Rail orbit six-number based ground detection platform simulation rail acceleration method and equipment

Abstract

The application discloses a simulation track acceleration method and equipment based on a track six-number ground detection platform, and the method is characterized in that, in the process of load ground testing, a satellite platform sends broadcast data packets to the load through a data bus; a load control interface analyzes the received broadcast data packets and calculates the satellite position; whether to accelerate is judged according to the satellite position; whether to end the track acceleration is controlled according to whether the satellite reaches a hot spot position; and whether the hot spot detection is completed is judged when the track acceleration is ended. The method uses track six-number and time to determine the satellite running track according to the running curve of the whole track, divides the load operation scene according to the load operation time, calculates the next operation time point if the load is in a non-operation state, and imports corresponding parameters to simulate the test of the load, accelerates the track simulation process, and accelerates the track simulation process in the redundant time of some unnecessary load operation, so that the simulation scene of the 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 on a ground detection platform based on six orbital elements. Background Art

[0002] The high-resolution spaceborne hotspot monitor monitors greenhouse gas emissions in global hotspots from orbit using a nadir observation method. While in orbit, the hotspot monitor receives broadcast time packets sent by the satellite platform every second, which contain the satellite's orbital information. The monitor uses this orbital information to adjust the pointing mirror angle, aiming at the hotspot, and then taking images. The payload does not operate continuously during its orbital operation; it enters standby mode in dark areas below the satellite or in areas with fewer hotspots.

[0003] During payload development, a ground-based test platform is required to simulate the satellite platform and perform ground testing on the payload to ensure its proper operation while in orbit. Ground-based commissioning of the payload, if it follows the actual orbit, is time-consuming and inefficient.

[0004] The equipment used for ground-based payload commissioning is called a ground test platform. Satellites consist of two main components: the satellite platform and the payload, which communicate using a unified protocol. During payload development, the ground test platform simulates the satellite platform and conducts ground tests on the payload to verify its functionality, eliminate potential safety hazards, and ensure its normal operation in orbit.

[0005] The ground detection platform needs to measure and control the payload and receive service data. The measurement and control platform uses the measurement and control bus 1553B. 1553B communication includes internal commands, data injection, telemetry information, and broadcast time packets. Measurement and control is generally referred to as "digital communication," primarily encompassing direct commands, indirect commands, direct telemetry, and indirect telemetry. Service data is generally referred to as "digital transmission communication," primarily encompassing data transmission and the reception of broadcast signals.

[0006] Because the orbit is divided into different scenarios based on payload operation time, the satellite will not operate in some scenarios, such as when the satellite enters darkness or is not in a hotspot area. During the testing process, the payload's redundant time beyond the necessary operation time should be minimized. Therefore, orbit acceleration was designed as a method to reduce this redundant time.

[0007] The existing method mainly uses orbit simulation software such as STK to create orbital data for simulated satellite operation in advance, and then has the ground inspection platform analyze and transmit it. However, due to the large size of STK software, the software interface is difficult to call, and the calculation time is long, the output data cannot be used directly and requires secondary processing.

[0008] This invention discloses a method for simulating orbital acceleration on a ground-based testing platform based on six orbital numbers. During ground payload testing, the six orbital numbers and time are used to determine the satellite's orbit. The payload's operational scenarios are divided according to the payload's operational time. When the payload is in a non-operating period, the orbital simulation process is accelerated. This method quickly reaches the next operational moment of the payload, thereby reducing redundant time during payload simulation. (Hereinafter referred to as "orbital acceleration") This algorithm is customized according to the operating status of the satellite and payload and can be embedded in other programs. It 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 using a ground detection platform based on six track numbers, comprising the following steps:

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

[0013] S2: The payload central control system analyzes the received broadcast data packets and calculates the satellite position.

