Method and device for acquiring sun-synchronous satellite orbit point root number

Abstract

The present disclosure provides a sun synchronous satellite orbit point root acquisition method and device, which relates to the technical field of spacecraft and can solve the problem that the satellite orbit point root cannot be acquired based on the launch site position and the descending node local time in the related art. The specific technical solution is: calculating initial orbit parameters according to the sun synchronous satellite epoch time, the preset orbit height and the preset descending node local time; calculating the time of the satellite orbit point; adjusting the ascending node right ascension according to the position of the launch site; adjusting the perigee amplitude angle; acquiring the calculated value of the descending node local time according to the time of the satellite orbit point, the adjusted ascending node right ascension, the adjusted perigee amplitude angle and other parameters in the initial orbit parameters; when the calculated value of the descending node local time matches the preset descending node local time, determining that the orbit parameters corresponding to the calculated value of the descending node local time are the sun synchronous satellite orbit point root; when they do not match, re-executing the above adjustment steps until they match.

Description

Technical Field

[0001] This disclosure relates to the field of spacecraft, and in particular to a method and apparatus for obtaining the orbital elements of a sun-synchronous satellite. Background Technology

[0002] With the booming development of commercial spaceflight, spacecraft launch services are also increasing. Remote control commands need to be injected in the early stages of telemetry and control, making the prediction, calculation, and analysis of the initial visible telemetry and control arc segment at tracking stations crucial.

[0003] Currently, the initial visible arc segment list for satellite tracking and control can be calculated using the satellite's orbital insertion point (OIP) count and the geodetic coordinates of the tracking and control stations. Users can then arrange data injection and remote control commands based on the locations of different tracking and control stations, thereby achieving precise control of the satellite and payload activation. Users can also select appropriate tracking stations based on the satellite's OIP count, achieving precise control while avoiding duplicate applications and waste of tracking and control resources, thus reducing tracking and control costs.

[0004] However, it is currently impossible to obtain the satellite's orbital insertion points based on the launch site location and the local time of the descending node. Summary of the Invention

[0005] This disclosure provides a method and apparatus for obtaining the orbital elements of a sun-synchronous satellite, which solves the current problem that it is impossible to obtain the satellite's orbital elements based on the launch site location and the local time of the descending node. The technical solution is as follows:

[0006] According to a first aspect of the present disclosure, a method for obtaining the orbital insertion point elements of a sun-synchronous satellite is provided, the method comprising:

[0007] The initial orbital parameters are calculated based on the epoch time of the sun-synchronous satellite, the preset orbital altitude, and the local time of the preset descending node.

[0008] Calculate the time of satellite insertion point;

[0009] Based on the location of the launch site, adjust the right ascension of the ascending node in the initial orbital parameters so that the satellite's nadir trajectory passes near the geodetic coordinates of the launch site;

[0010] Adjust the perigee argument in the initial orbital parameters;

[0011] The calculated value of the local time of the descending node is obtained based on the time of the satellite's orbit insertion point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and other parameters in the initial orbital parameters;

[0012] When the calculated value of the local time at the descending node matches the preset local time at the descending node, the orbital parameter corresponding to the calculated value of the local time at the descending node is determined to be the element of the sun-synchronous satellite's orbital insertion point.

[0013] If the calculated value of the descending node local time does not match the preset descending node local time, the steps of adjusting the right ascension of the ascending node and adjusting the argument of the perigee are repeated until the orbital parameters corresponding to the determined calculated value of the descending node local time match the preset descending node local time.

[0014] The method disclosed herein can extrapolate the orbital elements of a sun-synchronous satellite based on information such as the satellite's orbital altitude, the local time of the descending node, and the location of the launch site. It has high accuracy and provides essential orbital data for early-stage satellite tracking and forecasting calculations, thus providing strong support for the scheduling and decision-making of commercial satellite tracking and control plans.

[0015] In one embodiment, the calculation of the satellite's orbital insertion point time includes:

[0016] The time of the satellite's orbital insertion point is obtained based on the launch time at the launch site and the duration of the active phase flight.

