A satellite constellation networking launching method, device, medium and equipment

Abstract

The application provides a satellite constellation networking launching method, comprising the following steps: determining a first launching Julian date which satisfies a target satellite orbit plane under a specified launching point; determining a target orbit entry position difference according to a first latitude amplitude angle of the target satellite at the first launching Julian date and a latitude amplitude angle of an orbit entry point; correcting the first launching Julian date according to the target orbit entry position difference to obtain a target launching Julian date; and determining a target launching point according to the target orbit entry position difference, a longitude of the specified launching point, an earth rotation angular rate, an advance angle rate of the target satellite orbit plane and an orbit motion angle rate of the target satellite. Thus, the target launching time and the target launching point which do not rely on the satellite itself to adjust the orbit phase are determined, and after the target satellite is launched according to the target launching time and the target launching point, the satellite itself does not need to adjust the phase to correctly enter the orbit, the purpose of networking when the satellite enters the orbit is achieved, and the efficiency of the satellite constellation networking is improved.

Description

Technical Field

[0001] This application relates to the field of aerospace vehicle technology, and in particular to a satellite constellation networking launch method, apparatus, medium and equipment. Background Technology

[0002] Traditional satellite constellation launch methods involve sending multiple satellites into the same orbit at once and releasing them simultaneously, then relying on the satellites' own propulsion to adjust to the correct position in the orbital plane. However, due to the limited velocity increment of the satellites, the phasing process often takes a long time. If orbital phase adjustment is performed using the rocket's final stage, it requires a long coasting time and multiple power-on / off operations, which presents various difficulties for rockets using cryogenic propellants, thus affecting launch efficiency. For rockets using ambient temperature propellants, it places higher demands on the rocket's final stage batteries and other hardware, and also reduces the rocket's payload capacity.

[0003] With the rise of efficient and flexible small solid-propellant launch vehicles, the launch point can theoretically vary within a certain range. Therefore, if a launch time and launch point that can be found without relying on the satellite's own power to adjust the orbital phase is found, it will be possible for the satellite to enter the correct position immediately after the launch vehicle separates from orbit.

[0004] Therefore, there is a need for a satellite constellation networking and launch method to meet the potential demand for rapid network replenishment of future satellite constellations and improve the efficiency of satellite constellation networking. Summary of the Invention

[0005] To address the problems existing in the prior art, embodiments of the present invention provide a satellite constellation networking launch method, apparatus, medium, and equipment to solve or partially solve the technical problem that the networking or supplementary network launch efficiency cannot be guaranteed when launching satellite constellations or supplementary networks in the prior art.

[0006] A first aspect of the present invention provides a method for launching satellite constellations, the method comprising:

[0007] Obtain the specified epoch of the target satellite and the instantaneous orbital elements at the specified epoch, convert the instantaneous orbital elements into the average orbital elements in the Earth-fixed coordinate system, and obtain the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is the satellite to be networked.

[0008] Based on the average orbital elements, determine the geographical longitude of the ascending node, the latitude argument of the insertion point, and the time from launch to insertion into orbit at the designated launch point;

[0009] The first Julian launch day that satisfies the target satellite's orbital plane is determined based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the launch time to orbit, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate.

[0010] Determine the first latitude argument of the target satellite on the first Julian day of launch, and determine the target orbital position difference based on the first latitude argument and the latitude argument of the orbital insertion point;

[0011] The first launch Julian date is corrected based on the target orbital position difference to obtain the target launch Julian date;

[0012] The target launch point is determined based on the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite.

[0013] In the above scheme, determining the first Julian launch day that satisfies the target satellite's orbital plane at the designated launch point based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the launch time to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate includes:

[0014] Extract the geographical longitude of the ascending node at a specified epoch from the average orbital elements, according to the formula ΔΩ=Ω m -Ω m0 Determine the geographical longitude difference ΔΩ;

[0015] According to the formula ΔT0=2kπ / (ω e -W Ω Determine the duration ΔT0 after the target satellite's orbital plane has precessed k revolutions relative to the Earth from the specified epoch;

[0016] According to the formula JD1=JD0+(ΔΩ / (ω) e -W Ω )-T inj +ΔT0) determines the first Julian launch day JD1 at the designated launch point that satisfies the target satellite's orbital plane; where...

[0017] The Ω m To specify the geographical longitude of the ascending node at a given epoch, the Ω m0 The geographical longitude of the ascending node of the target satellite launched into orbit from the specified launch point, k is the number of orbits the target satellite's orbital plane has precessed relative to the Earth since the specified epoch, JD0 is the second Julian day corresponding to the target satellite's orbital epoch, and ω... e The Earth's rotation angular rate, W ΩThe T is the precession angular rate of the target satellite's orbital plane. inj The time from launch to orbit insertion is denoted as .

