation angle when entering the observation range. The second approximation angle when leaving the observation interval. The orbital eccentricity of the satellite, The mean anterior angle when entering the observation interval. It is the mean anterior angle when leaving the observation interval.

[0085] The mean angle corresponding to the predicted start time within the current satellite operating cycle is calculated. Calculate the mean anterior angle difference from the prediction calculation starting point to the point of entry into and exit from the observation interval:

[0086]

[0087] in, This is the difference in the mean anterior angle when entering the observation interval. This represents the difference in the mean aperimeter angle when leaving the observation interval.

[0088] It should be noted that mod 2π represents taking the value modulo 2π. Solving for ΔM yields... in and ΔM out Then, ΔM in and ΔM out All are normalized to between 0 and 2π.

[0089] Based on the mean angular difference and the mean angular velocity, calculate the first time difference from the prediction calculation start point to entering the observation interval. and the second time difference from leaving the observation interval :

[0090]

[0091] in, The average angular velocity, The gravitational constant of Earth, For the satellite's semi-major axis, For the second-order zonal harmonic coefficients of the Earth's gravitational field perturbation, This is the radius of the Earth's equator.

[0092] In some embodiments, This means that the satellite has passed the start point of the observation window at its current position, "crossed" the boundary of its operational cycle, and entered the observation range within the current operational cycle, but will have to wait until the next operational cycle to leave the observation range. In this case, A complete orbital period needs to be added to accurately reflect the time required from the start of the prediction calculation to leaving the observation interval. Without this correction, it would lead to… This contradicts the physical fact that entry precedes exit.

[0093] Specifically, if the first time difference between the satellite's start point of prediction and its entry into the observation interval is greater than the second time difference between the satellite's start point of prediction and its departure from the observation interval, the second time difference is updated based on the length of the satellite's current operating cycle.

[0094] When the first time difference is greater than the second time difference, the second time difference is updated, that is, the original second time difference is added to the orbital period.

[0095]

[0096] Step 104: Determine the target observation mission parameters corresponding to the current satellite operating cycle based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle.

[0097] It should be understood that the estimated parameters are obtained based on the satellite's operating parameters at the predicted calculation start time; however, the satellite's operating parameters are dynamic. Therefore, this application's embodiment introduces orbital perturbation correction to eliminate time calculation errors caused by the drift of the right ascension Ω of the ascending node and the argument ω of perigee during flight, outputting high-precision target observation mission parameters that can be directly executed by the satellite.

[0098] In a feasible embodiment, the target observation mission parameters corresponding to the current satellite operating cycle are determined based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle. This includes: calculating the right ascension of the first ascending node and the first perigee argument when entering the observation interval, and the right ascension of the second ascending node and the second perigee argument when leaving the observation interval, based on the first time difference and the second time difference, as well as the rate of change of the right ascension of the ascending node and the rate of change of the argument of the perigee; and performing correction calculations based on the right ascension of the first ascending node and the right ascension of the second ascending node to obtain the first corrected amplitude and the first corrected phase when entering the observation interval, and the second corrected phase when leaving the observation interval. The first correction amplitude and the second correction phase are used to calculate the first correction latitude argument when entering the observation interval and the second correction latitude argument when leaving the observation interval, based on the first correction latitude argument, the first perigee argument, the second correction latitude argument and the second perigee argument. The first correction time difference from the prediction calculation start point to entering the observation interval and the second correction time difference to leaving the observation interval are calculated based on the prediction calculation start time, the first correction time difference and the second correction time difference. The first target time to enter the observation interval and the second target time to leave the observation interval are determined based on the prediction calculation start time, the first correction time difference and the second correction time difference.

[0099] Specifically, for satellite orbital parameters, the semi-major axis *a*, eccentricity *e*, and orbital inclination *i* remain essentially constant throughout one orbital period. However, the right ascension of the ascending node *Ω* and the argument of perigee *ω* vary to some extent, affecting the calculation of the mean perigee angle *M* and the time difference. Therefore, correction calculations are required. The right ascension of the ascending node *Ω* is calculated for correction when the satellite enters and leaves the observation interval. in and Ω out Argument of perigee ω in and ω outThe expression is shown below.