[0017] In one embodiment, obtaining the satellite's orbital insertion point time based on the launch time from the launch site and the duration of the active phase flight includes:

[0018] The sum of the launch time at the launch site and the active phase flight duration is determined as the time of the satellite's orbital insertion point.

[0019] In one embodiment, the step of obtaining the calculated value of the local time of the descending node based on the time of the satellite's orbit insertion point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and the initial orbital parameters includes:

[0020] The satellite's nadir parameters are calculated based on the satellite's orbital insertion time, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and the initial orbital parameters.

[0021] The calculated value of the local time of the descending node is obtained by dividing the longitude of the nadir point in the nadir point parameters corresponding to the latitude of the nadir point being 0 by the Earth's rotation angular velocity.

[0022] According to a second aspect of the present disclosure, a device for obtaining the orbital insertion point elements of a sun-synchronous satellite is provided, the device comprising:

[0023] The first calculation module is used to calculate the initial orbit parameters based on the epoch time of the sun-synchronous satellite, the preset orbital altitude, and the local time of the preset descending node.

[0024] The second calculation module is used to calculate the time of the satellite's orbital insertion point;

[0025] The first adjustment module is used to adjust the right ascension of the ascending node in the initial orbital parameters according to the location of the launch site, so that the satellite's nadir trajectory passes near the geodetic coordinates of the launch site;

[0026] The second adjustment module is used to adjust the perigee argument in the initial orbital parameters;

[0027] The acquisition module is used to obtain the calculated value of the local time of the descending node based on the time of the satellite's entry point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and other parameters in the initial orbital parameters;

[0028] The determination module is used to determine the orbital parameters corresponding to the calculated value of the descending node local time as the sun-synchronous satellite's orbital element number when the calculated value of the descending node local time matches the preset descending node local time.

[0029] The loop module is used to re-execute the steps of adjusting the right ascension of the ascending node and adjusting the argument of the perigee when the calculated value of the local time of the descending node does not match the preset local time of the descending node, until the orbital parameters corresponding to the calculated value of the local time of the descending node match the preset local time of the descending node.

[0030] In one embodiment, the second computing module includes:

[0031] The first acquisition submodule is used to acquire the time of the satellite's orbital insertion point based on the launch time at the launch site and the duration of the active phase flight.

[0032] In one embodiment, the first acquisition submodule includes:

[0033] The acquisition sub-unit is used to determine the time of the satellite's orbit insertion point as the sum of the launch time at the launch site and the active phase flight duration.

[0034] In one embodiment, the acquisition module includes:

[0035] The calculation submodule is used to calculate the satellite's nadir parameters based on the satellite's orbital insertion time, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and the initial orbital parameters.

[0036] The second acquisition submodule is used to divide the longitude of the nadir point in the nadir point parameters corresponding to the latitude of the nadir point being 0 by the Earth's rotation angular velocity to obtain the calculated value of the local time of the descending node.

[0037] According to a third aspect of the present disclosure, a device for obtaining the orbital elements of a sun-synchronous satellite is provided, comprising:

[0038] processor;

[0039] Memory used to store processor-executable instructions;

[0040] The processor is configured to execute the method described in any of the above embodiments.

[0041] According to a fourth aspect of the present disclosure, a computer-readable storage medium is provided that stores computer instructions thereon, which, when executed by a processor, implement the method described in any of the preceding embodiments.

[0042] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit this disclosure. Attached Figure Description

[0043] The accompanying drawings, which are incorporated in and form a part of this specification, illustrate embodiments consistent with this disclosure and, together with the description, serve to explain the principles of this disclosure.

