Constellation satellite deployment method and device, electronic equipment, storage medium and product

Abstract

The application discloses a constellation satellite deployment method and device, electronic equipment, storage medium and product. In the multi-satellite launch task of a large constellation, the ascending node right ascension and phase angle of the satellite are synchronously adjusted by using the continuous small thrust configured on the satellite. First, a timing control strategy is proposed, which uses the influence of the semi-major axis on the ascending node right ascension and phase angle to ensure that the injected satellite is sequentially transferred from the starting orbit to the target orbit. In addition, the coupling relationship between the ascending node right ascension and the phase angle during the maneuver is studied, a mathematical model describing the drift of the ascending node right ascension, the drift of the phase angle, the orbit inclination compensation and the relationship between the maneuver time interval is established, the least square method is used to calculate the mathematical model, and the maneuver time interval and the orbit inclination compensation value are obtained. Finally, the timing control and inclination compensation strategies are combined to realize high-precision synchronous deployment of constellation satellites, and the application potential of small thrust in large-scale orbit control is shown.

Description

Technical Field

[0001] This application belongs to the field of aerospace technology, and in particular relates to a method, apparatus, electronic equipment, storage medium and product for deploying a constellation of satellites. Background Technology

[0002] Currently, an increasing number of mega-constellations are being deployed in low Earth orbit, providing high-speed telecommunications services globally. The deployment of mega-constellations refers to placing satellites at predetermined positions for their ascending node, right ascension, and phase angle.

[0003] However, both the right ascension of the ascending node and the phase angle are related to the semi-major axis of the satellite's orbit, and the right ascension of the ascending node and the phase angle are coupled together. During the deployment of a constellation of satellites, it is impossible to indirectly control both the right ascension of the ascending node and the phase angle simultaneously by controlling the semi-major axis.

[0004] In related technologies, a timing-based control method is typically employed. This method utilizes the Earth's non-spherical perturbation J2 term to first separate the right ascension of the ascending node. After the target satellite enters the target orbital plane, a three-stage maneuver (i.e., ascending / descending orbit—waiting—descending / ascending orbit) is used to adjust the phase angle on the same orbital plane. However, this method cannot synchronously adjust the right ascension of the ascending node and the phase angle, thus reducing the efficiency of constellation satellite deployment. Summary of the Invention

[0005] This application provides a method, apparatus, electronic device, storage medium, and product for deploying constellation satellites, which can compensate for the phase angle and right ascension of the ascending node of the constellation satellites during maneuvers, thereby achieving simultaneous adjustment of the phase angle and right ascension of the ascending node and improving the deployment efficiency of the constellation satellites.

[0006] In a first aspect, embodiments of this application provide a method for deploying a constellation of satellites. The method includes: determining the satellite maneuvering time for the constellation satellites to move from their current orbit to the target orbit based on the satellite maneuvering position information in their current orbit and the target position information in the target orbit; determining a phase adjustment time for adjusting the phase of the satellites based on the satellite maneuvering time and the phase difference between two adjacent constellation satellites in the current orbit; determining a right ascension adjustment time for adjusting the right ascension of the ascending node based on the satellite maneuvering time and the difference in right ascension of the ascending node, wherein the difference in right ascension of the ascending node is the difference between the right ascension of the ascending node of the current orbit and the right ascension of the ascending node of the target orbit; determining an orbital inclination compensation value for the current orbit based on the right ascension offset corresponding to the phase adjustment time, wherein the orbital inclination compensation value is used to compensate for the phase angle and right ascension of the ascending node of the current orbit during the satellite maneuvering time; and deploying the constellation of satellites in the target orbit based on the satellite maneuvering time, the phase adjustment time, the right ascension adjustment time, and the orbital inclination compensation value.

[0007] Secondly, embodiments of this application provide a satellite constellation deployment apparatus, comprising: a time determination module, used to determine the satellite maneuvering time for the constellation satellite to move from the current orbit to the target orbit based on the satellite's parking position information in the current orbit and the target position information in the target orbit; a first determination module, used to determine the phase adjustment time for adjusting the phase of the satellite based on the satellite maneuvering time and the phase difference between two adjacent constellation satellites in the current orbit; a second determination module, used to determine the right ascension adjustment time for adjusting the right ascension of the ascending node based on the satellite maneuvering time and the right ascension difference of the ascending node, wherein the right ascension difference of the ascending node is the difference between the right ascension of the ascending node of the current orbit and the right ascension of the ascending node of the target orbit; a third determination module, used to determine the orbital inclination compensation value of the current orbit based on the right ascension offset of the ascending node corresponding to the phase adjustment time, wherein the orbital inclination compensation value is used to compensate for the phase angle and right ascension of the ascending node of the current orbit during the satellite maneuvering time; and a satellite deployment module, used to deploy the constellation satellites in the target orbit based on the satellite maneuvering time, the phase adjustment time, the right ascension adjustment time, and the orbital inclination compensation value.

[0008] Thirdly, embodiments of this application provide an electronic device, which includes: a processor and a memory storing computer program instructions; the processor executes the computer program instructions to implement the deployment method of the constellation satellites as described in the first aspect.

[0009] Fourthly, embodiments of this application provide a computer-readable storage medium storing computer program instructions that, when executed by a processor, implement the satellite constellation deployment method as described in the first aspect.

[0010] Fifthly, embodiments of this application provide a computer program product in which instructions, when executed by a processor of an electronic device, cause the electronic device to perform the satellite constellation deployment method as described in the first aspect.

[0011] As described above, in this embodiment, the orbital inclination compensation value is determined based on the shift in right ascension of the ascending node caused by phase adjustment. This compensation value is then used to compensate for the ascending node's right ascension and phase angle during the constellation satellite's maneuvers, enabling the satellite to accurately reach its target position in the target orbit. This process only requires compensation for the satellite's phase angle and ascending node's right ascension during maneuvers, achieving synchronous adjustment of both and improving the constellation satellite deployment efficiency. Attached Figure Description

[0012] To more clearly illustrate the technical solutions of the embodiments of this application, the accompanying drawings used in the embodiments of this application will be briefly introduced below. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0013] Figure 1 This is a schematic diagram of satellite deployment in related technologies;

[0014] Figure 2 This is a flowchart illustrating a method for deploying a constellation of satellites according to an embodiment of this application;

[0015] Figure 3 This is a flowchart illustrating a method for deploying a constellation of satellites according to an embodiment of this application;

[0016] Figure 4 This is a timing diagram of the semi-major axis of the lifting orbit of a constellation satellite provided in one embodiment of this application;

[0017] Figure 5 This is a timing diagram of phase angle changes provided in one embodiment of this application;

[0018] Figure 6 This is a schematic diagram of satellite position changes during the phase deployment process provided in one embodiment of this application;

[0019] Figure 7 This is a time-series diagram illustrating the change in the right ascension interval of the ascending node provided in one embodiment of this application;

[0020] Figure 8 This is a schematic diagram showing the deployment results of phase angles and right ascension of the ascending node of 80 satellites in two orbital planes according to one embodiment of this application;

[0021] Figure 9 This is a schematic diagram of the structure of a constellation satellite deployment device provided in another embodiment of this application;

[0022] Figure 10 This is a schematic diagram of the structure of an electronic device provided in another embodiment of this application. Detailed Implementation

[0023] The features and exemplary embodiments of various aspects of this application will be described in detail below. To make the objectives, technical solutions, and advantages of this application clearer, the application will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only intended to explain this application and not to limit it. For those skilled in the art, this application can be implemented without some of these specific details. The following description of the embodiments is merely to provide a better understanding of this application by illustrating examples.