[0100]

[0101] The right ascension Ω of the ascending node calculated by correction in and Ω out Substituting into the equation of step S102, the corresponding amplitude is obtained through correction calculation. and Phase and Then, the latitude argument was corrected and calculated when entering and leaving the observation interval. , Its expression is shown below.

[0102]

[0103] Correct the calculated latitude angle , and perigee argument , Substituting the values ​​into the conversion formulas for true anomaly / off-center anomaly / mean anomaly in step S3, the calculation of the mean anomaly difference, and the time difference are used to correct and calculate the time difference between the satellite entering the observation interval and the starting point. and the time difference between the time of leaving the observation interval and the time of the starting point The final calculated time to enter the observation interval. and the time of leaving the observation interval for:

[0104]

[0105] Where t is the starting time of the prediction calculation within this operating cycle, and dt inc dt outc The expression is shown below.

[0106]

[0107] Based on the final calculated entry time t of the observation interval inc and the time t of departure from observation outc The satellite's mission planning system autonomously plans and executes the observation start point and duration for the current operational cycle based on the time results.

[0108] Furthermore, if the first correction time difference from the start of the prediction calculation to the entry into the observation interval is greater than the second correction time difference from the departure from the observation interval, the second correction time difference is updated according to the length of the satellite's current operating cycle.

[0109] Considering Δt in >Δt outThis means that there are situations where a satellite enters the observation interval during its operational cycle but does not leave the observation interval until the next operational cycle. Therefore, when the calculated Δt... in >Δt out At that time, Δt out The value is updated to the original Δt. out Add cycle

[0110] Therefore, the rapid calculation method for satellite autonomous astronomical observation mission parameters provided in this application converts satellite orbital elements into right ascension and declination and then uses spherical trigonometry to calculate the included angles. It achieves analytical determination of observation conditions for the entire operating cycle through amplitude analysis of the first-order harmonic equation and realizes high-precision calculation of observation window time through a two-stage "prediction-correction" architecture. While ensuring calculation accuracy, it reduces the amount of calculation to less than one ten-thousandth of that of traditional methods, making the planning of on-board autonomous astronomical observation missions not only theoretically feasible but also engineering feasible.

[0111] In one specific embodiment, the center of the Milky Way, Cygnus X-1, and the Crab Nebula were selected as astronomical observation targets, with right ascension and declination of (266.41°, -29.0°), (299.59°, +35.2°), and (83.63°, +22.0°), respectively. During observation, celestial objects with an angle greater than 120° to the Earth's center (i.e., an angle less than 60° to the satellite's position vector) were preferred, and these objects should also have an angle greater than 60° to the Sun. For the test satellite, the epoch time was 00:00:00 GMT on January 1, 2026, and the satellite orbit was a sun-synchronous orbit with an altitude of 561 km and local times at the descending node of 6:00 am and 12:00 pm. For satellites with a descending node local time of 6:00 AM, GMT 19:00:03 on January 1, 2026 was selected as the starting point for prediction calculation; for satellites with a descending node local time of 12:00 PM, GMT 9:24:02 on January 1, 2026 was selected as the starting point for prediction calculation. Table 1 shows a comparison between the results of the satellite's autonomous astronomical observation prediction calculations and the actual verification results.

[0112] Table 1 Comparison of Satellite Autonomous Astronomical Observation Prediction Calculation Results

[0113]

[0114] The entire calculation process was performed on an i9, 2.2GHz computer hardware environment and a Windows 11 64-bit system software environment, using Visual Studio 2022, with the compiler set to disable optimization ( / Od), and the program took 0.0000185 seconds.

[0115] Practical application results show that the satellite's autonomous astronomical observation algorithm has high accuracy, speed and efficiency, and fully meets the needs of actual onboard missions.