[0044] Figure 1 This is a flowchart of a method for obtaining the orbital elements of a sun-synchronous satellite according to an embodiment of this disclosure;

[0045] Figure 2 This is a flowchart of a method for obtaining the orbital elements of a sun-synchronous satellite according to an embodiment of this disclosure;

[0046] Figure 3 This is a structural diagram of a device for obtaining the orbital elements of a sun-synchronous satellite according to an embodiment of this disclosure;

[0047] Figure 4 This is a structural diagram of the second calculation module in a device for obtaining the orbital elements of a sun-synchronous satellite provided in this embodiment of the present disclosure;

[0048] Figure 5 This is a structural diagram of the first acquisition submodule in a device for acquiring the orbital elements of a sun-synchronous satellite provided in an embodiment of this disclosure;

[0049] Figure 6 This is a structural diagram of the acquisition module in a device for acquiring the orbital elements of a sun-synchronous satellite provided in an embodiment of this disclosure. Detailed Implementation

[0050] Exemplary embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. When the following description relates to the drawings, unless otherwise indicated, the same numerals in different drawings denote the same or similar elements. The embodiments described in the following exemplary embodiments do not represent all embodiments consistent with this disclosure. Rather, they are merely examples of apparatuses and methods consistent with some aspects of this disclosure as detailed in the appended claims.

[0051] This disclosure employs a method of adjusting the right ascension of the ascending node and the argument of perigee of the instantaneous true equatorial elements to continuously iterate the satellite's orbit, thereby obtaining the orbital elements of the sun-synchronous satellite. The following embodiments detail the technical solution of this disclosure.

[0052] This disclosure provides a method for obtaining the orbital elements of a sun-synchronous satellite, such as... Figure 1 As shown, the method includes:

[0053] 101. Calculate the initial orbital parameters based on the epoch time of the sun-synchronous satellite, the preset orbital altitude, and the local time of the preset descending node.

[0054] Specifically, this step is used to calculate the initial trajectory.

[0055] The following parameters can be given: sun-synchronous satellite epoch time, preset orbital altitude, and preset descending node local time.

[0056] Given the parameters of the sun-synchronous satellite epoch time, preset orbital altitude, and preset descending node local time, the initial orbit is calculated based on the publicly available method for calculating sun-synchronous orbital elements. At this point, the Kepler elements in the instantaneous true equatorial coordinate system at that epoch can be obtained.

[0057] The method for calculating the initial orbit based on the publicly available method for calculating the sun-synchronous orbit elements is the same as that in related technologies, and will not be repeated here.

[0058] 102. Calculate the time of satellite insertion into orbit;

[0059] In one embodiment, the satellite's orbital insertion time can be obtained based on the launch time at the launch site and the duration of the active phase flight.

[0060] For example, the launch time at the launch site and the duration of the active phase flight can be used to determine the time of the satellite's orbital insertion point.

[0061] Specifically, this step is used to calculate the epoch time.

[0062] Because the eccentricity of a sun-synchronous orbit is small and it is a near-circular orbit, the time of satellite insertion can be obtained by adding the launch time from the launch site to the active phase flight time.

[0063] 103. Adjust the right ascension of the ascending node in the initial orbital parameters according to the location of the launch site, so that the satellite's nadir trajectory passes near the geodetic coordinates of the launch site;

[0064] Depending on the location of the launch site, the right ascension of the ascending node can be adjusted to adjust the geographical longitude of the descending node, allowing the satellite's nadir trajectory to pass near the geodetic coordinates of the launch site.

[0065] The geodetic coordinates of the launch site can be understood as the location at a preset distance from the geodetic coordinates of the launch site. The preset distance can be defined and modified according to the actual situation, or it can be predefined. This disclosure does not limit its specific data.

[0066] Various methods can be used to adjust the right ascension of the ascending node in the initial orbital parameters. For example, the user can manually adjust the parameters, and the terminal can receive the user's adjustment information and adjust the right ascension of the ascending node in the initial orbital parameters based on the adjustment information. Alternatively, an adjustment program can be preset in the terminal to automatically adjust the right ascension of the ascending node in the initial orbital parameters according to the adjustment program.

[0067] In one feasible approach, adjusting the right ascension of the ascending node in the initial orbital parameters includes:

[0068] Receive adjustment information input by the user;

[0069] Adjust the right ascension of the ascending node in the initial orbital parameters according to the adjustment information.

[0070] 104. Adjust the perigee argument in the initial orbital parameters;

[0071] When the orbit is a near-circular orbit, the perigee argument refers to the satellite's position on its elliptical orbit.

