Ground detection platform simulation track acceleration method and device based on six tracks

Abstract

The invention discloses a ground detection platform simulation orbit acceleration method and device based on six orbit elements, and the method comprises the steps: enabling a simulation satellite platform to transmit a broadcast data packet to a load through a data bus in a load ground test process; the load central controller inputs the received broadcast data packet into an orbit acceleration system for analysis, and calculates the satellite position; judging whether to accelerate according to the satellite position; controlling orbital acceleration to be finished according to whether the satellite reaches the hotspot position; and when the track acceleration is finished, judging whether the hot spot detection is finished. According to the method, six orbit elements and time are used for representing an operation curve of a whole orbit to determine a satellite operation orbit, a load operation scene is divided according to load operation time, if the orbit is in a non-operation state, a system can calculate a next operation time point, corresponding parameters are imported to carry out simulation test on the load, and an orbit simulation process is accelerated; in the redundant time of unnecessary operation of some loads, acceleration can be carried out, so that a simulation scene of next operation can be reached more quickly.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite orbit simulation, and in particular to a method and device for simulating orbit acceleration of a ground detection platform based on six orbit elements. Background Art

[0002] The high-resolution satellite-borne hotspot monitor monitors greenhouse gas emissions in hotspot areas around the world in orbit. It uses nadir observation. During its in-orbit operation, the hotspot monitor receives the broadcast time packet sent by the satellite platform every second. The broadcast time packet contains the satellite's orbit information. The monitor adjusts the pointing mirror angle based on the orbit information, aims at the hotspot area, and then takes pictures. During the in-orbit operation of the payload, the entire orbit does not work continuously. In areas with low hotspots or dark nights under the satellite, the payload is in standby mode.

[0003] During the development of the payload, a ground test platform is needed to simulate the satellite platform and conduct ground tests on the payload to ensure that the payload can operate normally while in orbit. During the ground debugging of the payload, if it runs completely according to the real orbit, it will take a long time and the debugging efficiency will be low.

[0004] The equipment used for ground debugging of payloads is called a ground test platform. Satellites are divided into two main components: satellite platform and payload. The platform and payload communicate according to a unified protocol. During the development of the payload, the ground test platform simulates the satellite platform and conducts ground tests on the payload to verify whether the payload functions are realized, eliminate other safety hazards, and ensure its normal operation in orbit.

[0005] The ground detection platform needs to measure and control the payload and receive business data. The measurement and control platform uses the measurement and control bus 1553B. The communication content of 1553B includes internal instructions, data injection, telemetry information, and broadcast time packets, etc. Measurement and control is generally called "digital communication" and its main contents include direct instructions, indirect instructions, direct telemetry, and indirect telemetry, etc. Business data is generally called "digital transmission communication" and its main contents include data transmission and reception of broadcast signals.

[0006] Since the orbit is divided into different scenarios according to the payload operation time, the satellite will not work in some scenarios, such as when the satellite enters the night or non-hotspot area. During the test, the payload should reduce the remaining redundant time as much as possible after the necessary operation time, so orbit acceleration is designed as a method to reduce redundant time.

[0007] The existing method is to use orbit simulation software such as STK in advance to create orbit data for simulating satellite operation in orbit, and then have the ground inspection platform parse and send it. However, due to the large size of STK software, the software interface is difficult to call, the calculation time is long, and the output data cannot be used directly, and the data needs to be processed again.

[0008] The present invention discloses a method for simulating orbital acceleration of a ground detection platform based on six orbital numbers. During ground test of payload, the six orbital numbers and time are used to determine the satellite orbit, and the payload operation scenes are divided according to the payload operation time. When the payload is in a non-operating time period, the orbital simulation process is accelerated. The next moment when the payload is about to operate is quickly reached, thereby reducing the redundant time during payload simulation. (Hereinafter referred to as: orbital acceleration) This algorithm is customized according to the working status of the satellite and payload, can be embedded in other programs, is easy to operate, and has good real-time performance. Summary of the invention

[0009] The present invention proposes a method, device and storage medium for simulating track acceleration on a ground detection platform based on six track numbers, which can solve at least one of the technical problems in the background technology.

