Spacecraft rendezvous and docking method and device based on step-by-step orbit-raising phase control

Abstract

The application discloses a spacecraft rendezvous and docking method and device based on step-by-step orbit raising phase modulation control, and relates to the technical field of spaceflight. The spacecraft rendezvous and docking method comprises the following steps: determining the coplanar phase difference between a target spacecraft and a visiting spacecraft corresponding to each launch window based on a preset launch window range; determining a phase error reference based on a determination date; determining a target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference and a target phase difference; and launching the visiting spacecraft in the target launch window based on a step-by-step orbit raising phase modulation control strategy. The application solves the technical problem that the launch window cannot be accurately determined in the spacecraft rendezvous and docking task in the related art, resulting in low phase modulation control efficiency.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and more specifically, to a spacecraft rendezvous and docking method and apparatus based on step-by-step orbital ascent and phasing control. Background Technology

[0002] During rendezvous and docking missions, due to the limitations of the rocket's carrying capacity, the visiting spacecraft, after separating from the rocket, is in a lower orbit relative to the target spacecraft. Operating in this lower orbit, the visiting spacecraft's orbital angular velocity is much greater than that of the target spacecraft. During long-range guidance after entering orbit, the visiting spacecraft gradually approaches the target spacecraft, entering a pursuit process.

[0003] For different long-range guidance processes, firstly, a suitable phase difference is needed to provide the visiting spacecraft with space for long-range orbital ascent and guidance; secondly, the phase difference is closely coupled with the target spacecraft's orbital altitude and rendezvous / docking duration, requiring a phase difference that matches the long-range guidance strategy given a clear rendezvous / docking duration; finally, considering the uncertainty of the rendezvous time, the target spacecraft needs to be in a perfectly circular orbit to handle situations such as delayed rendezvous / docking. Therefore, for various current rendezvous / docking modes, the target spacecraft's orbital altitude, eccentricity, and phase difference must be strictly matched with the visiting spacecraft at the moment of its orbital insertion. To accommodate the visiting spacecraft's requirements for the target orbit, a phase modulation control strategy is generally designed for the visiting spacecraft's launch window before launch.

[0004] In related technologies, launch windows are typically planned before launch based on the target spacecraft's orbital parameters and predicted space environment conditions to achieve the desired phase difference, thereby enabling phasing control during subsequent orbit adjustments. However, these technologies struggle to fully account for uncertainties in the space environment, the impact of collision avoidance, and the need for long-term orbit maintenance. This leads to inaccurate launch window determination, resulting in lower efficiency and effectiveness of phasing control and unnecessary propellant consumption.

[0005] There is currently no effective solution to the above problems. Summary of the Invention

[0006] This invention provides a spacecraft rendezvous and docking method and apparatus based on step-by-step orbital elevation phase modulation control, which at least solves the technical problem in related technologies where the launch window cannot be accurately determined in spacecraft rendezvous and docking missions, resulting in low phase modulation control efficiency.

[0007] According to one aspect of the present invention, a spacecraft rendezvous and docking method based on step-by-step orbital phasing control is provided, comprising: determining the coplanar phase difference between a target spacecraft and an incoming spacecraft corresponding to each launch window range based on a preset launch window range, wherein the preset launch window range includes: multiple launch windows; the launch window is the date for launching the incoming spacecraft; the incoming spacecraft is the spacecraft that will rendezvous and dock with the target spacecraft; determining a phase error reference based on a determined date, wherein the determined date refers to the date used to determine the launch time of the incoming spacecraft; the phase error reference includes at least: prediction error and collision avoidance influence; determining a target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference, and the target phase difference, wherein the target phase difference is the phase difference when the incoming spacecraft rendezvous and docks with the target spacecraft; and launching the incoming spacecraft in the target launch window based on a step-by-step orbital phasing control strategy, wherein the step-by-step orbital phasing control strategy is used to control the orbital parameters of the target spacecraft so that the incoming spacecraft will rendezvous and dock with the target launch window.