[0024] It should be noted that, in this document, relational terms such as "first" and "second" are used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising..." does not exclude the presence of additional identical elements in the process, method, article, or apparatus that includes said element.

[0025] To facilitate understanding, before explaining the solution provided in this application, the background of the solution provided in this application will be explained first.

[0026] Currently, an increasing number of mega-constellations are being deployed into low Earth orbit to provide high-speed telecommunications services globally. These mega-constellations, consisting of hundreds or even thousands of satellites, have opened up new possibilities for the development of international spaceflight. However, the excessive number of satellites also presents numerous challenges and threatens sustainability. In fact, many constellation providers face serious financial problems, partly due to high development and operating costs, highlighting the importance of adopting highly optimized solutions, especially during the constellation's satellite deployment phase.

[0027] The deployment of a giant constellation involves placing its satellites at predetermined positions for their ascending nodes, right ascension, and phase angles. For example, in... Figure 1 The diagram illustrating satellite deployment shows that the deployment process mainly consists of three phases: Phase 1, Phase 2, and Phase 3. In Phase 1, the satellite launch vehicle (e.g., a rocket) carries the satellite to its parking orbit, waits for a period of time, and then proceeds to Phase 2. In Phase 2, the launch vehicle deploys a group of satellites from the constellation onto orbital plane 1, waits for a period of time, and then proceeds to Phase 3. In Phase 3, the launch vehicle deploys a group of satellites from the constellation onto orbital plane 2, waits for a period of time, and then proceeds to the next orbital plane for satellite deployment, until all satellites are deployed. The total time from Phase 1 to the completion of satellite deployment is the total deployment time.

[0028] Traditional satellite deployment methods include the following three categories:

[0029] (1) Direct deployment of satellites carried by rockets. In this method, satellites are launched directly into the target orbit by a launch vehicle, and the minimum number of launches required is equal to the number of orbital planes of the constellation. After entering orbit, satellites can adjust the phase between adjacent satellites through in-plane maneuvers. However, since the number of satellite launches increases the deployment cost, this method is not suitable for giant constellations.

[0030] (2) Indirect deployment of satellites by rocket. In this method, the satellite is launched directly into the target orbit by a launch vehicle and moves to the target position through a pre-defined control strategy. This method can significantly reduce launch costs. The satellite uses a high-performance electric thruster to move to the target position by taking advantage of the Earth's non-spherical perturbation. In extreme cases, a single launch is sufficient to deploy an entire constellation.

[0031] (3) Rocket-carried satellite joint deployment. This method combines the first two methods, but requires additional steps to find the optimal distribution between launch and on-orbit maneuvering.

[0032] Regarding satellite mobile transfer and deployment, the average orbital motion equations show that satellites can reach a designated location in two ways:

[0033] (1) Direct control. This method directly changes the right ascension of the ascending node and the argument of latitude through small thrust maneuvers.

[0034] (2) Indirect Manipulation. This method alters the semi-major axis and orbital inclination through small-thrust maneuvers, thereby utilizing the Earth's non-spherical J2 perturbation term to change the right ascension of the ascending node and the phase angle. In related technologies, satellites can be deployed sequentially using launch vehicles, with time-sequential adjustments made between parking orbits and target orbits, thus separating the right ascension of the ascending node through the Earth's non-spherical J2 perturbation term. Alternatively, the satellite control strategy can be set to a three-stage approach, with all satellites maneuvering from their initial orbits to "drift" orbits, using the Earth's non-spherical J2 perturbation term to separate the right ascension of the ascending node, and finally, each satellite being moved to its working orbit to complete the constellation's orbital plane deployment. Furthermore, in related technologies, in-plane pulse maneuvers, coplanar eccentric orbits, and multi-objective optimization can be considered when deploying constellation satellites; however, this method involves pulse burn formulas and planar maneuvers, limiting its applicability.

[0035] In summary, existing satellite deployment methods mainly address the problem of separating the right ascension of the ascending node in constellation deployment, without addressing the joint deployment problem of phase angle and right ascension of the ascending node. Therefore, there is an urgent need in the existing technology for a time-series joint deployment method for giant constellations based on low thrust to solve the problem of rapid deployment of constellation satellites with low thrust.

[0036] To address the problems of the prior art, embodiments of this application provide a method, apparatus, electronic device, storage medium, and product for deploying a constellation of satellites. The method for deploying a constellation of satellites provided in this application will be described first.

[0037] Figure 2 A flowchart illustrating a method for deploying a constellation of satellites according to an embodiment of this application is shown. Figure 2 As shown, the method includes the following steps:

[0038] Step S201: Based on the parking position information of the constellation satellites in their current orbits and the target position information of the satellites operating in the target orbits, determine the satellite maneuvering time for the constellation satellites to move from their current orbits to the target orbits.

[0039] In step S201, the parking position of the constellation satellite in the current orbit may include, but is not limited to, the phase angle and right ascension of the ascending node of the constellation satellite in the current orbit; similarly, the target position information of the constellation satellite in the target orbit may include, but is not limited to, the phase angle and right ascension of the ascending node of the constellation satellite in the target orbit.

[0040] Furthermore, in step S201, the satellite maneuver time is the time it takes for the constellation satellites to begin moving from their current orbits.

[0041] As an example, after obtaining the parking position information and target position information of the constellation satellites, a line integral equation can be established based on these information. This line integral equation relates to the orbital inclination and the semi-major axis of the orbit. Using this line integral equation, the change in phase angle of the constellation satellites during the raising of their semi-major axis can be calculated. Based on this change in phase angle, the waiting time required for the constellation satellites to maneuver can be determined, and the maneuvering time of the constellation satellites can be determined accordingly.

[0042] Step S202: Determine the phase adjustment time for the satellite based on the satellite maneuver time and the phase difference between two adjacent constellation satellites in the current orbit.

[0043] In step S202, the phase adjustment time is the time for adjusting the phase of the constellation satellites. During this phase adjustment time, the phase angle of the constellation satellites in their current orbits is adjusted to minimize the angular difference between the phase angle of the constellation satellites in the target orbit and the target phase angle. In this embodiment, the phase adjustment time can be determined based on the phase difference between two adjacent constellation satellites in the same orbit within the constellation.

[0044] Step S203: Determine the right ascension adjustment time for adjusting the right ascension of the ascending node based on the satellite maneuver time and the difference in right ascension of the ascending node.

[0045] In step S203, the ascending node right ascension difference is the difference between the ascending node right ascension of the current orbit and the ascending node right ascension of the target orbit. Similar to the phase adjustment time, adjusting the ascending node right ascension of the constellation satellites during the right ascension adjustment time minimizes the deviation between the ascending node right ascension of the constellation satellites in the target orbit and the target ascending node right ascension.