[0010] To achieve the above object, the present invention adopts the following technical solutions:

[0011] A method for simulating track acceleration on a ground detection platform based on six track numbers, the specific steps are as follows:

[0012] S1, simulates the satellite platform sending broadcast data packets to the payload through the data bus;

[0013] S2, the payload central control enters the received broadcast data packet into the orbit acceleration system for analysis and calculates the satellite position;

[0014] S3, judging whether to accelerate according to the satellite position;

[0015] S4, controlling whether orbit acceleration ends according to whether the satellite reaches the hotspot position;

[0016] S5. When track acceleration is finished, determine whether hotspot detection is completed.

[0017] Furthermore, the track acceleration system in S2 of the present invention includes a simulated track visualization module, a number injection module, and an acceleration judgment module;

[0018] The injection module converts the time data in the broadcast data packet input by the system into JD time, and transmits it to the acceleration judgment module together with the six track number data;

[0019] The acceleration judgment module calculates the satellite position based on JD time and six orbital numbers and determines whether the satellite needs to be accelerated, and inputs the processed data into the visualization module.

[0020] Furthermore, the satellite position calculation method in S2 of the present invention is as follows:

[0021] S21, introducing auxiliary parameters to calculate the true anomaly angle;

[0022] Two auxiliary parameters, the eccentric anomaly angle E and the mean anomaly angle M, are introduced. M is an assumed quantity, then M = n(t-t0), where t0 is the time when the satellite passes the perigee, and t is the time when the satellite is observed.

[0023] The relationship between the mean anomaly angle and the eccentric anomaly angle is E = M + e sin E, so the cosine of the true anomaly angle is:

[0024]

[0025] S22, calculate the celestial coordinates of the satellite at any time by combining the true anomaly angle;

[0026] The origin of the rectangular coordinate system is taken to coincide with the center of mass of the earth, the X-axis points to the perigee, the Z-axis is perpendicular to the orbital plane and points upward, and the Y-axis is perpendicular to the X-axis on the orbital plane to form a right-hand system. The celestial coordinates of the satellite at any time are:

[0027]

[0028] Where r = a(1-ecosE);

[0029] S23, converting the celestial coordinates of the satellite at any time into geodetic coordinates;

[0030] The method to convert the satellite's celestial coordinates at any time into geodetic coordinates is as follows:

[0031]

[0032] The precession matrix is:

[0033]

[0034] in, θ is the rotation angle, Ry is the coordinate transformation matrix around the Y axis, and Rz is the coordinate transformation matrix around the Z axis;

[0035] Among them, time T is the naturalized time calculated from the J2000.0 epoch, that is,

[0036]

[0037] where JD(TDB) represents the Julian day expressed by the barycentric dynamical time (TDB);

[0038] The nutation matrix is:

[0039] [NR]=R x (-ε s -Δε)·R z (-ΔΨ)·R x (ε s )

[0040] where Δε and Ψ are the nutation of the ecliptic longitude and the nutation of the intersection angle respectively; ε s is the obliquity of the mean ecliptic, Rx is the coordinate transformation matrix around the X axis;

[0041] Day and night rotation matrix:

[0042] [ER]=R y (θ G )

[0043] in,

[0044] The polar shift matrix is:

[0045] [EP]=R y (-x p )R x (-y P ).

[0046] Furthermore, the method for determining whether to accelerate according to the satellite position in S3 of the present invention is specifically as follows:

[0047] S31. Define monitoring area;

[0048] S32, judging the working status according to the relationship between the satellite position and the monitoring area position;

[0049] S33. Control orbit acceleration according to the satellite working status.

[0050] Furthermore, the monitoring area in step S31 of the present invention

[0051] According to actual requirements, the area where the load monitors a specific position is the hotspot area, which is represented by a hotspot in the geodetic coordinate system;

[0052] Starting from point A, a hot spot on the Earth, draw a straight line AD toward the center of the Earth;

[0053] Starting from hotspot point A, draw two lines with an angle of 30° outward along line AD. These two lines intersect with the satellite's orbit at points B and C respectively.