[0072] The Earth rotates at a speed of 15 degrees per hour. By subtracting the launch time from the local time of the descending node based on the launch time at the launch site, the length of time the Earth rotates can be obtained. Therefore, this parameter can be adjusted to control the time it takes to pass through a given geographical longitude.

[0073] Various methods can be used to adjust the perigee argument in the initial orbit parameters. For example, the user can manually adjust the parameters, and the terminal can receive the user's adjustment information and adjust the perigee argument in the initial orbit parameters based on the adjustment information. Alternatively, a pre-set adjustment program can be used in the terminal to automatically adjust the perigee argument in the initial orbit parameters according to the adjustment program.

[0074] In one feasible approach, adjusting the perigee argument in the initial orbital parameters includes:

[0075] Receive debugging information input by the user;

[0076] Adjust the perigee argument in the initial orbital parameters based on the aforementioned debugging information.

[0077] 105. Obtain the calculated value of the local time of the descending node based on the satellite's entry point time, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and other parameters in the initial orbital parameters;

[0078] In one possible implementation, such as Figure 2 As shown, step 105 includes the following sub-steps 1051-1052:

[0079] 1051. Calculate the satellite's nadir parameters based on the satellite's entry point time, adjusted right ascension of the ascending node, adjusted argument of perigee, and other parameters in the initial orbital parameters;

[0080] 1052. Divide the longitude of the nadir point in the nadir point parameters when the latitude of the nadir point is 0 by the Earth's rotational angular velocity to obtain the local time of the descending node.

[0081] 106. When the calculated value of the local time at the descending node matches the preset local time at the descending node, the orbital parameter corresponding to the calculated value of the local time at the descending node is determined as the orbital element of the sun-synchronous satellite.

[0082] 107. If the calculated local time of the descending node does not match the preset local time of the descending node, repeat the steps of adjusting the right ascension of the ascending node and adjusting the argument of the perigee until the orbital parameters corresponding to the calculated local time of the determined descending node match the preset local time of the descending node.

[0083] Specifically, based on the initial orbital parameters, the satellite's nadir parameters can be extrapolated. When the nadir latitude is 0, the corresponding nadir longitude is divided by the Earth's rotational angular velocity of 15 degrees / hour to obtain the accurate descending node local time. Based on the designed descending node local time, steps 103 and 104 are repeated until the designed descending node local time is met.

[0084] This disclosure provides a method for obtaining the elements of a sun-synchronous satellite's orbital insertion point, comprising: calculating initial orbital parameters based on the epoch time of the sun-synchronous satellite, a preset orbital altitude, and a preset local time of the descending node; calculating the time of the satellite's orbital insertion point; adjusting the right ascension of the ascending node in the initial orbital parameters according to the location of the launch site, so that the satellite's nadir trajectory passes near the geodetic coordinates of the launch site; adjusting the argument of perigee in the initial orbital parameters; obtaining the calculated value of the local time of the descending node based on the time of the satellite's orbital insertion point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and other parameters in the initial orbital parameters; when the calculated value of the local time of the descending node matches the preset local time of the descending node, determining the orbital parameters corresponding to the calculated value of the local time of the descending node as the elements of the sun-synchronous satellite's orbital insertion point; when the calculated value of the local time of the descending node does not match the preset local time of the descending node, re-executing the steps of adjusting the right ascension of the ascending node and adjusting the argument of perigee until the orbital parameters corresponding to the determined calculated value of the local time of the descending node match the preset local time of the descending node.

[0085] The method disclosed herein can extrapolate the orbital elements of a sun-synchronous satellite based on information such as the satellite's orbital altitude, the local time of the descending node, and the location of the launch site. It has high accuracy and provides essential orbital data for early-stage satellite tracking and forecasting calculations, thus providing strong support for the scheduling and decision-making of commercial satellite tracking and control plans.

[0086] Based on the above Figure 1 The method for obtaining the orbital elements of a sun-synchronous satellite as described in the corresponding embodiments is described below as an embodiment of the apparatus of this disclosure, which can be used to execute the method embodiments of this disclosure.