[0046] Step S204: Determine the orbit inclination compensation value of the current orbit based on the right ascension offset of the ascending node corresponding to the phase adjustment time.

[0047] In step S204, the orbital inclination compensation value is used to compensate for the phase angle and right ascension of the ascending node of the current orbit during the satellite maneuver time. In this embodiment, the orbital inclination compensation value can be determined using the least squares method based on the right ascension offset of the ascending node.

[0048] Step S205: Based on the satellite maneuver time, phase adjustment time, right ascension adjustment time, and orbital inclination compensation value, deploy the constellation satellites in the target orbit.

[0049] After obtaining the orbital inclination compensation value, it is substituted into the formulas for calculating the right ascension offset of the ascending node and the phase offset. The right ascension offset and phase offset are then calculated. During the maneuvering of the constellation satellites, the right ascension of the ascending node and the phase angle of the constellation satellites are compensated using these offsets, respectively, to obtain the compensated right ascension of the ascending node and the compensated phase angle. For example, if the right ascension of the ascending node of a constellation satellite in its current orbit is 0° and the phase angle is also 0°; and the right ascension offset is 1° and the phase offset is 0.5°, then during the constellation satellite maneuvering, the right ascension of the ascending node of the constellation satellite in its current orbit is adjusted to 1° (i.e., the compensated right ascension of the ascending node), and the phase angle of the constellation satellite in its current orbit is adjusted to 0.5° (i.e., the compensated phase angle).

[0050] After determining the satellite maneuver time, phase adjustment time, and right ascension adjustment time, during the satellite maneuver time, the constellation satellites are controlled to maneuver at the positions determined by the compensated phase angle and the compensated right ascension of the ascending node. During the deployment of the constellation satellites, during the phase adjustment time, the compensated phase angle is adjusted to the target phase angle (i.e., the phase angle of the constellation satellites in the target orbit); during the right ascension adjustment time, the compensated right ascension of the ascending node is adjusted to the target right ascension of the ascending node (i.e., the right ascension of the ascending node corresponding to the target orbit), so as to achieve the deployment of the constellation satellites in the target orbit.

[0051] Based on the scheme defined in steps S201 to S205 above, it can be understood that in this embodiment, the orbital inclination compensation value for compensating the orbital inclination of the current orbit is determined according to the offset of the right ascension of the ascending node caused by the phase adjustment. This orbital inclination compensation value is then used to compensate the right ascension of the ascending node and the phase angle of the constellation satellites during their maneuvers, enabling the constellation satellites to accurately reach their target positions on the target orbits. This process only requires compensation of the phase angle and right ascension of the ascending node during the maneuvers of the constellation satellites, thus achieving synchronous adjustment of the right ascension of the ascending node and the phase angle, thereby improving the deployment efficiency of the constellation satellites.

[0052] The methods provided in the embodiments of this application will be explained in detail below.

[0053] In this embodiment of the application, before deploying constellation satellites, it is necessary to determine the orbital inclination compensation value in order to compensate for the phase angle and right ascension of the ascending node during maneuvers.

[0054] Specifically, firstly, based on the correlation between the orbital inclination deviation and the right ascension of the ascending node, a first function is constructed between the orbital inclination compensation value and the right ascension offset of the ascending node for the current orbit; then, based on the correlation between the orbital inclination deviation and the phase angle of the constellation satellites, a second function is constructed between the orbital inclination compensation value and the phase offset; finally, based on the orbital information of the current orbit and the target orbit, the first and second functions are solved simultaneously to obtain the orbital inclination compensation value.

[0055] In the above embodiments, the right ascension offset of the ascending node is the offset of the right ascension of the ascending node of the constellation satellite during the process of adjusting the phase angle of the constellation satellite; the phase offset is the offset of the phase angle of the constellation satellite during the process of adjusting the right ascension of the ascending node; the orbital information (including the orbital information of the current orbit and the orbital information of the target orbit) may include, but is not limited to, the orbital inclination and the semi-major axis of the orbit.

[0056] It should be noted that the deviation of track inclination refers to the angular difference between the track inclination after compensation and the track inclination before compensation. This angular difference is the track inclination compensation value.

[0057] Furthermore, the right ascension of the ascending node is related to the semi-major axis of the orbit, and the phase angle is also related to the semi-major axis of the orbit. Therefore, the right ascension of the ascending node and the phase angle are coupled. Deploying a constellation of satellites refers to placing them in the target orbit at positions determined by the target's right ascension of the ascending node and the target's phase angle. Therefore, in the actual deployment of constellation satellites, the right ascension of the ascending node and the phase angle need to be adjusted multiple times, which reduces the efficiency of satellite deployment.

[0058] In this embodiment, after determining the target ascending node right ascension and target phase angle, the ascending node right ascension and phase angle at the moment of constellation satellite maneuver are compensated by the orbital inclination compensation value. In the subsequent process of raising the semi-major axis, the constellation satellites are deployed directly based on the compensated ascending node right ascension and compensated phase angle, without the need to make multiple adjustments to the ascending node right ascension and phase angle, thereby improving the efficiency of satellite deployment.

[0059] For the right ascension of the ascending node, with the orbital inclination compensation value being Δi, the rate of change deviation of the right ascension of the ascending node can be expressed by formula (1):

[0060]

[0061] Will Expanding according to Taylor's formula, the above formula (1) can be simplified to formula (2):

[0062]

[0063] Integrating equation (2), we obtain equation (3):

[0064]

[0065] After simplification, we obtain formula (4), which is the first function:

[0066]

[0067] Among them, Ω drift Δi is the right ascension offset of the ascending node, representing the impact of the orbital compensation value on the right ascension of the ascending node during the adjustment of the semi-major axis of the orbit; μ is the Earth's gravitational constant; m is the mass of the constellation satellite; F is the thrust of the constellation satellite's propulsion system; a0 is the length of the semi-major axis of the current orbit; a f J1 is the semi-major axis length of the target orbit; J2 is the Earth's non-spherical perturbation term; R is the Earth's average radius; i is the orbital inclination of the current orbit.

[0068] By rearranging formula (3), we can obtain formula (5), which is the second function:

[0069]

[0070] Among them, u drift The phase offset represents the effect of the track compensation value on the phase during the track semi-major axis adjustment process.

[0071] As an example, after constructing the first and second functions, a matrix can be built based on the coefficients of the first and second functions. This matrix can then be used to establish an equation relating the right ascension of the ascending node and the phase angle to the orbital inclination compensation value and the time interval between constellation satellite maneuvers. Solving this equation yields the orbital inclination compensation value. Here, the time interval between the aforementioned constellation satellite maneuvers is the difference between the maneuver times of two satellites.

[0072] After constructing the first function and the second function using formulas (4) and (5) respectively, an equation is constructed based on the coefficients of the first function and the second function to determine the track inclination compensation value.

[0073] Specifically, firstly, based on the orbital information of the current orbit and the target orbit, a third function relating the first duration of the constellation satellites to the orbital inclination is constructed. Then, based on the first, second, and third functions, the least squares method is used to construct the objective equation corresponding to the orbital inclination compensation value. Solving the objective equation yields the orbital inclination compensation value.