[0018] In the above scheme, determining the first latitude argument of the target satellite on the first Julian launch day includes:

[0019] According to formula u s =u m +ΔΩ×W orbit / (ω e -W Ω )+ΔT0×W orbit Determine the first latitude argument u s ;in,

[0020] The u m The target satellite is the average latitude argument in the Earth-fixed coordinate system, ΔΩ is the geographic longitude difference, and W is the mean latitude argument. orbit The angular rate of motion of the target satellite's orbital plane, ω e The Earth's rotation angular rate, W Ω The target satellite's orbital plane precession angular rate is ΔT0, which is the time elapsed after the target satellite's orbital plane has precessed k revolutions relative to the Earth since the specified epoch.

[0021] In the above scheme, before determining the first latitude argument of the target satellite on the first Julian day of launch, the method further includes:

[0022] According to formula W orbit =W M +W ω Determine the angular rate W of motion on the target satellite's orbital plane. orbit ,in,

[0023] The W M The W represents the rate of change of the mean perihelion angle of the target satellite's orbit. ω The perigee angle precession rate of the target satellite.

[0024] In the above scheme, determining the target orbital insertion position difference based on the first latitude argument and the orbital insertion point latitude argument includes:

[0025] Using the formula Δu=u0-u s Determine the initial orbital position difference Δu;

[0026] Using formula The initial orbital insertion position difference is corrected to obtain the target orbital insertion position difference Δu′; wherein,

[0027] u0 is the latitude argument of the orbital entry point; u s This is the first latitude argument.

[0028] In the above scheme, the step of correcting the first launch Julian date based on the orbital position difference to obtain the target launch Julian date includes:

[0029] According to the formula The target launch point is determined to be Julian Japan JD; whereby,

[0030] JD1 is the first Julian launch day at the designated launch point that satisfies the target satellite's orbital plane; Δu′ is the target's orbital insertion position difference; W... orbit The angular velocity of the target satellite's orbital plane is t, where t is the number of seconds contained in a day.

[0031] In the above scheme, determining the target launch point based on the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite includes:

[0032] According to the formula L=L0-Δu′×(ω) e -W Ω ) / W orbit Determine the longitude L of the target launch point;

[0033] The latitude of the target launch point is determined based on the latitude of the designated launch point; the latitude of the target launch point is consistent with the latitude of the designated launch point.

[0034] The target launch point is determined based on its longitude and latitude; wherein...

[0035] L0 is the longitude of the designated launch point, Δu′ is the target orbital position difference, and ω e The Earth's rotation angular rate, W Ω The W represents the precession angular rate of the target satellite's orbital plane. orbit The angular rate of motion of the target satellite's orbital plane.

[0036] A second aspect of the present invention provides a satellite constellation networking launch device, the device comprising:

[0037] The acquisition unit is used to acquire a specified epoch of the target satellite and the instantaneous orbital elements at the specified epoch, convert the instantaneous orbital elements into the average orbital elements in the Earth-fixed coordinate system, and acquire the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is a satellite to be networked.

[0038] The first determining unit is configured to determine the geographical longitude of the ascending node, the latitude argument of the launch point, and the time taken from launch to orbit insertion based on the average orbital elements; determine the first Julian date for launch at the designated launch point that satisfies the target satellite's orbital plane based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the time taken from launch to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate; determine the first latitude argument of the target satellite on the first Julian date; and determine the target orbit insertion position difference based on the first latitude argument and the latitude argument of the launch point.

[0039] The correction unit is used to correct the first launch Julian date based on the target orbital position difference to obtain the target launch Julian date;

[0040] The second determining unit is used to determine the target launch point based on the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite.

[0041] A third aspect of the present invention provides a computer-readable storage medium having a computer program stored thereon that, when executed by a processor, implements the steps of the method described in any of the first aspects.

[0042] A fourth aspect of the present invention provides a computer device including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor, when executing the program, implements the steps of the method described in any of the first aspects.

[0043] This invention provides a satellite constellation networking launch method, comprising: obtaining a specified epoch of a target satellite and instantaneous orbital elements at the specified epoch; converting the instantaneous orbital elements into average orbital elements in an Earth-fixed coordinate system; and obtaining the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; wherein the target satellite is a satellite to be networked; determining the geographical longitude of the ascending node of the launch point, the latitude argument of the launch point, and the launch-to-orbit duration based on the average orbital elements; determining the first Julian day for launch at the specified launch point that satisfies the target satellite's orbital plane based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the launch-to-orbit duration, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate; and determining the target satellite's launch date on the first launch date. The first latitude argument of the Julian Day is used to determine the target orbit insertion position difference based on the first latitude argument and the latitude argument of the insertion point. The first Julian Day of launch is corrected based on the target orbit insertion position difference to obtain the target Julian Day of launch. The target launch point is determined based on the target orbit insertion position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite. In this way, by determining the target launch time and target launch point without relying on the satellite's own power to adjust the orbital phase, the satellite can correctly enter orbit without needing to adjust its phase after launch, achieving the goal of satellite constellation networking upon entering orbit, thereby improving the efficiency of satellite constellation networking. Attached Figure Description

[0044] Various other advantages and benefits will become apparent to those skilled in the art upon reading the following detailed description of preferred embodiments. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings.