[0116] It should be noted that although the operation of the method of the present invention is described in a specific order in the accompanying drawings, this does not require or imply that the operations must be performed in that specific order, or that all the operations shown must be performed in order to achieve the desired result.

[0117] Figure 3 A schematic diagram of the structure of a device for determining satellite astronomical observation mission parameters provided in an embodiment of this application is shown.

[0118] like Figure 3 As shown, the device 10 for determining the parameters of the satellite astronomical observation mission includes:

[0119] The included angle calculation module 11 is used to calculate the first included angle and the second included angle corresponding to the current operating cycle of the satellite based on the satellite orbit elements, the right ascension and declination of the target celestial body and the right ascension and declination of the sun. The first included angle is the angle between the satellite position vector and the target celestial body position vector, and the second included angle is the angle between the target celestial body position vector and the sun position vector.

[0120] The observation condition judgment module 12 is used to determine whether the current operating cycle of the satellite meets the observation conditions for the target celestial body based on the first included angle and the second included angle.

[0121] The prediction parameter determination module 13 is used to determine the prediction observation task parameters corresponding to the current satellite operating cycle based on the satellite operating parameters at the predicted start time of the current satellite operating cycle when the current satellite operating cycle meets the observation conditions for the target celestial body; the prediction observation task parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval;

[0122] The target parameter determination module 14 is used to determine the target observation task parameters corresponding to the current operating cycle of the satellite based on the estimated observation task parameters and the dynamic satellite operating parameters of the current operating cycle of the satellite.

[0123] In some embodiments, the parameter estimation module 13 is specifically used for:

[0124] When the current operating cycle of the satellite meets the observation conditions for the target celestial body, the satellite operating parameters at the predicted start time of the current operating cycle are used to determine the first latitude argument when the satellite enters the observation interval and the second latitude argument when it leaves the observation interval.

[0125] Based on the first latitude argument and the second latitude argument, determine the first time difference from the satellite's starting point of the prediction calculation to its entry into the observation interval and the second time difference from its exit from the observation interval.

[0126] In some embodiments, the parameter estimation module 13 is specifically used for:

[0127] Based on the first latitude argument and the second latitude argument, determine the first true anomaly angle when the satellite enters the observation range and the second true anomaly angle when it leaves the observation range;

[0128] The first true anomaly angle and the second true anomaly angle are converted into the first mean anomaly angle when the satellite enters the observation range and the second mean anomaly angle when it leaves the observation range;

[0129] Based on the mean aperimeter angle at the predicted calculation start time, calculate the difference in mean aperimeter angle from the predicted calculation start time to entering the observation interval and leaving the observation interval;

[0130] Based on the mean apogee angle difference and the mean angular velocity, the first time difference from the start of the prediction calculation to entering the observation interval and the second time difference to leaving the observation interval are calculated.

[0131] In some embodiments, the parameter estimation module 13 is further configured to:

[0132] If the first time difference between the satellite's start point of prediction calculation and its entry into the observation interval is greater than the second time difference between the satellite's start point of prediction calculation and its departure from the observation interval, the second time difference is updated based on the length of the satellite's current operating cycle.

[0133] In some embodiments, the target parameter determination module 14 is specifically used for:

[0134] Based on the first time difference and the second time difference, as well as the rate of change of right ascension of the ascending node and the rate of change of argument of perigee, calculate the right ascension of the first ascending node and the argument of the first perigee when entering the observation interval, and the right ascension of the second ascending node and the argument of the second perigee when leaving the observation interval;

[0135] Substituting the right ascension of the first ascending node and the right ascension of the second ascending node into the harmonic equation, the first corrected amplitude and the first corrected phase when entering the observation interval, and the second corrected amplitude and the second corrected phase when leaving the observation interval are calculated.

[0136] Based on the first correction amplitude and the first correction phase, as well as the second correction amplitude and the second correction phase, the first correction latitude argument when entering the observation range and the second correction latitude argument when leaving the observation range are calculated.

[0137] Based on the first corrected latitude argument, the first perigee argument, the second corrected latitude argument, and the second perigee argument, the first corrected time difference from the prediction calculation start point to entering the observation interval and the second corrected time difference to leaving the observation interval are calculated.