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

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

[0016] S5. When the track acceleration ends, determine whether the 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 recorded 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, introduce auxiliary parameters to calculate the true anomaly;

[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 and the eccentric anomaly is E = M + e sin E, so the cosine of the true anomaly is:

[0024]

[0025] S22, calculate the satellite's arbitrary time-controlled celestial coordinates based on the true anomaly;

[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-handed 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 in 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's operating 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 location is called the hotspot area and is represented by a hotspot in the geodetic coordinate system;

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

[0053] Starting from hotspot A, draw two lines with an angle of 30° 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 status in step S32 of the present invention is as follows:

[0056] The orbital acceleration system determines the satellite position based on time. If the angle between the line connecting the satellite and hotspot 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 orbital acceleration system is operating normally and there is no need for orbital acceleration. The satellite is in normal working condition.

[0057] If the angle between the line connecting the satellite and hotspot A and line AD is greater than 30°, indicating that the surface satellite is moving away from the effective observation range of point A, the payload central control computer requests a time modification instruction from the simulated satellite platform, adjusting the orbit through orbit acceleration to ensure that it returns to or remains above the BC arc, i.e., within the monitoring area, as soon as possible. At this point, the satellite is in an abnormal operating state. Furthermore, the method for controlling orbit acceleration based on the satellite operating state, described in S33 of the present invention, is as follows:

[0058] Before entering the monitoring area (Point B): When the satellite's orbital position approaches Point B, meaning the satellite is about to enter the monitoring area, the system detects this and instructs the satellite to stop orbital acceleration. 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, indicating it is about to leave the monitoring area, the payload control unit first requests a time modification command from the satellite simulation platform. The satellite simulation platform sends the time modification command to the payload control unit, and the system initiates the orbit acceleration mechanism. The purpose of the acceleration is to allow the satellite to more quickly enter the orbital position of the next hotspot point A, ensuring continuous or expedited resumption of monitoring of hotspot point A.

[0060] In another aspect, 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, 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 above method.

[0062] As can be seen from the above technical solution, the present invention's method and system for simulating orbital acceleration using the six orbital parameters of a ground-based inspection platform modifies simulation parameters based on the payload's operational status, accelerating the payload to the point where it is about to operate. This reduces redundant simulation time. Since the six orbital parameters remain unchanged when a satellite is operating in a fixed orbit, the ground-based inspection platform calculates the satellite's orbital position based on the acceleration time and transmits this information to the payload.

[0063] This method uses six orbital numbers to represent the entire orbital curve, and then uses time to determine the satellite's orbital position at that moment. If the satellite is in an inoperative state, the system calculates the next operational time point and imports the corresponding parameters to simulate the payload. Compared to traditional methods, it has the following advantages:

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

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

[0066] 3. Instead of spending the entire track running time, the redundant time for some payloads that are not necessary to run will be accelerated to reach the next simulation scenario more quickly, saving payload detection time.

[0067] 4. The function of accelerating 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 using 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 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 fully determine the motion of an object in orbit. For satellites or other celestial bodies orbiting the Earth, the orbital elements typically include the following six parameters:

[0080] 1. Semi-major axis (a): The length of the semi-major axis of an 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 a value between 0 and 1. 0 indicates a circular orbit, and close to 1 indicates an extremely flat ellipse.

[0082] 3. Inclination (i): The tilt of the orbital plane relative to a reference plane, usually relative to the Earth's equatorial plane. It ranges from 0° to 180°.

[0083] 4. Right Ascension of the Ascending Node (RAAN): The angular position 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) in 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 orbit). It defines the orientation of the orbital ellipse within the orbital plane.

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

[0086] S2: The payload central control system analyzes the received broadcast data packets 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 recorded 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] Earth-centered inertial reference frame (ICRF): ICRF is used in both Cesium and STK software, and is also formally designated by the International Celestial Reference Frame (IERS): ICRF (or ICRS) (International Celestial Reference Frame). This coordinate system differs slightly from the commonly used J2000 inertial reference frame in terms of constant matrix. For the Earth-centered inertial reference frame, ICRF is equivalent to GCRS. The coordinate transformation matrix from the Earth-fixed reference frame (ITRF) to the Earth-centered inertial reference frame (ICRF or GCRS) consists of polar motion, rotation, and precession and nutation.