[0045] In the attached diagram:

[0046] Figure 1 A schematic flowchart of a satellite constellation networking launch method according to an embodiment of the present invention is shown;

[0047] Figure 2 A schematic diagram of a satellite constellation networking launch device according to an embodiment of the present invention is shown. Detailed Implementation

[0048] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.

[0049] To better understand the technical solution of this invention, the principles upon which this invention is based will be introduced first:

[0050] The constraint of a satellite's orbital plane is essentially a constraint of the geographic longitude of its ascending node. If the launch point is fixed, considering only descending orbit launches, there exists a solution within each Earth rotation cycle. If the requirement of satellite orbital phase is added, the set of launch points satisfying the conditions will form a curve on the ground resembling a nadir trajectory. Therefore, for any location on the ground, to meet the requirement of completing the network upon entering orbit, theoretically these trajectories need to exactly sweep across the designated location. However, because these trajectories are discrete, filling the entire sphere (a strip of land within a certain latitude range) would require an infinite amount of time. Therefore, theoretically, an infinite amount of time is needed to ensure that the trajectories exactly sweep across the designated location. However, if the launch vehicle itself is allowed to maneuver within a certain range (allowing the launch point to change within a certain range), then a suitable launch point and corresponding launch time can be found within each Earth rotation cycle.

[0051] Based on this, to meet the requirements of rapid satellite constellation deployment and replenishment, this invention provides a satellite constellation deployment and launch method, and a method for calculating launch time and launch point that simultaneously satisfies the requirements of satellite orbital plane and orbital phase, while ensuring that the target launch point is not far from the pre-specified launch location, and providing a solution within each Earth rotation cycle. Furthermore, this method is compatible with special cases such as weak constraints where the launch point satisfies orbital plane constraints, and constraints where the geographical longitude of the specified ascending node is specified, providing the correct launch time under these special circumstances.

[0052] Specifically, such as Figure 1 As shown, the method mainly includes the following steps:

[0053] S110, obtain the specified epoch of the target satellite and the instantaneous orbital elements at the specified epoch, convert the instantaneous orbital elements into the average orbital elements in the Earth-fixed coordinate system, and obtain the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is a satellite to be networked.

[0054] First, the latitude and longitude coordinates (L0, B0) of the designated launch point need to be obtained. As mentioned above, the designated launch point cannot simultaneously satisfy the constraints of the satellite orbital plane and phase. Therefore, this invention will determine the longitude of the target launch point under the same latitude B0.

[0055] Then, obtain the instantaneous orbital elements (J2000 orbital six elements) at a specified epoch and at a specified epoch (year Y0, month Mth0, day D0, hour H0, minute MIN0, second SEC0) of the target satellite. The instantaneous orbital elements include: the semi-major axis a of the first orbit. 2k First orbital eccentricity e 2k First orbital inclination angle i 2k Argument ω of the first perigee 2k The geographical longitude of the first ascending node Ω 2k First angle of approach M 2k .

[0056] The designated launch point, designated epoch, and instantaneous orbital elements of the target satellite can be determined based on corresponding data from historically successfully launched satellites (the designated launch point of the target satellite must be the same as the launch point of historically successfully launched satellites, the designated epoch of the target satellite must be the same as the epoch of historically successfully launched satellites, and the instantaneous orbital elements at the designated epoch must be the same as the instantaneous orbital elements of historically successfully launched satellites). It should be noted that the orbital shape of historically successfully launched satellites must be identical to the orbital shape of the target satellite.

[0057] The instantaneous orbital elements are converted into average orbital elements in the Earth-fixed coordinate system. The average orbital elements include: the second orbital semi-major axis a. m Second orbital eccentricity e m Second orbital inclination angle i m Second perigee argument ω m Geographic longitude Ω of the ascending node at a specified epoch m The second near-point angle M m .

[0058] It should be noted that after converting the instantaneous orbital elements to average orbital elements in the Earth-fixed coordinate system, the average latitudinal argument u of the target satellite in the Earth-fixed coordinate system will also be obtained. m .

[0059] S111, based on the average orbital elements, determine the geographical longitude of the ascending node of the launch point, the latitude argument of the launch point, and the time from launch to orbit insertion.

[0060] In one implementation, determining the geographical longitude of the ascending node, the latitude argument of the orbit insertion point, and the time from launch to orbit insertion at a designated launch point based on the average orbital elements includes:

[0061] The ballistic program calculates the trajectory based on the average orbital elements and obtains the calculation results, which include: the geographical longitude of the ascending node of the launch point, the latitude argument of the launch point, and the time from launch to orbit.

[0062] Specifically, once the average orbital elements are determined, the target orbital parameters of the target satellite can be set based on these elements. Specifically, the semi-major axis of the target orbit is set to a. m Set the eccentricity of the target orbit to e m Set the inclination angle of the target orbit to i. m Set the perigee argument of the target orbit to ω. m .