[0074] In the above embodiments, the orbit information of the current orbit includes at least the orbit inclination of the current orbit and the semi-major axis length of the current orbit, the orbit information of the target orbit includes at least the semi-major axis length of the target orbit, and the first duration is the duration between the current time and the satellite maneuver time of the constellation satellite.

[0075] As an example, a third function can be constructed in the following two steps:

[0076] Step 1: Construct a line integral equation based on the parking position information of the constellation satellites and the target position information. The line integral equation can be expressed by formula (6):

[0077]

[0078] In formula (6), u c The current phase angle of the current orbit; u tar The target phase angle of the target orbit; t is the rate of change of the phase angle of the current orbit; t0 is the time before the constellation satellites maneuver; t f T represents the time following the satellite maneuver of the constellation. f The first duration is the difference between the current time and the maneuvering time of the constellation satellites; This represents the rate of change of phase angle of the constellation satellites during the semi-major axis lift.

[0079] In formula (6), and It can be expressed by formulas (7) and (8):

[0080]

[0081]

[0082] In formula (7), n0 is the angular velocity of the constellation satellite in its current orbit; a0 is the semi-major axis of the current orbit; in formula (8), n is the angular velocity of the constellation satellite during the semi-major axis lift; a is the semi-major axis of the orbit in which the constellation satellite is located during the semi-major axis lift.

[0083] Step 2: Derive the rate of change of phase angle of constellation satellites during semi-major axis lifting.

[0084] The phase angle change of the constellation satellites during the raising of their semi-major axis is calculated using the curve integral method, as shown in formula (9):

[0085]

[0086] Simplifying formula (9), we get formula (10):

[0087]

[0088] Substituting formula (10) into formula (6) will yield the first duration T. f That is, the third function can be represented by formula (11):

[0089]

[0090] Furthermore, after constructing the first function, the second function, and the third function, the objective equation can be constructed based on the coefficients of the first function, the second function, and the third function. The objective equation can be represented by formula (12):

[0091]

[0092] In formula (12), Δt is the time interval for constellation satellite maneuvers; Δi is the orbital inclination compensation value;

[0093]

[0094]

[0095] Among them, Ω man The change in right ascension of the ascending node during the process of raising the semi-major axis of the constellation satellites; Ω c Ω represents the right ascension of the ascending node corresponding to the current orbit. tar The right ascension of the ascending node corresponding to the target orbit; This represents the rate of change of right ascension at the ascending node corresponding to the current orbit. ΔΩ represents the rate of change of right ascension at the ascending node corresponding to the target orbit. req The right ascension interval of the ascending node is preset.

[0096] Solving formula (12) yields the maneuver time interval and orbital inclination compensation value for the right ascension and phase angle of the ascending node of each constellation satellite joint control, as shown in formula (13):

[0097]

[0098] After calculating the orbital inclination compensation value, substituting this value into formulas (5) and (6) will compensate for the right ascension of the ascending node and the phase angle, resulting in the compensated right ascension of the ascending node and the compensated phase angle. Then, based on the satellite maneuver time, the compensated phase angle, and the compensated right ascension of the ascending node, a constellation of satellites can be deployed in the target orbit.

[0099] After obtaining the orbital inclination compensation value, the phase angle and right ascension of the ascending node of the constellation satellites can be compensated based on the orbital inclination compensation value during the satellite maneuver time, so as to obtain the compensated phase angle and the compensated right ascension of the ascending node.

[0100] After obtaining the orbital inclination compensation value, substitute the orbital inclination compensation value into the above formulas (4) and (5) to obtain the ascending node right ascension offset and phase offset. Then, at the maneuvering time of the constellation satellite, use the ascending node right ascension offset and phase offset to compensate for the ascending node right ascension and phase angle of the constellation satellite, respectively, to obtain the compensated ascending node right ascension and compensated phase angle. For example, if the ascending node right ascension of the constellation satellite in the current orbit is 0° and the phase angle is also 0°; the ascending node right ascension offset is 1° and the phase offset is 0.5°, then at the maneuvering time of the constellation satellite, adjust the ascending node right ascension of the constellation satellite in the current orbit to 1° (i.e., the compensated ascending node right ascension) and adjust the phase angle of the constellation satellite in the current orbit to 0.5° (i.e., the compensated phase angle).

[0101] After determining the satellite maneuver time, the compensated phase angle, and the compensated right ascension of the ascending node, the compensated phase angle is adjusted to the target phase angle of the target orbit during the phase adjustment time, and the compensated right ascension of the ascending node is adjusted to the target right ascension of the ascending node of the target orbit during the right ascension adjustment time, so that the constellation satellites move to the target position on the target orbit, thereby realizing the deployment of the constellation satellites on the target orbit.

[0102] In the above embodiments, the difference in right ascension of the ascending node is the difference between the right ascension of the ascending node of the current orbit and the right ascension of the ascending node of the target orbit.

[0103] Phase adjustment time can be calculated based on the phase difference between adjacent satellites in the same orbit within the constellation. Specifically, first, the phase angle change of the target orbit and the phase difference between the constellation satellites in the target orbit are obtained; then, based on the semi-major axis length of the current orbit, the semi-major axis length of the target orbit, and the thrust of the constellation satellites' propulsion systems, the second duration for raising the semi-major axis of the constellation satellites is calculated; next, based on the phase angle change rate of the target orbit, the corresponding phase difference of the target orbit, the second duration, and the phase angle change, the phase angle separation duration is determined; finally, the phase adjustment time can be determined based on the satellite maneuver time and the phase angle separation duration.

[0104] In the above embodiments, the phase angle separation duration is the difference between the satellite maneuver time and the time for adjusting the phase angle of the constellation satellites. As can be seen from the above embodiments, before calculating the phase angle separation duration, it is first necessary to calculate the second duration for raising the semi-major axis of the constellation satellites based on the satellites' semi-major axis elevation and the thrust of the electric thrusters. The second duration can be expressed by formula (14):

[0105]

[0106] In formula (14), T man This is the second duration; other parameters have been described above and will not be repeated here.

[0107] After determining the second duration, the phase angle separation duration is determined based on the phase difference between satellites in the same orbit within the constellation. The phase angle separation duration can be expressed by formula (15):

[0108]

[0109] In formula (15), Phase angle separation time; The phase angle change of the target orbit; u req The preset phase interval is a set value, which varies depending on the different phases of constellation satellite deployment. req It can be the phase interval within the same orbit, or the phase interval between adjacent satellites in different orbits; T man This is the second duration. and Equations (16) and (17) represent the following respectively:

[0110]

[0111]

[0112] Substituting formulas (16), (17), (14), and (10) into formula (15) yields the final expression for the phase angle separation time:

[0113]

[0114] After calculating the phase angle separation time, the sum of the satellite maneuver time and the phase angle separation time is calculated, and the resulting time is the phase adjustment time.