[0138] Based on the predicted start time, the first correction time difference, and the second correction time difference, the first target time for entering the observation interval and the second target time for leaving the observation interval are determined.

[0139] In some embodiments, the target parameter determination module 14 is specifically used for:

[0140] If the first correction time difference from the start of the prediction calculation to the entry into the observation interval is greater than the second correction time difference from the departure from the observation interval, the second correction time difference is updated according to the length of the current operating cycle of the satellite.

[0141] It should be understood that the modules or modules described in the satellite astronomical observation mission parameter determination device 10 are related to the reference... Figure 1 The steps in the described method correspond accordingly. Therefore, the operations and features described above for the method are also applicable to the satellite astronomical observation mission parameter determination device 10 and its included modules, and will not be repeated here. The satellite astronomical observation mission parameter determination device 10 can be pre-implemented in the browser or other secure applications of an electronic device, or it can be loaded into the browser or its secure applications of an electronic device through download or other means. The corresponding modules in the satellite astronomical observation mission parameter determination device 10 can cooperate with the modules in the electronic device to implement the scheme of the embodiments of this application.

[0142] The division of modules or units mentioned in the detailed description above is not mandatory. In fact, according to the embodiments of this disclosure, the features and functions of two or more modules or units described above can be embodied in one module or unit. Conversely, the features and functions of one module or unit described above can be further divided and embodied by multiple modules or units.

[0143] The following is for reference. Figure 4 , Figure 4 A schematic diagram of the structure of a computer system suitable for implementing the embodiments of this application is shown.

[0144] like Figure 4As shown, the computer system 400 includes a central processing unit (CPU) 401, which can perform various appropriate actions and processes based on programs stored in read-only memory (ROM) 402 or programs loaded from storage section 408 into random access memory (RAM) 403. RAM 403 also stores various programs and data required for the system's operating instructions. CPU 401, ROM 402, and RAM 404 are interconnected via bus 403. Input / output (I / O) interface 405 is also connected to bus 404.

[0145] The following components are connected to I / O interface 405: an input section 406 including a keyboard, mouse, etc.; an output section 407 including a cathode ray tube (CRT), liquid crystal display (LCD), etc., and speakers, etc.; a storage section 408 including a hard disk, etc.; and a communication section 409 including a network interface card such as a LAN card, modem, etc. The communication section 409 performs communication processing via a network such as the Internet. Drive 410 is also connected to I / O interface 405 as needed. Removable media 411, such as a disk, optical disk, magneto-optical disk, semiconductor memory, etc., are installed on drive 410 as needed so that computer programs read from them can be installed into storage section 408 as needed.

[0146] Specifically, according to embodiments of this application, the processes described above with reference to the flowcharts can be implemented as computer software programs. For example, embodiments of this application include a computer program product comprising a computer program carried on a computer-readable medium, the computer program containing program code for performing the methods shown in the flowcharts. In such embodiments, the computer program contains program code for performing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network via communication section 409, and / or installed from removable medium 411. When the computer program is executed by central processing unit (CPU) 401, it performs the functions defined above in the system of this application.

[0147] It should be noted that the computer-readable medium shown in this application can be a computer-readable signal medium or a computer-readable storage medium, or any combination of the two. A computer-readable storage medium can be, for example,—but not limited to—an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples of a computer-readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination thereof. In this application, a computer-readable storage medium can be any tangible medium containing or storing a program that can be used by or in conjunction with an instruction execution system, apparatus, or device. In this application, a computer-readable signal medium can include a data signal propagated in baseband or as part of a carrier wave, carrying computer-readable program code. Such propagated data signals can take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. Computer-readable signal media can also be any computer-readable medium other than computer-readable storage media, which can send, propagate, or transmit a program for use by or in connection with an instruction execution system, apparatus, or device. The program code contained on the computer-readable medium can be transmitted using any suitable medium, including but not limited to: wireless, wire, optical fiber, RF, etc., or any suitable combination thereof.