[0054] Arc BC defines the monitoring area where the satellite stops accelerating to ensure effective coverage of point A.

[0055] Furthermore, the method for determining the working state in step S32 of the present invention is as follows:

[0056] The orbit acceleration system determines the satellite position according to time. If the angle between the line connecting the satellite and the hotspot point A and the AD line is less than or equal to 30°, it indicates that the satellite is in or about to enter the ideal observation position for point A. At this time, the orbit acceleration system operates normally and there is no need for orbit acceleration. The satellite is in normal working condition.

[0057] If the angle between the line connecting the satellite and the hotspot point A and the AD line is greater than 30°, the surface satellite is moving away from the effective observation range of point A, and the payload central control machine applies to the simulated satellite platform to send a time modification instruction, and adjusts the orbit through orbit acceleration to ensure that it returns to or remains above the BC arc as soon as possible, that is, within the monitoring area. At this time, the satellite is in an abnormal working state. Further, the method of controlling orbit acceleration according to the satellite working state in S33 of the present invention is as follows:

[0058] Before entering the monitoring area (point B): When the satellite orbit position approaches point B, that is, the satellite is about to enter the monitoring area, the system will instruct to stop orbit acceleration after detecting this state. This is because within the BC arc, the satellite is already in a position to effectively monitor point A and no further acceleration is required.

[0059] After leaving the monitoring area (point C): Once the satellite passes point C, it means that it is about to leave the monitoring area. At this time, the payload central control machine first requests the satellite simulation platform to send a modification time instruction, and the satellite simulation platform sends a modification time instruction to the payload central control machine, and the system starts the orbit acceleration mechanism. The purpose of acceleration is to allow the satellite to enter the orbit position of the next overheating point A more quickly, ensuring continuous or quick resumption of monitoring of hot spot point A.

[0060] On the other hand, the present invention further discloses a computer-readable storage medium storing a computer program, wherein when the computer program is executed by a processor, the processor executes the steps of the above method.

[0061] On the other hand, the present invention further discloses a computer device, including a memory and a processor, wherein the memory stores a computer program, and when the computer program is executed by the processor, the processor executes the steps of the above method.

[0062] It can be seen from the above technical solution that the method and system for simulating orbital acceleration based on the six orbital numbers of the ground detection platform of the present invention can modify the simulation parameters according to the operation of the payload and accelerate to the time point when the payload is about to operate. Reduce the redundant time generated by the simulation. Since the six orbital parameters of the satellite remain unchanged when it is operating in a fixed orbit, the ground detection platform calculates the satellite orbital position according to the time point of acceleration and sends it to the payload.

[0063] This method uses six orbital numbers to represent the running curve of the entire orbit, and then determines the orbital position of the satellite at this moment based on the time. If it is in a non-operating state, the system will calculate the next operating time point and import the corresponding parameters to simulate the load test. Compared with the traditional method, it has the following advantages:

[0064] 1. Use the six track numbers and time to reduce the amount of data transmission and reduce data redundancy.

[0065] 2. A track acceleration method is implemented using the six track numbers and time to speed up the load detection time.

[0066] 3. It does not need to spend the entire track running time. In the redundant time of some unnecessary load running, it will be accelerated to reach the simulation scene of the next operation faster, saving the load detection time.

[0067] 4. The function of accelerating the observation sequence has been added, which can simulate certain working periods in multiple orbits, examine the long-term operating performance of satellites in orbit, and improve test efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0068] Figure 1 It is a flowchart of the present invention;

[0069] Figure 2 This is a flow chart of satellite orbit acceleration of the present invention;

[0070] Figure 3 This is the relationship diagram between satellite platform and payload;

[0071] Figure 4 Schematic diagram for satellite position calculation;

[0072] Figure 5 This is a schematic diagram of the satellite monitoring area of ​​the present invention;

[0073] Figure 6 This is a schematic diagram of the satellite orbit acceleration principle of the present invention;

[0074] Figure 7 Schematic diagram of the simulated satellite orbit of the present invention. DETAILED DESCRIPTION

[0075] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments.