[0115] To determine the right ascension adjustment time, first, obtain the right ascension of the current ascending node corresponding to the current orbit and the right ascension of the target ascending node corresponding to the target orbit. Then, based on the semi-major axis length of the current orbit, the orbital inclination of the current orbit, the semi-major axis length of the target orbit, and the thrust of the constellation satellites' propulsion system, calculate the change in the right ascension of the ascending node during the process of raising the semi-major axis of the constellation satellites. Next, based on the change in the right ascension of the ascending node, the second duration, and the difference in the right ascension of the ascending nodes, determine the separation duration of the right ascension of the ascending nodes. Finally, the right ascension adjustment time can be determined based on the satellite maneuver time and the separation duration of the right ascension of the ascending nodes.

[0116] In the above embodiments, the separation time of the right ascension of the ascending node is the difference between the satellite maneuver time and the time for adjusting the right ascension of the ascending node. As can be seen from the above embodiments, before calculating the separation time of the right ascension of the ascending node, it is first necessary to calculate the change in the right ascension of the ascending node during the process of raising the semi-major axis of the constellation satellites. The change in the right ascension of the ascending node can be expressed by formula (19):

[0117]

[0118] In formula (18), Ω man This represents the change in right ascension at the ascending node.

[0119] The duration of separation of the right ascension of the ascending node can be expressed by formula (19):

[0120]

[0121] In formula (19), The duration of separation of the right ascension at the ascending node. and It can be expressed by formula (20) and formula (21) respectively:

[0122]

[0123]

[0124] Substituting formulas (20), (21), and (14) into formula (19), we can obtain the final expression for the duration of separation of the right ascension of the ascending node:

[0125]

[0126] The following examples illustrate the method provided in this application. This process can be carried out by... Figure 3To indicate, by Figure 3 It can be seen that the deployment of constellation satellites mainly includes the following five steps:

[0127] Step S301: Determine the satellite maneuvering time based on the docking positions of the constellation satellites and the target position;

[0128] Step S302: Determine the phase adjustment time based on the phase difference between adjacent satellites in the same orbit within the constellation;

[0129] Step S303: Determine the right ascension adjustment time based on the difference in right ascension of the ascending nodes of satellites in different orbits within the constellation;

[0130] Step S304: Based on the right ascension drift of the ascending node during the phase adjustment process, determine the orbital inclination compensation value and the time interval of constellation satellite maneuvers using the least squares method;

[0131] Step S305: Deploy the entire constellation based on the maneuver time, phase adjustment time, right ascension adjustment time, and orbital inclination compensation value of all satellites in the constellation.

[0132] The method provided in this application embodiment is simulated, and the constellation satellites are set as OneWeb constellation satellites. The OneWeb constellation configuration parameter table is shown in Table 1. Through simulation calculation, the following can be obtained: Figure 4 The diagram shows the timing of the semi-major axis of the satellites' lift orbits. Figure 4 In the diagram, lines of different colors represent different constellation satellites.

[0133] Table 1 OneWeb Constellation Configuration Parameters

[0134] parameter symbol numerical values unit Total number of satellites N 720 Number of orbital planes P 18 Number of orbital satellites S 40 track inclination i 87.9 deg orbital height h 1200 km Ascending node right ascension interval ΔΩ 10.2 deg Phase spacing within the same orbital plane <![CDATA[Δu intra ]]> 9 deg Phase spacing between adjacent satellites on different orbital planes <![CDATA[Δu inter ]]> 4.5 deg Thrust of electric propulsion system F 154 mN Specific impulse of electric propulsion system <![CDATA[I sp ]]> 2035 s

[0135] For step S301, the initial orbital altitude of the satellite is set to 500 km, the initial phase angle to 180 degrees, and the target phase angle to 0 degrees. Substituting into formula (11), the maneuver time of the constellation satellites is 227 seconds later.

[0136] For step S302, the initial orbital altitude of the satellite is 500 km, the orbital inclination is 87.9 degrees, the target orbital altitude is 1200 km, and the phase angle interval between adjacent satellites on the same orbital plane is 9 degrees. The phase modulation time interval can be calculated as 1052.023 seconds according to formula (18).

[0137] For step S303, the satellite's initial orbital altitude is 500 km, the orbital inclination is 87.9 degrees, the target orbital altitude is 1200 km, and the right ascension interval of the ascending nodes on different orbital planes is 10.2 degrees. The time interval for the right ascension modulation of the ascending nodes can be calculated as 10535158.446 seconds according to formula (22).

[0138] For step S304, the initial orbital altitude of the satellite is 500 km, the orbital inclination is 87.9 degrees, the target orbital altitude is 1200 km, the phase angle interval between adjacent satellites on the same orbital plane is 9 degrees, and the phase angle interval between adjacent satellites on different orbital planes is 4.5 degrees. The orbital inclination compensation value can be calculated as 0.002 degrees according to formula (12).

[0139] For step S305, the entire constellation is deployed based on the maneuver duration of all satellites in the constellation, the phase adjustment timing, the ascending node right ascension adjustment timing, and the preset orbital inclination. The orbital semi-major axis variation is as follows: Figure 4 As shown, the timing of phase angle changes is as follows: Figure 5 As shown, the satellite position changes during the phase deployment process are as follows: Figure 6 As shown, the temporal sequence of the changes in the right ascension interval of the ascending node is as follows: Figure 7 As shown, the phase angles and right ascension of the ascending nodes of 80 satellites in the two orbital planes are as follows: Figure 8 As shown.

[0140] In summary, the method provided in this application can effectively achieve synchronous adjustment of the right ascension of the ascending node and the phase angle of satellites in multi-satellite launch missions of large constellations by utilizing the continuous small thrust configured on the satellites. First, this application proposes a timing control strategy that utilizes the influence of the semi-major axis on the right ascension of the ascending node and the phase angle to ensure that the injected satellites can be sequentially transferred from the initial orbit to the target orbit. Based on this, the coupling relationship between the right ascension of the ascending node and the phase angle during maneuvers is studied, and a mathematical model (e.g., Equation 12) is established to describe the drift of the right ascension of the ascending node, the drift of the phase angle, the orbital inclination error, and the maneuver time interval. The mathematical model is precisely calculated using the least squares method to obtain the optimal maneuver time interval and the orbital inclination compensation value. Finally, by combining the timing control and inclination compensation strategies, high-precision synchronous deployment of constellation satellites is achieved, demonstrating the application potential of small thrust in large-scale orbit control.

[0141] This application also provides a constellation satellite deployment device, such as... Figure 9 As shown, the device 900 includes: a time determination module 901, a first determination module 902, a second determination module 903, a third determination module 904, and a satellite deployment module 905.

[0142] The time determination module 901 is used to determine the satellite maneuvering time for the constellation satellites to move from the current orbit to the target orbit based on the parking position information of the constellation satellites in the current orbit and the target position information of the satellites running in the target orbit.

[0143] The first determining module 902 is used to determine the phase adjustment time for adjusting the phase of the satellite based on the satellite maneuver time and the phase difference between two adjacent constellation satellites in the current orbit.

[0144] The second determining module 903 is used to determine the right ascension adjustment time for adjusting the right ascension of the ascending node based on the satellite maneuver time and the difference in right ascension of the ascending node, wherein the difference in right ascension of the ascending node is the difference between the right ascension of the ascending node of the current orbit and the right ascension of the ascending node of the target orbit.