[0063] Then, based on the target orbital parameters, the ballistic calculation program is run to obtain the geographical longitude Ω of the ascending node of the launch point into orbit. m0 The latitude and argument of the orbital insertion point, u0, and the time T from launch to orbital insertion. inj .

[0064] As can be seen, the target trajectory parameters are in the Earth-fixed coordinate system, so they are not affected by the predetermined launch time when performing ballistic calculations.

[0065] Then, the precession rate W of the target satellite's orbital plane is calculated based on the average orbital elements. Ω Target satellite perigee angle precession rate W ω and the rate of change of the mean perihelion angle of the target satellite orbit W M Among them, W Ω W ω and

[0066] W M The specific calculation method is well known to those skilled in the art, so it will not be elaborated here.

[0067] S112, determine the first Julian launch date that satisfies the target satellite's orbital plane at the designated launch point based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the launch time to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate.

[0068] In one implementation, the first Julian launch day satisfying the target satellite's orbital plane at the designated launch point is determined based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the time from launch to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate. This includes:

[0069] Extract the geographical longitude of the ascending node at a specified epoch from the mean orbital elements, using the formula ΔΩ=Ω m -Ω m0 Determine the geographical longitude difference ΔΩ;

[0070] According to the formula ΔT0=2kπ / (ω e -W Ω Determine the duration ΔT0 after the target satellite's orbital plane has precessed k revolutions relative to the Earth since the specified epoch.

[0071] According to the formula JD1=JD0+(ΔΩ / (ω) e -W Ω )-T inj +ΔT0) determines the first Julian launch day JD1 at the designated launch point that satisfies the target satellite's orbital plane; where...

[0072] Ω m Ω represents the geographical longitude of the ascending node in the mean orbital elements. m0 The geographical longitude of the ascending node of the target satellite launched from the designated launch point, k is the number of orbital precessions of the target satellite relative to the Earth starting from the designated epoch, JD0 is the second Julian day corresponding to the target satellite's orbital epoch, and ω e W is the Earth's rotational angular rate. Ω T is the precession angular rate of the target satellite's orbital plane; inj This refers to the time from launch to orbit insertion.

[0073] Specifically, after running the ballistic calculation program, the geographical longitude Ω of the ascending node of the launch point into orbit can also be obtained. m0 Then obtain the geographical longitude Ω of the ascending node at the specified epoch. m Therefore, Ω can be determined according to formula (1). m With Ω m0 The difference in geographical longitude between them is ΔΩ:

[0074] ΔΩ=Ω m -Ω m0 (1)

[0075] To ensure the orbital plane meets the requirements, the Earth needs to rotate relative to a specified epoch. As the Earth rotates, the Ω of the target satellite... m It will change, and Ω m0 This remains unchanged. When the two coincide, it represents the launch time at the designated launch point that satisfies the orbital plane requirements. Here, we introduce another variable, ΔT0, to describe the duration after the target satellite's orbital plane has precessed k revolutions relative to the Earth from the designated epoch:

[0076] ΔT0=2kπ / (ω e -W Ω (2)

[0077] The k value is used to adjust the date. The expression for the first Julian launch date JD1 that satisfies the target satellite's orbital plane at the specified launch point is shown in formula (3):

[0078] JD1=JD0+(ΔΩ / (ω e -W Ω )-T inj +ΔT0) (3)

[0079] This allows the second launch Julian day corresponding to the target satellite's orbital epoch to be directly converted into the first launch Julian day JD1, expressed in the form of year, month, day, hour, minute, and second.

[0080] The above steps yield the launch date and time that meet the orbital plane requirements at the designated launch point. When the target satellite can adjust its phase automatically, the above steps have already yielded the launch time that meets the orbital plane requirements (the launch point is a pre-specified location), which is the case of weak constraints. However, when the requirement is that the satellite is in the correct position upon entering orbit (completing network formation), the phase requirements of the satellite at the orbital insertion point must also be met. This requires adjusting the launch point and continuing with the following steps to determine the launch time and launch point that simultaneously meets both the orbital plane and phase requirements.

[0081] S113, determine the first latitude argument of the target satellite on the first launch Julian day, and determine the difference in target orbit insertion position based on the first latitude argument and the latitude argument of the insertion point.

[0082] In one implementation, before determining the first latitude argument of the target satellite on the first Julian launch day, the method further includes:

[0083] According to formula W orbit =W M +W ω Determine the angular rate W of motion on the target satellite's orbital plane. orbit ,in,

[0084] W M W represents the rate of change of the mean anomaly angle of the target satellite's orbit. ω The perigee angle precession rate of the target satellite.