[0076] like Figure 1 , Figure 2 As shown, the method for simulating track acceleration of a ground detection platform based on six track numbers described in this embodiment comprises the following specific steps:

[0077] S1, simulates the satellite platform sending broadcast data packets to the payload through the data bus;

[0078] A satellite platform can have multiple payloads, and the platform sends broadcast data packets to the payloads in the form of broadcasts via the data bus. The broadcast data packets contain the six orbital numbers, time and other parameters determined by the orbit;

[0079] Orbital Elements: A set of parameters used to describe the shape, size, and spatial position of an orbit. They can fully determine the motion state of an object in orbit. For artificial satellites or other celestial bodies orbiting the Earth, the orbital elements usually include the following six parameters:

[0080] 1. Semi-major axis (a): The length of the semi-major axis of the orbital ellipse determines the size of the orbit. For a circular orbit, the semi-major axis is the orbital radius.

[0081] 2. Eccentricity (e): describes the shape of the orbital ellipse, with values ​​between 0 and 1. 0 represents a circular orbit, and close to 1 represents an extremely flat ellipse.

[0082] 3. Inclination (i): The inclination angle of the orbital plane relative to a reference plane, usually relative to the Earth's equatorial plane. The range is 0° to 180°.

[0083] 4. Right Ascension of the Ascending Node (RAAN): The position angle of the ascending node (one of the intersections of the orbital plane and the equatorial plane, the point where an object crosses the equator from south to north) on the equatorial plane, usually measured from the vernal equinox.

[0084] 5. Argument of Perigee (ω): The angle between the ascending node and the perigee (the point closest to the Earth in the orbit). It defines the direction of the orbital ellipse in the orbital plane.

[0085] 6. Mean Anomaly (M): The position of the satellite relative to the perigee at a given time. Its relationship with the actual position is described by Kepler's equations.

[0086] S2, the payload central control enters the received broadcast data packet into the orbit acceleration system for analysis and calculates the satellite position;

[0087] like Figure 3 As shown, the track acceleration system includes a simulation track visualization module, a number injection module, and an acceleration judgment module;

[0088] The injection module converts the time data in the broadcast data packet input by the system into JD time, and transmits it to the acceleration judgment module together with the six track number data;

[0089] The acceleration judgment module calculates the satellite position based on JD time and six orbital numbers and determines whether the satellite needs to be accelerated, and inputs the processed data into the visualization module.

[0090] ICRF: ICRF is used in Cesium and STK software, and is also the official expression in IERS: ICRF (or ICRS) (international celestial reference frame). This coordinate system has a constant matrix difference from the J2000 inertial system we often talk about, which is also very small. For the geocentric inertial system, ICRF is equivalent to GCRS. The coordinate transformation matrix from the earth-fixed coordinate system (ITRF) to the geocentric inertial system (ICRF or GCRS) consists of polar motion, rotation, and precession and nutation.

[0091] Earth-fixed coordinate system (ITRF): This is the Earth-fixed system. It is represented by Fixed in Cesium and STK software. The official representation in IERS is: ITRF (or ITRS) (international terrestrial reference frame). This coordinate system is basically the same as the common WGS84 system, with only minor differences.

[0092] like Figure 4 As shown, the satellite position calculation method is as follows:

[0093] S21, introducing auxiliary parameters to calculate the true anomaly angle;

[0094] Two auxiliary parameters, the eccentric anomaly angle E and the mean anomaly angle M, are introduced. M is an assumed quantity, then M=n(t-t0), where t0 is the time when the satellite passes the perigee, and t is the time when the satellite is observed.