[0008] Furthermore, the step of determining the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window, based on a preset launch window range, includes: determining the coplanar time, wherein the coplanar time is the time when the visiting spacecraft is launched; for each launch window, determining the first phase of the visiting spacecraft at the coplanar time of the launch window and the second phase of the target spacecraft at the coplanar time of the launch window; and determining the coplanar phase difference corresponding to each launch window based on the first phase and the second phase.

[0009] Further, the steps for determining the coplanar moment include: determining the right ascension of the first ascending node of the visiting spacecraft in the instantaneous true equatorial coordinate system; transforming the right ascension of the first ascending node to obtain the right ascension of the second ascending node of the visiting spacecraft in the inertial coordinate system; determining the virtual coplanar condition, wherein the virtual coplanar condition refers to the condition that the virtual target plane of the target spacecraft's orbit is coplanar with the orbital plane of the visiting spacecraft's entry into orbit, and the virtual target plane refers to the plane considering the drift difference of the ascending node; and determining the coplanar moment based on the virtual coplanar condition and the right ascension of the second ascending node.

[0010] Furthermore, the step of determining the phase error benchmark based on the determined date includes: determining the preset range in which the determined date falls, and determining the number of avoidance control operations based on the preset range; determining the collision avoidance impact amount during each avoidance control operation based on the target spacecraft's average daily orbits and the target spacecraft's operating speed; and determining the phase error benchmark based on the prediction error and the collision avoidance impact amount during each avoidance control operation.

[0011] Furthermore, before determining the phase error benchmark based on the forecast error and the collision avoidance impact during each avoidance control, the process also includes: constructing an atmospheric drag model, wherein the atmospheric drag model includes at least: atmospheric density variable, atmospheric drag coefficient, equivalent area variable, and spacecraft motion vector; and determining the forecast error based on the magnitude of the spacecraft perturbation factors using the atmospheric drag model.

[0012] Further, the step of determining the target transmission window within a preset transmission window range based on the coplanar phase difference, the phase error reference, and the target phase difference includes: determining the sum of the phase error reference and the target phase difference to obtain a comparison phase difference; selecting candidate coplanar phase differences greater than the comparison phase difference from all coplanar phase differences to obtain a set of candidate coplanar phase differences; determining the absolute difference between each candidate coplanar phase difference and the comparison phase difference; determining the candidate coplanar phase difference indicated by the smallest absolute difference as the target coplanar phase difference, and determining the transmission window indicated by the target coplanar phase difference as the target transmission window.

[0013] Furthermore, before launching the visiting spacecraft within the target launch window based on the step-by-step orbital phasing control strategy, the process includes: determining a first relationship between orbital phase, orbital angular velocity, and flight time, where the orbital phase is determined based on the orbital semi-major axis; determining a second relationship between orbital phase change and orbital semi-major axis change based on the first relationship; determining a third relationship between orbital phase change and velocity increment based on the second relationship and the relationship between orbital semi-major axis change and velocity increment; and constructing a phasing control model based on the first, second, and third relationships.

[0014] Furthermore, based on the step-by-step orbital ascent and phasing control strategy, the steps for launching the visiting spacecraft within the target launch window include: determining multiple constraints for phasing control, wherein the multiple constraints for phasing control include at least: phase difference constraints, eccentricity constraints, and semi-major axis constraints; determining nonlinear programming variables based on the phasing control model, wherein the nonlinear programming variables include: multiple orbital parameters; and adjusting the orbital parameters based on the multiple constraints for phasing control so that the phase difference between the visiting spacecraft and the target launch window during rendezvous and docking is equal to the target phase difference.

[0015] According to another aspect of the present invention, a spacecraft rendezvous and docking method based on step-by-step orbital ascent and phasing control is also provided, comprising: a first determining unit, configured to determine the coplanar phase difference between a target spacecraft and an incoming spacecraft corresponding to each launch window based on a preset launch window range, wherein the preset launch window range includes: multiple launch windows; the launch window is the date of launching the incoming spacecraft; and the incoming spacecraft is the spacecraft that will rendezvous and dock with the target spacecraft; and a second determining unit, configured to determine a phase error reference based on a determined date, wherein the determined date refers to the date used to determine the launch date of the incoming spacecraft. The date of the spacecraft's launch; the phase error reference includes at least: prediction error and collision avoidance impact; the third determining unit is used to determine the target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference, and the target phase difference, wherein the target phase difference is the phase difference when the visiting spacecraft rendezvous and docks with the target spacecraft; the launch unit is used to launch the visiting spacecraft within the target launch window based on the step-by-step orbital phasing control strategy, wherein the step-by-step orbital phasing control strategy is used to control the orbital parameters of the target spacecraft so that the visiting spacecraft rendezvous and docks with the target launch window.