[0148] The flowcharts and block diagrams in the accompanying drawings illustrate the architecture, functionality, and operational instructions of possible implementations of systems, methods, and computer program products according to various embodiments of this application. In this regard, each block in a flowchart or block diagram may represent a module, segment, or portion of code containing one or more executable instructions for implementing a specified logical function. It should also be noted that in some alternative implementations, the functions indicated in the blocks may occur in a different order than those indicated in the drawings. For example, two connected blocks may actually be executed substantially in parallel, or they may sometimes be executed in reverse order, depending on the functions involved. It should also be noted that each block in the block diagrams and / or flowcharts, and combinations of blocks in the block diagrams and / or flowcharts, can be implemented using a dedicated hardware-based system that performs the specified functions or operational instructions, or using a combination of dedicated hardware and computer instructions.

[0149] The units or modules described in the embodiments of this application can be implemented in software or hardware. The described units or modules can also be housed in a processor; for example, a processor can be described as including an angle calculation module, an observation condition judgment module, a predicted parameter determination module, and a target parameter determination module. The names of these units or modules do not necessarily limit the specific unit or module itself. For example, the angle calculation module can also be described as "calculating the first and second angles corresponding to the current orbital period of the satellite based on the satellite orbital elements, the right ascension and declination of the target celestial body and the Sun."

[0150] In another aspect, this application also provides a computer-readable storage medium, which may be included in the electronic device described in the above embodiments, or may exist independently and not assembled into the electronic device. The aforementioned computer-readable storage medium stores one or more programs that, when used by one or more processors, execute the method for determining satellite astronomical observation mission parameters described in this application.

[0151] The above description is merely a preferred embodiment of this application and an explanation of the technical principles employed. Those skilled in the art should understand that the scope of disclosure in this application is not limited to technical solutions formed by specific combinations of the above-described technical features, but should also cover other technical solutions formed by arbitrary combinations of the above-described technical features or their equivalents without departing from the foregoing disclosed concept. For example, technical solutions formed by substituting the above features with (but not limited to) technical features with similar functions disclosed in this application.

Claims

1. A method for determining parameters of a satellite astronomical observation mission, characterized in that, include: Based on the satellite's orbital elements, the right ascension and declination of the target celestial body, and the right ascension and declination of the Sun, calculate the first and second included angles corresponding to the satellite's current orbital period. The first included angle is the angle between the satellite's position vector and the target celestial body's position vector, and the second included angle is the angle between the target celestial body's position vector and the Sun's position vector. Based on the first included angle and the second included angle, determine whether the current operating cycle of the satellite meets the observation conditions for the target celestial body; When the current operating cycle of the satellite meets the observation conditions for the target celestial body, the estimated observation task parameters corresponding to the current operating cycle of the satellite are determined based on the satellite operating parameters at the predicted start time of the current operating cycle; the estimated observation task parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval; Based on the estimated observation mission parameters and the dynamic satellite operation parameters of the current satellite operation cycle, the target observation mission parameters corresponding to the current satellite operation cycle are determined.

2. The method for determining satellite astronomical observation mission parameters according to claim 1, characterized in that, When the satellite's current operating cycle meets the observation conditions for the target celestial body, the estimated observation mission parameters corresponding to the current operating cycle are determined based on the satellite operating parameters at the predicted start time of the current operating cycle, including: When the current operating cycle of the satellite meets the observation conditions for the target celestial body, the satellite operating parameters at the predicted start time of the current operating cycle are used to determine the first latitude argument when the satellite enters the observation interval and the second latitude argument when it leaves the observation interval. Based on the first latitude argument and the second latitude argument, determine the first time difference from the satellite's starting point of the prediction calculation to its entry into the observation interval and the second time difference from its exit from the observation interval.