[0095] The relationship between the mean anomaly angle and the eccentric anomaly angle is E = M + e sin E. From this, the cosine of the true anomaly angle can be obtained as:

[0096]

[0097] S22, calculate the celestial coordinates of the satellite at any time by combining the true anomaly angle;

[0098] The origin of the rectangular coordinate system is taken to coincide with the center of mass of the earth, the X-axis points to the perigee, the Z-axis is perpendicular to the orbital plane and points upward, and the Y-axis is perpendicular to the X-axis on the orbital plane to form a right-hand system. The celestial coordinates of the satellite at any time are:

[0099]

[0100] Where r = a(1-ecosE);

[0101] S23, converting the celestial coordinates of the satellite at any time into geodetic coordinates;

[0102] The method to convert the satellite's celestial coordinates at any time into geodetic coordinates is as follows:

[0103]

[0104] The celestial coordinate system assumes that the earth is a sphere with a uniform texture and there is no perturbation from other celestial bodies, that is, it assumes that the direction of the earth's rotation axis in space is fixed, so the position of the vernal equinox on the celestial sphere remains unchanged. But in reality, the earth is an ellipsoid with an equatorial bulge. When it revolves, the rotation axis will be affected by the gravity of other planets, causing the vernal equinox to shift westward on the ecliptic. This phenomenon is called precession. The main causes of precession are equatorial precession and ecliptic precession, which cause the rotation axis to swing. Observing from above the North Pole, its motion trajectory approximates a circle with a circumference of 50.39″ / a in a clockwise direction centered on the North Pole. The precession of the ecliptic causes the vernal equinox to move about 0.1″ eastward on the celestial equator.

[0105] In order to calculate the coordinate transformation, the coordinate transformation matrices around the X-axis, Y-axis, and Z-axis are Rx, Ry, and Rz respectively. Take the rotation angle θ as an example:

[0106]

[0107] Then the precession matrix is:

[0108]

[0109] in

[0110]

[0111] The time T is the naturalized time calculated from the J2000.0 epoch, that is,

[0112]

[0113] where JD(TDB) represents the Julian day expressed by the barycentric dynamical time (TDB);

[0114] Since the positions of the moon, the sun and the earth are constantly changing, and the positions of the planets relative to the earth are also constantly changing, the ecliptic plane undergoes periodic changes. This causes the vernal equinox, the north celestial pole, the obliquity of the ecliptic, etc. to produce additional periodic swings on the basis of the precession of the ecliptic. This swing is called nutation.

[0115] The nutation matrix is:

[0116] [NR]=R x (-ε s -Δε)·R z (-ΔΨ)·R x (ε s )

[0117] where Δε and Ψ are the nutation of the ecliptic longitude and the nutation of the intersection angle respectively; ε s is the obliquity of the mean ecliptic, which can be expressed as

[0118] ε S =84381.448″-46.8150″T-0.00059″T 2 +0.001813″T 3

[0119] The nutation of the ecliptic longitude and the nutation of the angular radius are calculated using the model provided by the IAU (IAU1980 model).

[0120]

[0121] in,

[0122]

[0123] The nutation matrix and the data measured by IAS1980 released by IERS can be found on the IERS official website.

[0124] The position parameters of ICRF can be converted into CEP parameters according to the precession matrix and nutation matrix;

[0125] The CEP to ITRF conversion is affected by the Earth's rotation and polar axis movement. Since the Earth is in a state of rotation and is a spherical body with uneven mass, the polar axis will produce irregular rotation offsets due to gravity under the influence of various celestial bodies.

[0126] So there is a day and night rotation matrix:

[0127] [ER]=R y (θ G )

[0128] in,

[0129]

[0130] The phenomenon of the Earth's rotation pole moving on the Earth's surface is caused by the periodic swing of the Earth's instantaneous rotation axis inside the Earth. The factors of the polar shift mechanism are divided into two categories: external factors and internal factors. External factors include the gravity of the sun and the moon, as well as the effects of the atmosphere and the ocean, while internal factors involve various theoretical models, so there is a polar shift matrix:

[0131] [EP]=R y (-x p )R x (-y P ).

[0132] S3, judging whether to accelerate according to the satellite position;

[0133] The method for determining whether to accelerate based on the satellite position is as follows:

[0134] S31. Define monitoring area:

[0135] According to actual requirements, the area where the load monitors a specific position is the hotspot area, which is represented by a hotspot in the geodetic coordinate system;

[0136] like Figure 5 , Figure 6 As shown in the figure, the schematic diagram of satellite position acceleration determination is as follows: starting from the hot spot point A on the earth, a straight line AD is drawn toward the center of the earth.