[0145] The third determining module 904 is used to determine the orbital inclination compensation value of the current orbit based on the right ascension offset of the ascending node corresponding to the phase adjustment time. The orbital inclination compensation value is used to compensate for the phase angle and right ascension of the ascending node of the current orbit during the satellite maneuver time.

[0146] Satellite deployment module 905 is used to deploy constellation satellites in a target orbit based on satellite maneuver time, phase adjustment time, right ascension adjustment time, and orbital inclination compensation value.

[0147] In one example, the third determining module specifically includes: a first building module, a second building module, and a solving module. The first building module constructs a first function relating the orbital inclination compensation value and the right ascension offset of the ascending node, based on the correlation between the orbital inclination deviation and the right ascension of the ascending node. The right ascension offset is the offset of the ascending node right ascension of the constellation satellites during the adjustment of their phase angles. The second building module constructs a second function relating the orbital inclination compensation value and the phase offset, based on the correlation between the orbital inclination deviation and the phase angles of the constellation satellites. The phase offset is the offset of the phase angles of the constellation satellites during the adjustment of their ascending node right ascension. The solving module solves the first and second functions simultaneously based on the orbital information of the current orbit and the target orbit to obtain the orbital inclination compensation value.

[0148] In one example, the first function is represented by the following formula:

[0149]

[0150] Among them, Ω drift Δi is the right ascension offset of the ascending node; μ is the orbital inclination compensation value; m is the mass of the constellation satellite; F is the thrust of the constellation satellite's propulsion system; a0 is the semi-major axis length of the current orbit; a f J1 is the semi-major axis length of the target orbit; J2 is the Earth's non-spherical perturbation term; R is the Earth's average radius; i is the orbital inclination of the current orbit.

[0151] In one example, the second function is represented by the following formula:

[0152]

[0153] Among them, u drift This represents the phase offset.

[0154] In one example, the solution module is specifically used to construct a third function relating the first duration of the constellation satellites to the orbital inclination, based on the orbital information of the current orbit and the target orbit. The first duration is the duration between the current time and the satellite maneuver time. Based on the first function, the second function, and the third function, the least squares method is used to construct the objective equation corresponding to the orbital inclination compensation value. The objective equation is solved to obtain the orbital inclination compensation value.

[0155] In one example, the orbit information of the current orbit includes at least the orbit inclination angle and the semi-major axis length of the current orbit, and the orbit information of the target orbit includes at least the semi-major axis length of the target orbit, wherein the third function is represented by the following formula:

[0156]

[0157] Among them, T f The first duration; u tar The target phase angle of the target orbit; u c This represents the current phase angle of the current orbit. This represents the rate of change of the phase angle of the current orbit.

[0158] In one example, the objective equation is expressed as follows:

[0159]

[0160] Where Δt is the time interval for constellation satellite maneuvers; Δi is the orbital inclination compensation value;

[0161]

[0162]

[0163] Among them, Ω man The change in right ascension of the ascending node during the process of raising the semi-major axis of the constellation satellites; Ω c Ω represents the right ascension of the ascending node corresponding to the current orbit. tar The right ascension of the ascending node corresponding to the target orbit; This represents the rate of change of right ascension at the ascending node corresponding to the current orbit. ΔΩ represents the rate of change of right ascension at the ascending node corresponding to the target orbit. req The right ascension interval of the ascending node is preset.

[0164] In one example, the first determining module is specifically used to obtain the phase angle change of the target orbit and the phase difference of the constellation satellites in the target orbit; calculate the second duration for the constellation satellites to raise the semi-major axis based on the semi-major axis length of the current orbit, the semi-major axis length of the target orbit, and the thrust of the constellation satellites' propulsion system; determine the phase angle separation duration based on the phase angle change rate of the target orbit, the phase difference corresponding to the target orbit, the second duration, and the phase angle change, wherein the phase angle separation duration is the difference between the satellite maneuver time and the time for adjusting the phase angle of the constellation satellites; and determine the phase adjustment time based on the satellite maneuver time and the phase angle separation duration.

[0165] In one example, the phase angle separation time is determined by the following formula:

[0166]

[0167] in, Phase angle separation time; The phase angle change of the target orbit; u req The preset phase interval; T man This is the second duration.

[0168] In one example, the second determining module is specifically used to obtain the right ascension of the current ascending node corresponding to the current orbit and the right ascension of the target ascending node corresponding to the target orbit; calculate the change in the right ascension of the ascending node during the process of raising the semi-major axis of the constellation satellites based on the semi-major axis length of the current orbit, the orbital inclination of the current orbit, the semi-major axis length of the target orbit, and the thrust of the constellation satellites' propulsion system; determine the separation duration of the right ascension of the ascending node based on the change in the right ascension of the ascending node, the second duration, and the difference in the right ascension of the ascending node, wherein the separation duration of the right ascension of the ascending node is the difference between the satellite maneuver time and the time for adjusting the right ascension of the ascending node; and determine the right ascension adjustment time based on the satellite maneuver time and the separation duration of the right ascension of the ascending node.

[0169] In one example, the duration of separation of the right ascension of the ascending node is determined by the following formula:

[0170]

[0171] in, The duration of separation of the right ascension at the ascending node.

[0172] In one example, the satellite deployment module is specifically used to compensate the phase angle and right ascension of the ascending node of the constellation satellites during the satellite maneuvering time based on the orbital inclination compensation value, so as to obtain the compensated phase angle and the compensated right ascension of the ascending node; during the phase adjustment time, the compensated phase angle is adjusted to the target phase angle of the target orbit, and during the right ascension adjustment time, the compensated right ascension of the ascending node is adjusted to the target right ascension of the ascending node of the target orbit, so as to move the constellation satellites to the target position on the target orbit.

[0173] The constellation satellite deployment device provided in this application embodiment can realize the various processes implemented in the aforementioned method embodiments, and will not be repeated here to avoid repetition.

[0174] Those skilled in the art will clearly understand that, for the sake of convenience and brevity, the above-described division of functional units and modules is merely an example. In practical applications, the above functions can be assigned to different functional units and modules as needed, that is, the internal structure of the device can be divided into different functional units or modules to complete all or part of the functions described above. The functional units and modules in the embodiments can be integrated into one processing unit, or each unit can exist physically separately, or two or more units can be integrated into one unit. The integrated unit can be implemented in hardware or as a software functional unit. Furthermore, the specific names of the functional units and modules are only for easy differentiation and are not intended to limit the scope of protection of this application. The specific working process of the units and modules in the above system can be referred to the corresponding process in the foregoing method embodiments, and will not be repeated here.

[0175] Figure 10 A schematic diagram of the hardware structure of the electronic device provided in an embodiment of this application is shown.

[0176] The electronic device may include a processor 1001 and a memory 1002 storing computer program instructions.

[0177] Specifically, the processor 1001 may include a central processing unit (CPU), an application-specific integrated circuit (ASIC), or one or more integrated circuits that can be configured to implement the embodiments of this application.

[0178] Memory 1002 may include mass storage for data or instructions. For example, and not limitingly, memory 1002 may include a hard disk drive (HDD), floppy disk drive, flash memory, optical disk, magneto-optical disk, magnetic tape, or Universal Serial Bus (USB) drive, or a combination of two or more of these. Where appropriate, memory 1002 may include removable or non-removable (or fixed) media. Where appropriate, memory 1002 may be internal or external to the integrated gateway disaster recovery device. In a particular embodiment, memory 1002 is non-volatile solid-state memory.