[0085] In one implementation, determining the first latitude argument of the target satellite on the first Julian launch day includes:

[0086] According to formula u s =u m +ΔΩ×W orbit / (ω e -W Ω )+ΔT0×W orbit Determine the first latitude argument u s ;in,

[0087] u mW represents the average latitudinal argument of the target satellite in the Earth-fixed coordinate system, ΔΩ is the difference in geographic longitude, and W is the mean latitude angle of the target satellite in the Earth-fixed coordinate system. orbit The angular rate of motion of the target satellite's orbital plane, ω e W is the Earth's rotational angular rate. Ω ΔT0 is the precession angular rate of the target satellite's orbital plane, and ΔT0 is the time elapsed after the target satellite's orbital plane has precessed k revolutions relative to the Earth since the specified epoch.

[0088] Specifically, the first step is to determine the angular rate W of the target satellite's orbital plane according to formula (4). orbit :

[0089] W orbit =W M +W ω (4)

[0090] Based on the angular rate of motion of the target satellite's orbital plane, the first latitude argument of the target satellite on the first Julian day is determined according to formula (5):

[0091] u s =u m +ΔΩ×W orbit / (ω e -W Ω )+ΔT0×W orbit (5)

[0092] Adjust the first latitude argument to the range of (-2π, 2π), and compare the orbital insertion latitude argument u0 (in Earth-fixed system) with the first latitude argument u at the time of the first Julian day JD1. s The initial orbital position difference Δu between the two orbits is calculated, and the shortest path processing is performed on the initial orbital position difference Δu.

[0093] The shortest path processing is as follows: Since the satellite orbit is circular with an angle of 360 degrees, when the orbital position difference is 10 degrees, it can reach the correct position by rotating 10 degrees forward or 350 degrees backward. Therefore, in order to minimize the movement of the launch point position at each k value, it is necessary to use formula (6) to process the initial orbital position difference Δu with the shortest path, and obtain the processed target orbital position difference Δu′, as follows:

[0094]

[0095] That is, the target orbital position difference is determined based on the first latitude argument and the orbital entry point latitude argument, including:

[0096] Using the formula Δu=u0-u s Determine the initial orbital position difference Δu;

[0097] Using formula The initial orbital insertion position difference is corrected to obtain the target orbital insertion position difference Δu′; where,

[0098] u0 is the latitude argument of the Earth-fixed system at the orbital entry point; u s This is the first latitude argument.

[0099] S114, the first launch Julian date is corrected according to the target orbital position difference to obtain the target launch Julian date; the target launch point is determined according to the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite.

[0100] Once the target orbital position difference is determined, the target launch Julian date and launch point can be determined based on the target orbital position difference.

[0101] In one implementation, the first launch Julian date is corrected based on the target orbital position difference to obtain the target launch Julian date, including:

[0102] According to the formula The target for the launch of the Julian Japan JD was determined; among them,

[0103] JD1 is the first Julian launch day from the specified launch point that satisfies the target satellite's orbital plane; Δu′ is the target's orbital position difference; W orbit Let t be the angular velocity of the target satellite's orbital plane, and t be the number of seconds contained in a day, with a value of 86400.

[0104] In one implementation, determining the target launch point based on the target orbital insertion position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the target satellite's orbital motion angular rate includes:

[0105] According to the formula L=L0-Δu′×(ω) e -W Ω ) / W orbit Determine the longitude L of the target launch point;

[0106] The latitude of the target launch point is determined based on the latitude of the designated launch point; the latitude of the target launch point is consistent with the latitude of the designated launch point.

[0107] The target launch point is determined based on its longitude and latitude; among which...

[0108] L0 is the longitude of the designated launch point, Δu′ is the target orbital insertion position difference, ω e W is the Earth's rotational angular rate. Ω W is the precession angular rate of the target satellite's orbital plane. orbit The angular rate of motion of the target satellite's orbital plane.

[0109] After the above steps, the target launch point and target launch time that meet the requirements for rapid networking are determined, thus fulfilling the requirement of networking upon entering orbit.

[0110] It should be noted that the above k value can be any integer, allowing for flexible adjustment of the launch date. However, since Earth's gravitational perturbations include not only the J2 perturbation term but also other smaller long-term perturbation terms, it is recommended that the k value be taken near the specified orbital epoch.

[0111] Furthermore, when this method involves longitude shifts, the elevation changes along with the position. If the elevation change is too large, the trajectory of the incoming ballistic missile will change. In this case, by repeating steps S111 to S114 with known elevation information, the target launch time and launch point that meet the requirements can be found.

[0112] Furthermore, the above method employs shortest path processing, ensuring that the movement of longitude does not exceed the angle the Earth rotates within half a satellite cycle. For example, with a circular orbit at an altitude of 500 km, the Earth's rotation angle does not exceed 12° in half a satellite cycle, meaning that the east-west movement at a given latitude does not exceed one time zone. To achieve an even smaller movement, different launch times can be selected.

[0113] For example, suppose the instantaneous six roots of the J2000 orbit of the satellite to be supplemented at 8:22:37.0 on November 15, 2023 are:

[0114] a 2k =6881266.8m

[0115] e 2k =0.000748873

[0116] i 2k =45.066033°

[0117] ω 2k =141.802324°

[0118] Ω 2k =157.083130°

[0119] M 2k =13.640075°

[0120] The designated launch point has the following coordinates: E100.343°, N40.969°, and an elevation of 1011m.