[0091] Earth-Fixed Reference Frame (ITRF): This coordinate system is represented by the term "Fixed" in Cesium and STK. Its formal designation in the International Terrestrial Reference Frame (IERS) is ITRF (or ITRS). This coordinate system is essentially the same as the commonly used WGS84 system, with only minor differences.

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

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

[0094] Two auxiliary parameters, the eccentric anomaly E and the mean anomaly 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 and the eccentric anomaly is E = M + e sin E. From this, the cosine of the true anomaly is:

[0096]

[0097] S22, calculate the satellite's arbitrary time-controlled celestial coordinates based on the true anomaly;

[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-handed 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 uniform sphere without the perturbations of other celestial bodies, that is, it assumes that the direction of the Earth's rotation axis in space is fixed, and therefore the position of the vernal equinox on the celestial sphere remains unchanged. However, the Earth is actually an ellipsoid with an equatorial bulge. As it orbits, the rotation axis is affected by the gravitational pull 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 wobble. Observed from above the North Pole, its 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 approximately 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 in barycentric dynamical time (TDB);

[0114] The constantly changing positions of the Moon, Sun, and Earth, as well as the positions of the planets relative to Earth, cause periodic shifts in the ecliptic plane. This causes additional periodic oscillations in the vernal equinox, the North Celestial Pole, and the obliquity of the ecliptic, in addition to the precession of the ecliptic. This oscillation 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 ecliptic nutation and the angular nutation 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 transition is affected by the Earth's rotation and polar axis shift. Because the Earth is rotating and its mass is uneven, the polar axis will experience irregular rotational shifts due to the gravitational pull of various celestial bodies.

[0126] So we have the day and night rotation matrix:

[0127] [ER]=R y (θ G )

[0128] in,

[0129]

[0130] The movement of the Earth's rotation pole on the Earth's surface is caused by the periodic oscillation of the Earth's instantaneous rotation axis within the Earth. The factors that influence the polar shift mechanism are divided into two categories: external factors and internal factors. External factors include the gravitational forces of the sun and moon, as well as the effects of the atmosphere and oceans, while internal factors involve various theoretical models, resulting in the polar shift matrix:

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

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

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

[0134] S31. Define monitoring area:

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

[0136] like Figure 5 、 Figure 6 As shown in FIG, a schematic diagram of satellite position acceleration determination is shown, starting from the hotspot point A on the earth and drawing a straight line AD toward the center of the earth.

[0137] Starting from hotspot A, draw two lines with an angle of 30° 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 orbital acceleration system determines the satellite position based on time. If the angle between the line connecting the satellite and hotspot 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 orbital acceleration system is operating normally and there is no need for orbital 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 requests 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 operating status;

[0144] Before entering the monitoring area (Point B): When the satellite's orbital position approaches Point B, meaning the satellite is about to enter the monitoring area, the system detects this and instructs the system to stop orbital acceleration. 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, indicating it is about to leave the monitoring area, the payload control unit first requests a time modification command from the satellite simulation platform. The satellite simulation platform sends the time modification command to the payload control unit, and the system initiates the orbit acceleration mechanism. The purpose of the acceleration is to allow the satellite to more quickly enter the orbital position of the next hotspot point A, ensuring continuous or expedited resumption of monitoring of hotspot point A.

[0146] S4, controlling whether the orbit acceleration ends based on 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, the orbital acceleration is in progress. The angle will first increase until it reaches the orbital position of the next overheating point A and reaches the point with an angle of 30° again. 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 position 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 present invention's method and equipment for simulating orbital acceleration using the six orbital parameters of a ground-based inspection platform modifies simulation parameters based on the payload's operational status, accelerating the payload to the point where it is about to operate. This reduces redundant simulation time. Since the six orbital parameters remain unchanged when a satellite is operating in a fixed orbit, the ground-based inspection platform calculates the satellite's orbital position based on the acceleration time and transmits this information to the payload.