3. The method for determining satellite astronomical observation mission parameters according to claim 2, characterized in that, Based on the first latitude argument and the second latitude argument, the first time difference from the satellite's entry into the observation interval to its departure from the prediction calculation starting point and the second time difference from its entry into the observation interval are determined, including: Based on the first latitude argument and the second latitude argument, determine the first true anomaly angle when the satellite enters the observation range and the second true anomaly angle when it leaves the observation range; The first true anomaly angle and the second true anomaly angle are converted into the first mean anomaly angle when the satellite enters the observation range and the second mean anomaly angle when it leaves the observation range; Based on the mean aperimeter angle at the predicted calculation start time, calculate the difference in mean aperimeter angle from the predicted calculation start time to entering the observation interval and leaving the observation interval; Based on the mean apogee angle difference and the mean angular velocity, the first time difference from the start of the prediction calculation to entering the observation interval and the second time difference to leaving the observation interval are calculated.

4. The method for determining satellite astronomical observation mission parameters according to claim 2, characterized in that, Also includes: If the first time difference between the satellite's start point of prediction calculation and its entry into the observation interval is greater than the second time difference between the satellite's start point of prediction calculation and its departure from the observation interval, the second time difference is updated based on the length of the satellite's current operating cycle.

5. The method for determining satellite astronomical observation mission parameters according to claim 1, characterized in that, The step of determining the target observation mission parameters corresponding to the current satellite operating cycle based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle includes: Based on the first time difference and the second time difference, as well as the rate of change of right ascension of the ascending node and the rate of change of argument of perigee, calculate the right ascension of the first ascending node and the argument of the first perigee when entering the observation interval, and the right ascension of the second ascending node and the argument of the second perigee when leaving the observation interval; Correction calculations are performed based on the right ascension of the first and second ascending nodes to obtain the first corrected amplitude and the first corrected phase when entering the observation interval, and the second corrected amplitude and the second corrected phase when leaving the observation interval. Based on the first correction amplitude and the first correction phase, as well as the second correction amplitude and the second correction phase, the first correction latitude argument when entering the observation range and the second correction latitude argument when leaving the observation range are calculated. Based on the first corrected latitude argument, the first perigee argument, the second corrected latitude argument, and the second perigee argument, the first corrected time difference from the prediction calculation start point to entering the observation interval and the second corrected time difference to leaving the observation interval are calculated. Based on the predicted start time, the first correction time difference, and the second correction time difference, the first target time for entering the observation interval and the second target time for leaving the observation interval are determined.

6. The method for determining satellite astronomical observation mission parameters according to claim 5, characterized in that, Also includes: If the first correction time difference from the start of the prediction calculation to the entry into the observation interval is greater than the second correction time difference from the departure from the observation interval, the second correction time difference is updated according to the length of the current operating cycle of the satellite.

7. A device for determining parameters of a satellite astronomical observation mission, characterized in that, include: The angle calculation module is used to calculate the first angle and the second angle corresponding to the current orbital period of the satellite based on the satellite orbital elements, the right ascension and declination of the target celestial body and the right ascension and declination of the sun. The first angle is the angle between the satellite position vector and the target celestial body position vector, and the second angle is the angle between the target celestial body position vector and the sun position vector. The observation condition judgment module is used to determine whether the current operating cycle of the satellite meets the observation conditions for the target celestial body based on the first included angle and the second included angle. The prediction parameter determination module is used to determine the predicted observation task parameters corresponding to the current satellite operating cycle based on the satellite operating parameters at the predicted start time of the current satellite operating cycle when the current satellite operating cycle meets the observation conditions for the target celestial body; the predicted observation task parameters include at least the time when the satellite enters the observation interval and the time when it leaves the observation interval; The target parameter determination module is used to determine the target observation mission parameters corresponding to the current satellite operating cycle based on the estimated observation mission parameters and the dynamic satellite operating parameters of the current satellite operating cycle.

8. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the program, it implements the method for determining satellite astronomical observation mission parameters as described in any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by the processor, the program implements the method for determining satellite astronomical observation mission parameters as described in any one of claims 1-6.

10. A computer program product, comprising a computer program, characterized in that, When executed by a processor, the computer program implements the method for determining the parameters of a satellite astronomical observation mission as described in any one of claims 1-6.

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