[0121] Following the above method, the trajectory of a solid rocket launched from a designated launch point to orbit is calculated using a ballistic program for a certain type of rocket, yielding the geographical longitude Ω of the ascending node at the designated launch point. m0 =337.833013°, orbital insertion latitude argument u0=153.374065°, duration T from launch to orbital insertion inj = 847.0s. Since the terrain near the designated launch point is flat, the movement of the launch point can be considered as not changing the elevation. After the above operations, the target launch point and target launch time under different k values ​​are shown in Table 1:

[0122] Table 1

[0123]

[0124] As can be seen from Table 1, different k values ​​correspond to different launch points and launch dates, and the amount of movement of each launch point relative to the initially specified launch point is different.

[0125] As mentioned above, the designated launch point, designated epoch, and instantaneous orbital elements at the designated epoch of the satellite to be replenished are determined based on the launch parameters of historically successfully launched satellites with the same orbital shape. In the actual implementation of this invention, the orbit of the satellite that was actually successfully launched is a circular orbit with an inclination of 45° and a height of 500km launched at 8:08 AM on November 15th. This orbit is also used as the replenishment orbit for the target satellite.

[0126] As can be seen from Table 1, the method provided by this invention can reconstruct the actual launch time and actual launch location of historically successfully launched satellites under the condition of k=0, which shows that the algorithm has high accuracy.

[0127] To further verify the accuracy of the algorithm, the launch points and launch times with different k values ​​in Table 1 were used as inputs and substituted into the ballistic program to calculate the orbital trajectory. The corresponding orbital entry point information was then input into the Satellite Tool Kit (STK). The orbital parameters of the target satellite to be networked were also input into the STK tool to obtain the final simulation results.

[0128] Simulation results show that when simulated launches are performed at launch points and launch times with different k values, the orbits of all satellites entering orbit coincide, and the positions of the satellites also coincide. Therefore, it can be concluded that the accuracy of the satellite constellation networking launch method provided by this invention can be ensured.

[0129] Based on the same inventive concept as in the foregoing embodiments, this embodiment also provides a satellite constellation networking launch device, such as... Figure 2 As shown, the device includes:

[0130] The acquisition unit 21 is used to acquire a specified epoch of the target satellite and the instantaneous orbital elements at the specified epoch, convert the instantaneous orbital elements into the average orbital elements in the Earth-fixed coordinate system, and acquire the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is a satellite to be networked.

[0131] The first determining unit 22 is used to determine the geographical longitude of the ascending node of the launch point, the latitude argument of the launch point, and the time from launch to orbit insertion based on the average orbital elements; to determine the first Julian day of launch at the designated launch point that satisfies the target satellite's orbital plane based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the time from launch to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate; to determine the first latitude argument of the target satellite on the first Julian day of launch; and to determine the difference in target orbit insertion position based on the first latitude argument and the latitude argument of the launch point.

[0132] Correction unit 23 is used to correct the first launch Julian date based on the target orbital position difference to obtain the target launch Julian date;

[0133] The second determining unit 24 is used to determine the target launch point based on the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite.

[0134] Since the apparatus described in the embodiments of this invention is used to implement the satellite constellation networking and launch method of the embodiments of this invention, those skilled in the art can understand the specific structure and variations of the apparatus based on the method described in the embodiments of this invention, and therefore will not be described in detail here. All apparatuses used in the methods of the embodiments of this invention fall within the scope of protection of this invention.

[0135] Based on the same inventive concept, this embodiment provides a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements any step of the method described above.

[0136] Based on the same inventive concept, this embodiment provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the steps of any of the methods described above.

[0137] Through one or more embodiments of the present invention, the present invention has the following beneficial effects or advantages:

[0138] This invention provides a satellite constellation networking launch method, comprising: obtaining a specified epoch of a target satellite and instantaneous orbital elements at the specified epoch; converting the instantaneous orbital elements into average orbital elements in an Earth-fixed coordinate system; and obtaining the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is a satellite to be networked; determining the geographical longitude of the ascending node of the launch point, the latitude argument of the launch point, and the time from launch to orbit insertion based on the average orbital elements; determining the first launch Julian day at the specified launch point that satisfies the target satellite's orbital plane based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the time from launch to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate; and determining the first launch Julian day of the first launch Julian day. The first latitude argument of the launch date is determined, and the difference in target launch position is determined based on the first latitude argument and the latitude argument of the launch point. The first launch date is corrected based on the target launch position difference to obtain the target launch date. The target launch point is determined based on the target launch position difference, the longitude of the designated launch point, the Earth's rotation rate, the precession rate of the target satellite's orbital plane, and the orbital motion rate of the target satellite. In this way, by determining the target launch time and target launch point without relying on the satellite's own power to adjust the orbital phase, the satellite can correctly enter orbit without adjusting its own phase after launch, achieving the goal of satellite constellation networking upon entering orbit, thereby improving the efficiency of satellite constellation networking.