[0137] Starting from hotspot A, draw two lines with an angle of 30° outward along line AD. These two lines intersect with the satellite's orbit at points B and C respectively.

[0138] Arc BC defines the monitoring area where the satellite stops accelerating to ensure effective coverage of point A.

[0139] S32. Determine the working status based on the relationship between the satellite position and the monitoring area position:

[0140] The method for determining the working status is as follows:

[0141] The orbit acceleration system determines the satellite position according to time. If the angle between the line connecting the satellite and the hotspot point A and the AD line is less than or equal to 30°, it indicates that the satellite is in or about to enter the ideal observation position for point A. At this time, the orbit acceleration system operates normally and there is no need for orbit acceleration. The satellite is in normal working condition.

[0142] If the angle between the line connecting the satellite and hotspot point A and the AD line is greater than 30°, the surface satellite is moving away from the effective observation range of point A. The payload central control computer applies to the simulated satellite platform to send a time modification instruction, and adjusts the orbit through orbit acceleration to ensure that it returns to or remains above the BC arc as soon as possible, that is, within the monitoring area. At this time, the satellite is in an abnormal working state.

[0143] S33, controlling orbit acceleration according to the satellite working status;

[0144] Before entering the monitoring area (point B): When the satellite orbit position approaches point B, that is, the satellite is about to enter the monitoring area, the system will instruct to stop orbit acceleration after detecting this state. This is because within the BC arc, the satellite is already in a position to effectively monitor point A and no further acceleration is required.

[0145] After leaving the monitoring area (point C): Once the satellite passes point C, it means that it is about to leave the monitoring area. At this time, the payload central control machine first requests the satellite simulation platform to send a modification time instruction, and the satellite simulation platform sends a modification time instruction to the payload central control machine, and the system starts the orbit acceleration mechanism. The purpose of acceleration is to allow the satellite to enter the orbit position of the next overheating point A more quickly, ensuring continuous or quick resumption of monitoring of hot spot point A.

[0146] S4, controlling whether orbit acceleration ends according to whether the satellite reaches the hotspot position;

[0147] When the angle between the satellite and hotspot A is greater than 30°, it means that the satellite has not reached the hotspot position. At this time, orbital acceleration is in progress. The angle will increase until it reaches the point with an angle of 30° again after moving to the orbital position of the next overheating point A. This proves that the hotspot has been reached and the acceleration stops at this moment.

[0148] When the angle between the satellite and hotspot A is less than 30°, it means that the satellite has reached the hotspot and the acceleration ends.

[0149] S5, when track acceleration ends, determine whether hotspot detection is completed;

[0150] The visualization module in the orbit acceleration system completes the satellite orbit drawing, that is, the hotspot detection is completed. The orbit drawing is as follows Figure 7 shown.

[0151] In summary, the method and device for simulating orbital acceleration based on the six orbital numbers of the ground detection platform of the present invention modify the simulation parameters according to the operation of the payload and accelerate to the time point when the payload is about to operate. Reduce the redundant time generated by the simulation. Since the six orbital parameters of the satellite remain unchanged when it is operating in a fixed orbit, the ground detection platform calculates the satellite orbital position according to the time point of acceleration and sends it to the payload.

[0152] This method uses six orbital numbers to represent the running curve of the entire orbit, and then determines the orbital position of the satellite at this moment based on the time. If it is in a non-operating state, the system will calculate the next operating time point and import the corresponding parameters to simulate the load test. Compared with the traditional method, it has the following advantages:

[0153] 1. Use the six track numbers and time to reduce the amount of data transmission and reduce data redundancy.

[0154] 2. A track acceleration method is implemented using the six track numbers and time to speed up the load detection time.

[0155] 3. It does not need to spend the entire track running time. In the redundant time of some unnecessary load running, it will be accelerated to reach the simulation scene of the next operation faster, saving the load detection time.

[0156] 4. The function of accelerating the observation sequence has been added, which can simulate certain working periods in multiple orbits, examine the long-term operating performance of satellites in orbit, and improve test efficiency.