[0179] Memory may include read-only memory (ROM), random access memory (RAM), disk storage media devices, optical storage media devices, flash memory devices, and electrical, optical, or other physical / tangible memory storage devices. Therefore, typically, memory includes one or more tangible (non-transitory) computer-readable storage media (e.g., memory devices) encoded with software including computer-executable instructions, and when the software is executed (e.g., by one or more processors), it is operable to perform the operations described with reference to the methods according to one aspect of this disclosure.

[0180] The processor 1001 implements any of the constellation satellite deployment methods in the above embodiments by reading and executing computer program instructions stored in the memory 1002.

[0181] In one example, the electronic device may also include a communication interface 1003 and a bus 1010. For example, Figure 10 As shown, the processor 1001, memory 1002, and communication interface 1003 are connected through bus 1010 and complete communication with each other.

[0182] The communication interface 1003 is mainly used to realize communication between various modules, devices, units and / or equipment in the embodiments of this application.

[0183] Bus 1010 includes hardware, software, or both, that couples components of an electronic device together. For example, and not limitingly, the bus may include an Accelerated Graphics Port (AGP) or other graphics bus, an Enhanced Industry Standard Architecture (EISA) bus, a Front Side Bus (FSB), HyperTransport (HT) interconnect, an Industry Standard Architecture (ISA) bus, an Infinite Bandwidth Interconnect, a Low Pin Count (LPC) bus, a memory bus, a Microchannel Architecture (MCA) bus, a Peripheral Component Interconnect (PCI) bus, a PCI-Express (PCI-X) bus, a Serial Advanced Technology Attachment (SATA) bus, a Video Electronics Standards Association Local (VLB) bus, or other suitable buses, or combinations of two or more of these. Where appropriate, bus 1010 may include one or more buses. Although specific buses are described and illustrated in embodiments of this application, any suitable bus or interconnect is contemplated herein.

[0184] Furthermore, in conjunction with the satellite constellation deployment methods in the above embodiments, this application embodiment can provide a computer-readable storage medium for implementation. This computer-readable storage medium stores computer program instructions; when executed by a processor, these computer program instructions implement any of the satellite constellation deployment methods in the above embodiments.

[0185] Furthermore, in conjunction with the satellite constellation deployment methods described in the above embodiments, this application can provide a computer program product for implementation. When the instructions in this computer program product are executed by the processor of an electronic device, the electronic device performs the satellite constellation deployment method as described in any of the above embodiments.

[0186] It should be clarified that this application is not limited to the specific configurations and processes described above and shown in the figures. For the sake of brevity, detailed descriptions of known methods are omitted here. In the above embodiments, several specific steps are described and shown as examples. However, the method process of this application is not limited to the specific steps described and shown. Those skilled in the art can make various changes, modifications, and additions, or change the order of steps, after understanding the spirit of this application.

[0187] The functional modules shown in the above-described block diagram can be implemented as hardware, software, firmware, or a combination thereof. When implemented in hardware, they can be, for example, electronic circuits, application-specific integrated circuits (ASICs), appropriate firmware, plug-ins, function cards, etc. When implemented in software, the elements of this application are programs or code segments used to perform the required tasks. Programs or code segments can be stored on a machine-readable medium or transmitted over a transmission medium or communication link via data signals carried on a carrier wave. "Machine-readable medium" can include any medium capable of storing or transmitting information. Examples of machine-readable media include electronic circuits, semiconductor memory devices, ROM, flash memory, erasable ROM (EROM), floppy disks, CD-ROMs, optical disks, hard disks, fiber optic media, radio frequency (RF) links, etc. Code segments can be downloaded via computer networks such as the Internet, intranets, etc.

[0188] It should also be noted that the exemplary embodiments mentioned in this application describe methods or systems based on a series of steps or apparatus. However, this application is not limited to the order of the above steps; that is, the steps can be performed in the order mentioned in the embodiments, or in a different order, or several steps can be performed simultaneously.

[0189] The above flowcharts and / or block diagrams describing methods, apparatuses, electronic devices, storage media, and products for deploying constellation satellites according to embodiments of this disclosure have described various aspects of the present disclosure. It should be understood that each block in the flowcharts and / or block diagrams, and combinations of blocks in the flowcharts and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, or other programmable data processing apparatus to create a machine such that these instructions, executable via the processor of the computer or other programmable data processing apparatus, enable the implementation of the functions / actions specified in one or more blocks of the flowcharts and / or block diagrams. Such a processor can be, but is not limited to, a general-purpose processor, a special-purpose processor, a special application processor, or a field-programmable logic circuit. It is also understood that each block in the block diagrams and / or flowcharts, and combinations of blocks in the block diagrams and / or flowcharts, can also be implemented by dedicated hardware performing the specified functions or actions, or can be implemented by a combination of dedicated hardware and computer instructions.

[0190] The above description is merely a specific implementation of this application. Those skilled in the art will clearly understand that, for the sake of convenience and brevity, the specific working processes of the systems, modules, and units described above can be referred to the corresponding processes in the foregoing method embodiments, and will not be repeated here. It should be understood that the protection scope of this application is not limited thereto. Any person skilled in the art can easily conceive of various equivalent modifications or substitutions within the technical scope disclosed in this application, and these modifications or substitutions should all be covered within the protection scope of this application.

Claims

1. A method for deploying a constellation of satellites, characterized in that, include: Based on the parking position information of the constellation satellites in their current orbits and the target position information of the satellites operating in the target orbits, the satellite maneuvering time for the constellation satellites to move from their current orbits to the target orbits is determined. The phase adjustment time for adjusting the phase of the satellite is determined based on the satellite maneuver time and the phase difference between two adjacent constellation satellites in the current orbit. Based on the satellite maneuver time and the difference in right ascension of the ascending node, the right ascension adjustment time for adjusting the right ascension of the ascending node is determined, wherein the difference in right ascension of the ascending node is the difference between the right ascension of the ascending node of the current orbit and the right ascension of the ascending node of the target orbit; Based on the right ascension offset of the ascending node corresponding to the phase adjustment time, the orbital inclination compensation value of the current orbit is determined, wherein the orbital inclination compensation value is used to compensate for the phase angle and right ascension of the ascending node of the current orbit during the satellite maneuver time; Based on the satellite maneuver time, the phase adjustment time, the right ascension adjustment time, and the orbital inclination compensation value, the constellation satellites are deployed in the target orbit.

2. The method according to claim 1, characterized in that, Based on the right ascension shift of the ascending node corresponding to the phase adjustment time, the orbital inclination compensation value of the current orbit is determined, including: Based on the correlation between the deviation of the orbital inclination angle and the right ascension of the ascending node, a first function is constructed between the orbital inclination compensation value of the current orbit and the right ascension offset of the ascending node, wherein the right ascension offset of the ascending node is the offset of the right ascension of the ascending node corresponding to the constellation satellite during the process of adjusting the phase angle of the constellation satellite; Based on the correlation between the deviation of the orbital inclination angle and the phase angle of the constellation satellites, a second function is constructed between the orbital inclination angle compensation value and the phase offset, wherein the phase offset is the offset of the phase angle of the constellation satellites during the process of adjusting the right ascension of the ascending node; Based on the track information of the current track and the track information of the target track, the first function and the second function are solved simultaneously to obtain the track inclination compensation value.