[0139] The algorithms and displays provided herein are not inherently related to any particular computer, virtual system, or other device. Various general-purpose systems can also be used in conjunction with the teachings herein. The required structure for constructing such systems is apparent from the above description. Furthermore, this invention is not directed to any particular programming language. It should be understood that the contents of the invention described herein can be implemented using various programming languages, and the above description of specific languages ​​is for the purpose of disclosing the best mode of implementation of the invention.

[0140] Numerous specific details are set forth in the specification provided herein. However, it will be understood that embodiments of the invention may be practiced without these specific details. In some instances, well-known methods, structures, and techniques have not been shown in detail so as not to obscure the understanding of this specification.

[0141] Similarly, it should be understood that, in order to simplify this disclosure and aid in understanding one or more of the various aspects of the invention, in the above description of exemplary embodiments of the invention, various features of the invention are sometimes grouped together in a single embodiment, figure, or description thereof. However, this method of disclosure should not be construed as reflecting an intention that the claimed invention requires more features than are expressly recited in each claim. Rather, as reflected in the following claims, inventive aspects lie in fewer than all features of a single foregoing disclosed embodiment. Therefore, the claims following the detailed description are hereby expressly incorporated into this detailed description, wherein each claim itself is a separate embodiment of the invention.

[0142] Those skilled in the art will understand that modules in the device of the embodiments can be adaptively changed and placed in one or more devices different from that embodiment. Modules, units, or components in the embodiments can be combined into a single module, unit, or component, and further, they can be divided into multiple sub-modules, sub-units, or sub-components. Except where at least some of such features and / or processes or units are mutually exclusive, any combination can be used to combine all features disclosed in this specification (including the accompanying claims, abstract, and drawings) and all processes or units of any method or device so disclosed. Unless expressly stated otherwise, each feature disclosed in this specification (including the accompanying claims, abstract, and drawings) may be replaced by an alternative feature that serves the same, equivalent, or similar purpose.

[0143] Furthermore, those skilled in the art will understand that although some embodiments herein include certain features included in other embodiments but not others, combinations of features from different embodiments are intended to be within the scope of the invention and form different embodiments. For example, in the following claims, any of the claimed embodiments can be used in any combination.

[0144] The various component embodiments of the present invention can be implemented in hardware, or as software modules running on one or more processors, or a combination thereof. Those skilled in the art will understand that microprocessors or digital signal processors (DSPs) can be used in practice to implement some or all of the functions of some or all of the components of the gateway, proxy server, or system according to embodiments of the present invention. The present invention can also be implemented as a device or apparatus program (e.g., a computer program and computer program product) for performing some or all of the methods described herein. Such programs implementing the present invention can be stored on a computer-readable medium or can be in the form of one or more signals. Such signals can be downloaded from an Internet website, provided on a carrier signal, or provided in any other form.

[0145] It should be noted that the above embodiments are illustrative of the invention and not restrictive, and that those skilled in the art can devise alternative embodiments without departing from the scope of the appended claims. In the claims, any reference signs placed between parentheses should not be construed as limiting the claims. The word "comprising" does not exclude the presence of elements or steps not listed in the claims. The word "a" or "an" preceding an element does not exclude the presence of a plurality of such elements. The invention can be implemented by means of hardware comprising several different elements and by means of a suitably programmed computer. In the unit claims enumerating several means, several of these means may be embodied by the same item of hardware. The use of the words first, second, and third, etc., does not indicate any order. These words can be interpreted as names.

[0146] Although preferred embodiments of this application have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments as well as all changes and modifications falling within the scope of this application.

[0147] The above description is merely a preferred embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for launching satellite constellations, characterized in that, The method includes: Obtain the specified epoch of the target satellite and the instantaneous orbital elements at the specified epoch, convert the instantaneous orbital elements into the average orbital elements in the Earth-fixed coordinate system, and obtain the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is a satellite to be networked. Based on the average orbital elements, determine the geographical longitude of the ascending node, the latitude argument of the insertion point, and the time from launch to insertion into orbit at the designated launch point; The first Julian launch day that satisfies the target satellite's orbital plane is determined based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the launch time to orbit, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate. Determine the first latitude argument of the target satellite on the first Julian day of launch, and determine the target orbital position difference based on the first latitude argument and the latitude argument of the orbital insertion point; The first launch Julian date is corrected based on the target orbital position difference to obtain the target launch Julian date; The target launch point is determined based on the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite.