[0152] This method uses six orbital numbers to represent the entire orbital curve, and then uses time to determine the satellite's orbital position at that moment. If the satellite is in an inoperative state, the system calculates the next operational time point and imports the corresponding parameters to simulate the payload. Compared to traditional methods, it has the following advantages:

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

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

[0155] 3. Instead of spending the entire track running time, the redundant time for some payloads that are not necessary to run will be accelerated to reach the next simulation scenario more quickly, saving payload detection time.

[0156] 4. The function of accelerating 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] In another aspect, 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, 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 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 track acceleration based on the six track numbers on a ground detection platform 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 of 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 software is used for implementation, 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 can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions can 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 can be transmitted from a website, computer, server or data center to another website, computer, server or data center via a wired (e.g., coaxial cable, optical fiber, digital subscriber line (DSL)) or wireless (e.g., infrared, wireless, microwave, etc.) method. The computer-readable storage medium can 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 integrations. The available medium can 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 (SSD)).

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

[0163] Each embodiment in this specification is described in a related manner. Similar parts between the various embodiments can be referred to in conjunction with each other. Each embodiment focuses on the differences between the other embodiments. In particular, the system embodiment is generally similar to the method embodiment, so the description is relatively simple. For related parts, refer to the 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 they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. 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 various 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 system analyzes the received broadcast data packets and enters them into the orbit acceleration system to calculate the satellite position. S3, judging whether to accelerate based on the satellite position; S4, controlling whether the orbit acceleration ends based on whether the satellite reaches the hotspot position; S5. When track acceleration ends, determine whether hotspot detection is completed; The broadcast data packet contains six orbital parameters, including semi-major axis, eccentricity, orbit inclination, right ascension of ascending node, argument of perigee, and mean anomaly. The hotspot position is the area where the load monitors a specific position according to actual requirements. This area is the hotspot area and is represented by a hotspot in the geodetic coordinate system.

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 recorded 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, introduce auxiliary parameters to calculate the true anomaly; Introduce two auxiliary parameters, the eccentric anomaly angle E and the mean anomaly angle M, where M is an assumed quantity. ,in, is the time when the satellite passes the perigee, and t is the time when the satellite is observed; There is a relationship between the mean anomaly and the eccentric anomaly as follows: , then the cosine of the true anomaly is: S22, calculate the satellite's arbitrary time-controlled celestial coordinates based on the true anomaly; 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-handed system. The celestial coordinates of the satellite at any time are: in, ; 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 represented by the barycentric dynamical time TDB; The nutation matrix is: in and They are ecliptic nutation and angular nutation respectively; is the obliquity of the mean ecliptic, Rx is the coordinate transformation matrix around the X axis; Day and night rotation matrix: in, The polar shift matrix is: 。 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 based on 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's operating 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: The method for defining the monitoring area in step S31 includes: Starting from point A, a hot spot on Earth, draw a straight line AD toward the center of the Earth; Starting from hotspot A, draw two lines with an angle of 30° 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 status in step S32 is as follows: The orbital acceleration system determines the satellite position based on time. If the angle between the line connecting the satellite and hotspot 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 orbital acceleration system is operating normally and there is no need for orbital 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 requests 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: Point B before entering the monitoring area: When the satellite's orbital position approaches point B, that is, the satellite is about to enter the monitoring area, the system will detect this state and instruct to stop orbital acceleration; This is because within the BC arc, the satellite is already in a position where it can 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 computer first requests the satellite simulation platform to send a time modification instruction. The satellite simulation platform sends the time modification instruction to the payload central control computer, and the system starts the orbit acceleration mechanism. The purpose of the acceleration is to enable the satellite to enter the orbit position of the next overheating point A more quickly, ensuring that the monitoring of the hotspot point A is continued or resumed as soon as possible.

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 is caused to perform the steps of the method according to any one of claims 1 to 7.

Images