[0157] On the other hand, the present invention further discloses a computer-readable storage medium storing a computer program, wherein when the computer program is executed by a processor, the processor executes the steps of the above method.

[0158] On the other hand, the present invention further discloses a computer device, including a memory and a processor, wherein the memory stores a computer program, and when the computer program is executed by the processor, the processor executes the steps of the above method.

[0159] In another embodiment provided in the present application, a computer program product comprising instructions is also provided, which, when executed on a computer, enables the computer to execute any of the methods for simulating orbital acceleration of a ground detection platform based on six orbital numbers in the above-mentioned embodiments.

[0160] It is understandable that the system, device and storage medium provided in the embodiments of the present invention correspond to the method provided in the embodiments of the present invention, and the explanation, examples and beneficial effects of the relevant contents can refer to the corresponding parts in the above methods.

[0161] In the above embodiments, it can be implemented in whole or in part by software, hardware, firmware or any combination thereof. When implemented using software, it can be implemented in whole or in part in the form of a computer program product. The computer program product includes one or more computer instructions. When the computer program instructions are loaded and executed on a computer, the process or function described in the embodiment of the present application is generated in whole or in part. The computer may be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions may be stored in a computer-readable storage medium, or transmitted from one computer-readable storage medium to another computer-readable storage medium, for example, the computer instructions may be transmitted from a website site, computer, server or data center by wired (e.g., coaxial cable, optical fiber, digital subscriber line (DSL)) or wireless (e.g., infrared, wireless, microwave, etc.) mode to another website site, computer, server or data center. The computer-readable storage medium may be any available medium that a computer can access or a data storage device such as a server or data center that includes one or more available media integrated. The available medium may be a magnetic medium, (e.g., a floppy disk, a hard disk, a tape), an optical medium (e.g., a DVD), or a semiconductor medium (e.g., a solid-state drive Solid State Disk (SSD)), etc.

[0162] It should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the existence of other identical elements in the process, method, article or device including the elements.

[0163] Each embodiment in this specification is described in a related manner, and the same or similar parts between the embodiments can be referred to each other, and each embodiment focuses on the differences from other embodiments. In particular, for the system embodiment, since it is basically similar to the method embodiment, the description is relatively simple, and the relevant parts can be referred to the partial description of the method embodiment.

[0164] The above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit the same. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that the technical solutions described in the aforementioned embodiments may still be modified, or some of the technical features may be replaced by equivalents. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for simulating track acceleration on a ground detection platform based on six track numbers, characterized in that: The specific steps are as follows: S1, simulates the satellite platform sending broadcast data packets to the payload through the data bus; S2, the payload central control enters the received broadcast data packet into the orbit acceleration system for analysis and calculates the satellite position; S3, judging whether to accelerate according to the satellite position; S4, controlling whether orbit acceleration ends according to whether the satellite reaches the hotspot position; S5. When track acceleration is finished, determine whether hotspot detection is completed.

2. The method for simulating track acceleration using a ground detection platform based on six track numbers according to claim 1, characterized in that: The track acceleration system in S2 includes a simulated track visualization module, a number injection module, and an acceleration judgment module; The injection module converts the time data in the broadcast data packet input by the system into JD time, and transmits it to the acceleration judgment module together with the six track number data; The acceleration judgment module calculates the satellite position based on JD time and six orbital numbers and determines whether the satellite needs to be accelerated, and inputs the processed data into the visualization module.