2. The method as described in claim 1, characterized in that, The determination of the first Julian launch day satisfying the target satellite's orbital plane at the designated launch point, based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the launch time to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate, includes: Extract the geographical longitude of the ascending node at a specified epoch from the average orbital elements, according to the formula ΔΩ=Ω m -Ω m0 Determine the geographical longitude difference ΔΩ; According to the formula ΔT0=2kπ / (ω e -W Ω Determine the duration ΔT0 after the target satellite's orbital plane has precessed k revolutions relative to the Earth from the specified epoch; According to the formula JD1=JD0+(ΔΩ / (ω) e -W Ω )-T inj +ΔT0) determines the first launch Julian day JD1 at the designated launch point that satisfies the target satellite's orbital plane; where... The Ω m To specify the geographical longitude of the ascending node at a given epoch, the Ω m0 The geographical longitude of the ascending node of the target satellite launched into orbit from the specified launch point, k is the number of orbits the target satellite's orbital plane has precessed relative to the Earth since the specified epoch, JD0 is the second Julian day corresponding to the target satellite's orbital epoch, and ω... e The Earth's rotation angular rate, W Ω The T is the precession angular rate of the target satellite's orbital plane. inj The time from launch to orbit insertion is described.

3. The method as described in claim 1, characterized in that, Determining the first latitude argument of the target satellite on the first Julian launch day includes: According to formula u s =u m +ΔΩ×W orbit / (ω e -W Ω )+ΔT0×W orbit Determine the first latitude argument u s ;in, The u m The target satellite is the average latitude argument in the Earth-fixed coordinate system, ΔΩ is the geographic longitude difference, and W is the mean latitude argument. orbit The angular rate of motion of the target satellite's orbital plane, ω e The Earth's rotation angular rate, W Ω The target satellite's orbital plane precession angular rate is ΔT0, which is the time elapsed after the target satellite's orbital plane has precessed k revolutions relative to the Earth since the specified epoch.

4. The method as described in claim 3, characterized in that, Before determining the first latitude argument of the target satellite on the first Julian launch day, the method further includes: According to formula W orbit =W M +W ω Determine the angular rate W of motion on the target satellite's orbital plane orbit ,in, The W M The W represents the rate of change of the mean perihelion angle of the target satellite's orbit. ω The perigee angle precession rate of the target satellite.

5. The method as described in claim 1, characterized in that, The step of determining the target orbital insertion position difference based on the first latitude argument and the orbital insertion point latitude argument includes: Using the formula Δu=u0-u s Determine the initial orbital position difference Δu; Using formula The initial orbital insertion position difference is corrected to obtain the target orbital insertion position difference Δu′; wherein, u0 is the latitude argument of the orbital entry point; u s This is the first latitude argument.

6. The method as described in claim 1, characterized in that, The step of correcting the first launch Julian date based on the target orbital position difference to obtain the target launch Julian date includes: According to the formula The target launch point is determined to be Julian Japan JD; whereby, JD1 is the first Julian launch day at the designated launch point that satisfies the target satellite's orbital plane; Δu′ is the target's orbital insertion position difference; W... orbit The angular velocity of the target satellite's orbital plane is t, where t is the number of seconds contained in a day.

7. The method as described in claim 1, characterized in that, The step of determining the target launch point based on the target orbital insertion position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite includes: According to the formula L=L0-Δu′×(ω) e -W Ω ) / W orbit Determine the longitude L of the target launch point; The latitude of the target launch point is determined based on the latitude of the designated launch point; the latitude of the target launch point is consistent with the latitude of the designated launch point. The target launch point is determined based on its longitude and latitude; wherein... L0 is the longitude of the designated launch point, Δu′ is the target orbital position difference, and ω e The Earth's rotation angular rate, W Ω The W represents the precession angular rate of the target satellite's orbital plane. orbit The angular rate of motion of the target satellite's orbital plane.

8. A satellite constellation networking launch device, characterized in that, The device includes: The acquisition unit is used to acquire a specified epoch of the target satellite and the instantaneous orbital elements at the specified epoch, convert the instantaneous orbital elements into the average orbital elements in the Earth-fixed coordinate system, and acquire the average latitudinal argument of the target satellite in the Earth-fixed coordinate system; the target satellite is a satellite to be networked. The first determining unit is configured to: determine the geographical longitude of the ascending node of the launch point, the latitude argument of the launch point, and the time taken from launch to orbit insertion based on the average orbital elements; determine the first Julian date for launch at the designated launch point that satisfies the target satellite's orbital plane based on the average orbital elements, the geographical longitude of the ascending node of the launch point, the time taken from launch to orbit insertion, the precession rate of the target satellite's orbital plane, and the Earth's rotation rate; determine the first latitude argument of the target satellite on the first Julian date; and determine the target orbit insertion position difference based on the first latitude argument and the latitude argument of the launch point. The correction unit is used to correct the first launch Julian date based on the target orbital position difference to obtain the target launch Julian date; The second determining unit is used to determine the target launch point based on the target orbital position difference, the longitude of the designated launch point, the Earth's rotation angular rate, the precession angular rate of the target satellite's orbital plane, and the orbital motion angular rate of the target satellite.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by a processor, the program implements the steps of the method according to any one of claims 1-7.

10. A computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the program, it implements the steps of the method according to any one of claims 1-7.

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