[0087] This disclosure provides a device for obtaining the orbital elements of a sun-synchronous satellite, such as... Figure 3 As shown, the device for obtaining the orbital elements of a sun-synchronous satellite includes:

[0088] The first calculation module 11 is used to calculate the initial orbit parameters based on the epoch time of the sun-synchronous satellite, the preset orbital altitude, and the local time of the preset descending node.

[0089] The second calculation module 12 is used to calculate the time of the satellite's orbital insertion point;

[0090] The first adjustment module 13 is used to adjust the right ascension of the ascending node in the initial orbital parameters according to the location of the launch site, so that the nadir trajectory of the satellite passes near the geodetic coordinates of the launch site;

[0091] The second adjustment module 14 is used to adjust the perigee argument in the initial orbital parameters;

[0092] The acquisition module 15 is used to obtain the calculated value of the local time of the descending node based on the time of the satellite's entry point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and other parameters in the initial orbital parameters;

[0093] The determining module 16 is used to determine the orbital parameters corresponding to the calculated value of the descending node local time as the sun-synchronous satellite insertion point element when the calculated value of the descending node local time matches the preset descending node local time.

[0094] The loop module 17 is used to re-execute the steps of adjusting the right ascension of the ascending node and adjusting the argument of the perigee when the calculated value of the local time of the descending node does not match the preset local time of the descending node, until the orbital parameters corresponding to the calculated value of the local time of the descending node match the preset local time of the descending node.

[0095] In one embodiment, such as Figure 4 As shown, the second computing module 12 includes:

[0096] The first acquisition submodule 121 is used to acquire the time of the satellite's orbital insertion point based on the launch time at the launch site and the duration of the active phase flight.

[0097] In one embodiment, such as Figure 5 As shown, the first acquisition submodule 121 includes:

[0098] The acquisition subunit 1211 is used to determine the time of the satellite's orbit insertion point by summing the launch time at the launch site and the active phase flight duration.

[0099] In one embodiment, such as Figure 6 As shown, the acquisition module 15 includes:

[0100] The calculation submodule 151 is used to calculate the satellite's nadir parameters based on the time of the satellite's entry point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and the initial orbital parameters.

[0101] The second acquisition submodule 152 is used to divide the longitude of the nadir point in the nadir point parameters corresponding to the latitude of the nadir point being 0 by the Earth's rotation angular velocity to obtain the calculated value of the local time of the descending node.

[0102] Based on the above Figure 1 The method for obtaining the orbital elements of a sun-synchronous satellite as described in the corresponding embodiments also provides a computer-readable storage medium in this disclosure. For example, a non-transitory computer-readable storage medium may be a read-only memory (ROM), a random access memory (RAM), a CD-ROM, magnetic tape, a floppy disk, or an optical data storage device. This storage medium stores computer instructions for executing the above-described... Figure 1 The method for obtaining the orbital elements of a sun-synchronous satellite as described in the corresponding embodiments will not be repeated here.

[0103] Other embodiments of this disclosure will readily occur to those skilled in the art upon consideration of the specification and practice of the disclosure herein. This application is intended to cover any variations, uses, or adaptations of this disclosure that follow the general principles of this disclosure and include common knowledge or customary techniques in the art not disclosed herein. The specification and examples are to be considered exemplary only, and the true scope and spirit of this disclosure are indicated by the following claims.

[0104] It should be understood that this disclosure is not limited to the precise structures described above and shown in the accompanying drawings, and various modifications and changes can be made without departing from its scope. The scope of this disclosure is limited only by the appended claims.