3. The method for simulating track acceleration using a ground detection platform based on six track numbers according to claim 2, characterized in that: The satellite position calculation method in S2 is as follows: S21, introducing auxiliary parameters to calculate the true anomaly angle; Two auxiliary parameters, the eccentric anomaly angle E and the mean anomaly angle M, are introduced. M is an assumed quantity, then M = n(t-t0), where t0 is the time when the satellite passes the perigee, and t is the time when the satellite is observed. The relationship between the mean anomaly angle and the eccentric anomaly angle is E = M + e sin E, so the cosine of the true anomaly angle is: S22, calculate the celestial coordinates of the satellite at any time by combining the true anomaly angle; The origin of the rectangular coordinate system is taken to coincide with the center of mass of the earth, the X-axis points to the perigee, the Z-axis is perpendicular to the orbital plane and points upward, and the Y-axis is perpendicular to the X-axis on the orbital plane to form a right-hand system. The celestial coordinates of the satellite at any time are: Where r = a(1-ecosE); S23, converting the celestial coordinates of the satellite at any time into geodetic coordinates; The method to convert the satellite's celestial coordinates at any time into geodetic coordinates is as follows: The precession matrix is: in, θ is the rotation angle, Ry is the coordinate transformation matrix around the Y axis, and Rz is the coordinate transformation matrix around the Z axis; Among them, time T is the naturalized time calculated from the J2000.0 epoch, that is, where JD(TDB) represents the Julian day expressed by the barycentric dynamical time (TDB); The nutation matrix is: [NR]=R x (-e s -D)·R z (-ΔΨ)·R x (e s ) where Δε and Ψ are the nutation of the ecliptic longitude and the nutation of the intersection angle respectively; ε s is the obliquity of the mean ecliptic, Rx is the coordinate transformation matrix around the X axis; Day and night rotation matrix: [IS]=R y (θ G ) in, The polar shift matrix is: [EP]=R y (-x p )R x (-y P )。 4. The method for simulating track acceleration using a ground detection platform based on six track numbers according to claim 1, characterized in that: The method for determining whether to accelerate according to the satellite position in S3 is as follows: S31. Define monitoring area; S32, judging the working status according to the relationship between the satellite position and the monitoring area position; S33. Control orbit acceleration according to the satellite working status.

5. The method for simulating track acceleration using a ground detection platform based on six track numbers according to claim 4, characterized in that: In step S31, the monitoring area According to actual requirements, the area where the load monitors a specific position is the hotspot area, which is represented by a hotspot in the geodetic coordinate system; Starting from point A, a hot spot on the Earth, draw a straight line AD toward the center of the Earth; Starting from hotspot point A, draw two lines with an angle of 30° outward along line AD. These two lines intersect with the satellite's orbit at points B and C respectively. Arc BC defines the monitoring area where the satellite stops accelerating to ensure effective coverage of point A.

6. The method for simulating track acceleration using a ground detection platform based on six track numbers according to claim 4, characterized in that: The method for determining the working state in step S32 is as follows: The orbit acceleration system determines the satellite position according to time. If the angle between the line connecting the satellite and the hotspot point A and the AD line is less than or equal to 30°, it indicates that the satellite is in or about to enter the ideal observation position for point A. At this time, the orbit acceleration system operates normally and there is no need for orbit acceleration. The satellite is in normal working condition. If the angle between the line connecting the satellite and hotspot point A and the AD line is greater than 30°, the surface satellite is moving away from the effective observation range of point A. The payload central control computer applies to the simulated satellite platform to send a time modification instruction, and adjusts the orbit through orbit acceleration to ensure that it returns to or remains above the BC arc as soon as possible, that is, within the monitoring area. At this time, the satellite is in an abnormal working state.

7. The method for simulating track acceleration using a ground detection platform based on six track numbers according to claim 4, characterized in that: The method of controlling orbit acceleration according to the satellite working state in S33 is as follows: Before entering the monitoring area (point B): When the satellite orbit position approaches point B, that is, the satellite is about to enter the monitoring area, the system will instruct to stop orbit acceleration after detecting this state. This is because within the BC arc, the satellite is already in a position to effectively monitor point A and no further acceleration is required. After leaving the monitoring area (point C): Once the satellite passes point C, it means that it is about to leave the monitoring area. At this time, the payload central control machine first requests the satellite simulation platform to send a modification time instruction, and the satellite simulation platform sends a modification time instruction to the payload central control machine, and the system starts the orbit acceleration mechanism. The purpose of acceleration is to allow the satellite to enter the orbit position of the next overheating point A more quickly, ensuring continuous or quick resumption of monitoring of hot spot point A.

8. A computer device comprising a memory and a processor, wherein the memory stores a computer program, and when the computer program is executed by the processor, the processor executes the steps of the method according to any one of claims 1 to 7.

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