3. The method according to claim 2, characterized in that, The first function is represented by the following formula: Among them, Ω drift Δi is the right ascension offset of the ascending node; μ is the orbital inclination compensation value; m is the mass of the constellation satellite; F is the thrust of the constellation satellite's propulsion system; a0 is the semi-major axis length of the current orbit; a f J1 is the semi-major axis length of the target orbit; J2 is the Earth's non-spherical perturbation term; R is the Earth's average radius; and i is the orbital inclination angle of the current orbit.

4. The method according to claim 3, characterized in that, The second function is represented by the following formula: Among them, u drift This refers to the phase offset.

5. The method according to claim 4, characterized in that, Based on the orbit information of the current orbit and the orbit information of the target orbit, the first function and the second function are solved simultaneously to obtain the orbit inclination compensation value, including: Based on the orbit information of the current orbit and the orbit information of the target orbit, a third function is constructed between the first duration of the constellation satellite and the orbital inclination, wherein the first duration is the duration of the constellation satellite from the current time to the satellite maneuver time; Based on the first function, the second function, and the third function, the objective equation corresponding to the track inclination compensation value is constructed using the least squares method. Solve the objective equation to obtain the track inclination compensation value.

6. The method according to claim 5, characterized in that, The orbit information of the current orbit includes at least the orbit inclination angle and the semi-major axis length of the current orbit, and the orbit information of the target orbit includes at least the semi-major axis length of the target orbit, wherein the third function is represented by the following formula: Among them, T f The first duration; u tar The target phase angle of the target orbit; u c The current phase angle of the current orbit; The phase angle change rate of the current orbit.

7. The method according to claim 6, characterized in that, The objective equation is expressed by the following formula: Where Δt is the time interval for constellation satellite maneuvers; Δi is the orbital inclination compensation value; Among them, Ω man Ω represents the change in right ascension of the ascending node during the process of raising the semi-major axis of the satellites in the constellation; c Ω represents the right ascension of the ascending node corresponding to the current orbit. tar The right ascension of the ascending node corresponding to the target orbit; The rate of change of right ascension of the ascending node corresponding to the current orbit; ΔΩ is the rate of change of right ascension of the ascending node corresponding to the target orbit. req The right ascension interval of the ascending node is preset.

8. The method according to claim 7, characterized in that, Based on the satellite maneuver time and the phase difference between two adjacent constellation satellites in the current orbit, the phase adjustment time for adjusting the phase of the satellite is determined, including: The phase angle change of the target orbit and the phase difference of the constellation satellites in the target orbit are obtained; The second duration for raising the semi-major axis of the constellation satellites is calculated based on the semi-major axis length of the current orbit, the semi-major axis length of the target orbit, and the thrust of the satellites' propulsion system. Based on the phase angle change rate of the target orbit, the phase difference corresponding to the target orbit, the second duration, and the phase angle change amount, the phase angle separation duration is determined, wherein the phase angle separation duration is the difference between the satellite maneuver time and the time for adjusting the phase angle of the constellation satellites; The phase adjustment time is determined based on the satellite maneuver time and the phase angle separation duration.

9. The method according to claim 8, characterized in that, The phase angle separation time is determined by the following formula: in, The phase angle separation time; The change in phase angle of the target orbit; u req The preset phase interval; T man This is the second duration.

10. The method according to claim 9, characterized in that, Based on the satellite maneuver time and the difference in right ascension of the ascending node, the right ascension adjustment time for adjusting the right ascension of the ascending node is determined, including: Obtain the right ascension of the current ascending node corresponding to the current orbit, and the right ascension of the target ascending node corresponding to the target orbit; Based on the semi-major axis length of the current orbit, the orbital inclination angle of the current orbit, the semi-major axis length of the target orbit, and the thrust of the propulsion system of the constellation satellite, calculate the change in right ascension of the ascending node during the process of raising the semi-major axis of the constellation satellite; The separation duration of the right ascension of the ascending node is determined based on the change in the right ascension of the ascending node, the second duration, and the difference in the right ascension of the ascending node, wherein the separation duration of the right ascension of the ascending node is the difference between the satellite maneuver time and the time for adjusting the right ascension of the ascending node; The right ascension adjustment time is determined based on the satellite maneuver time and the right ascension separation time of the ascending node.

11. The method according to claim 10, characterized in that, The duration of separation of the right ascension of the ascending node is determined by the following formula: in, The time of separation of the right ascension of the ascending node.

12. The method according to claim 2, characterized in that, Based on the satellite maneuver time, the phase adjustment time, the right ascension adjustment time, and the orbital inclination compensation value, the constellation satellites are deployed in the target orbit, including: Based on the orbital inclination compensation value, the phase angle and right ascension of the ascending node of the constellation satellites are compensated during the satellite maneuver time to obtain the compensated phase angle and the compensated right ascension of the ascending node; During the phase adjustment time, the compensated phase angle is adjusted to the target phase angle of the target orbit, and during the right ascension adjustment time, the compensated ascending node right ascension is adjusted to the target ascending node right ascension of the target orbit, so that the constellation satellites move to the target position on the target orbit.

13. A deployment device for a constellation of satellites, characterized in that, include: The time determination module is used to determine the satellite maneuvering time for the constellation satellites to move from the current orbit to the target orbit based on the parking position information of the constellation satellites in the current orbit and the target position information of the satellites running in the target orbit. The first determining module is used to determine the phase adjustment time for adjusting the phase of the satellite based on the satellite maneuver time and the phase difference between two adjacent constellation satellites in the current orbit; The second determining module is used to determine the right ascension adjustment time for adjusting the right ascension of the ascending node based on the satellite maneuver time and the difference in right ascension of the ascending node, wherein the difference in right ascension of the ascending node is the difference between the right ascension of the ascending node of the current orbit and the right ascension of the ascending node of the target orbit; The third determining module is used to determine the orbital inclination compensation value of the current orbit based on the right ascension offset of the ascending node corresponding to the phase adjustment time, wherein the orbital inclination compensation value is used to compensate for the phase angle and right ascension of the ascending node of the current orbit during the satellite maneuver time; The satellite deployment module is used to deploy the constellation satellites in the target orbit based on the satellite maneuvering time, the phase adjustment time, the right ascension adjustment time, and the orbital inclination compensation value.

14. An electronic device, characterized in that, Electronic devices include: processors and memory storing computer program instructions; When the processor executes the computer program instructions, it implements the method for deploying constellation satellites as described in any one of claims 1-12.

15. A computer-readable storage medium, characterized in that, A computer-readable storage medium stores computer program instructions that, when executed by a processor, implement the deployment method of the constellation satellites as described in any one of claims 1-12.

16. A computer program product, characterized in that, When the instructions in the computer program product are executed by the processor of an electronic device, the electronic device causes the electronic device to perform the deployment method of the constellation satellites as described in any one of claims 1-12.

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