Claims

1. A method for obtaining the orbital point number of a sun-synchronous satellite, characterized in that, The method comprises: calculating initial orbit parameters according to a sun-synchronous satellite epoch time, a preset orbit height and a preset descending node local time; calculating a satellite orbit injection point time; adjusting the initial orbit parameters according to the position of the launch site, so that the subsatellite point trajectory of the satellite passes through the vicinity of the geodetic coordinates of the launch site; adjusting the argument of perigee in the initial orbit parameters; calculating a calculated value of the descending node local time according to the satellite orbit injection point time, the adjusted argument of the ascending node, the adjusted argument of the perigee and other parameters in the initial orbit parameters; when the calculated value of the descending node local time matches the preset descending node local time, determining that the orbit parameters corresponding to the calculated value of the descending node local time are the sun-synchronous satellite orbit injection point parameters; when the calculated value of the descending node local time does not match the preset descending node local time, re-executing the steps of adjusting the argument of the ascending node and adjusting the argument of the perigee until the orbit parameters corresponding to the calculated value of the descending node local time match the preset descending node local time.

2. The method of claim 1, wherein, The method comprises: calculating a satellite orbit injection point time according to the launch time of the launch site and the flight time of the main active stage.

3. The method of claim 2, wherein, The method comprises: determining the sum of the launch time of the launch site and the flight time of the main active stage as the satellite orbit injection point time.

4. The method of claim 1, wherein, The method comprises: calculating a satellite orbit injection point time according to the launch time of the launch site and the flight time of the main active stage. The method comprises:

5. A device for obtaining the orbital elements of a sun-synchronous satellite, characterized in that, calculating a satellite orbit injection point time according to the launch time of the launch site and the flight time of the main active stage. The method comprises: calculating a satellite orbit injection point time according to the launch time of the launch site and the flight time of the main active stage. The device comprises: a first calculation module configured to calculate initial orbit parameters according to a sun-synchronous satellite epoch time, a preset orbit height and a preset descending node local time; a second calculation module configured to calculate a satellite orbit injection point time; a first adjustment module configured to adjust the initial orbit parameters according to the position of the launch site, so that the subsatellite point trajectory of the satellite passes through the vicinity of the geodetic coordinates of the launch site; a second adjustment module configured to adjust the argument of perigee in the initial orbit parameters; an acquisition module configured to calculate a calculated value of the descending node local time according to the satellite orbit injection point time, the adjusted argument of the ascending node, the adjusted argument of the perigee and other parameters in the initial orbit parameters; a determination module configured to determine that the orbit parameters corresponding to the calculated value of the descending node local time are the sun-synchronous satellite orbit injection point parameters when the calculated value of the descending node local time matches the preset descending node local time; and a determination module configured to determine that the orbit parameters corresponding to the calculated value of the descending node local time are the sun-synchronous satellite orbit injection point parameters when the calculated value of the descending node local time matches the preset descending node local time. The circulation module is configured to re-perform the steps of adjusting the right ascension of the ascending node and adjusting the argument of perigee until the calculated value of the local mean time of the descending node corresponds to the preset local mean time of the descending node when the calculated value of the local mean time of the descending node does not match the preset local mean time of the descending node.

6. The apparatus of claim 5, wherein, The second calculation module comprises: The first acquisition sub-module is configured to acquire the time of the satellite orbit injection point according to the launch time of the launch site and the flight duration of the active stage.

7. The apparatus of claim 6, wherein, The first acquisition sub-module comprises: The acquisition sub-unit is configured to determine the sum of the launch time of the launch site and the flight duration of the active stage as the time of the satellite orbit injection point.

8. The apparatus of claim 5, wherein, The acquisition module comprises: The calculation sub-module is configured to calculate the subsatellite point parameters of the satellite according to the time of the satellite orbit injection point, the adjusted right ascension of the ascending node, the adjusted argument of perigee, and other parameters in the initial orbit parameters. The second acquisition sub-module is configured to divide the subsatellite longitude in the subsatellite point parameters corresponding to the subsatellite latitude being 0 by the earth rotation angular velocity to acquire the calculated value of the local mean time of the descending node.

9. A device for obtaining the orbital elements of a sun-synchronous satellite, characterized in that, comprise: a processor; a memory for storing processor-executable instructions; wherein the processor is configured to perform the method of any one of claims 1 to 4.

10. A computer readable storage medium having stored thereon computer instructions, wherein, The instructions, when executed by the processor, implement the method of any one of claims 1 to 4.

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