[0016] Further, the first determining unit includes: a first determining module, used to determine the coplanar time, wherein the coplanar time is the time of launching the visiting spacecraft; a second determining module, used to determine, for each launch window, a first phase of the visiting spacecraft at the coplanar time of the launch window and a second phase of the target spacecraft at the coplanar time of the launch window; and a third determining module, used to determine the coplanar phase difference corresponding to each launch window based on the first phase and the second phase.

[0017] Further, the first determining module includes: a first determining submodule, used to determine the right ascension of the first ascending node of the visiting spacecraft in the instantaneous true equatorial coordinate system; a first transformation submodule, used to transform the right ascension of the first ascending node to obtain the right ascension of the second ascending node of the visiting spacecraft in the inertial coordinate system; a second determining submodule, used to determine the virtual coplanarity condition, wherein the virtual coplanarity condition refers to the condition that the virtual target plane of the target spacecraft's orbit is coplanar with the orbital plane of the visiting spacecraft's entry into orbit, and the virtual target plane refers to the plane considering the drift difference of the ascending node; and a third determining submodule, used to determine the coplanarity time based on the virtual coplanarity condition and the right ascension of the second ascending node.

[0018] Furthermore, the second determining unit includes: a fourth determining module, used to determine the preset range in which the determined date is located, and based on the preset range, determine the number of avoidance control operations; a fifth determining module, used to determine the collision avoidance impact amount during each avoidance control operation based on the target spacecraft's average number of orbits per day and the target spacecraft's operating speed; and a sixth determining module, used to determine the phase error benchmark based on the prediction error and the collision avoidance impact amount during each avoidance control operation.

[0019] Furthermore, the spacecraft rendezvous and docking also includes: a first construction module, used to construct an atmospheric drag model before determining the phase error benchmark based on the prediction error and the collision avoidance impact amount during each avoidance control, wherein the atmospheric drag model includes at least: atmospheric density variable, atmospheric drag coefficient, equivalent area variable, and spacecraft motion vector; and a seventh determination module, used to determine the prediction error based on the magnitude of the spacecraft perturbation factors using the atmospheric drag model.

[0020] Furthermore, the third determining unit includes: an eighth determining module, used to determine the sum of the phase error reference and the target phase difference to obtain a comparison phase difference; a first screening module, used to screen candidate coplanar phase differences that are greater than the comparison phase difference from all coplanar phase differences to obtain a set of candidate coplanar phase differences; a ninth determining module, used to determine the absolute difference between each candidate coplanar phase difference and the comparison phase difference; and a tenth determining module, used to determine the candidate coplanar phase difference indicated by the smallest absolute difference as the target coplanar phase difference, and to determine the transmission window indicated by the target coplanar phase difference as the target transmission window.

[0021] Furthermore, the spacecraft rendezvous and docking also includes: an eleventh determination module, used to determine the first relationship between orbital phase, orbital angular velocity, and flight time before launching the visiting spacecraft within the target launch window, based on a step-by-step orbital phasing control strategy, wherein the orbital phase is determined based on the orbital semi-major axis; a twelfth determination module, used to determine the second relationship between orbital phase change and orbital semi-major axis change based on the first relationship; a thirteenth determination module, used to determine the third relationship between orbital phase change and velocity increment based on the second relationship and the relationship between orbital semi-major axis change and velocity increment; and a second construction module, used to construct a phasing control model based on the first, second, and third relationships.

[0022] Furthermore, the launch unit includes: a fourteenth determining module, used to determine multiple constraints for phase modulation control, wherein the multiple constraints for phase modulation control include at least: phase difference constraints, eccentricity constraints, and semi-major axis constraints; a fifteenth determining module, used to determine nonlinear programming variables based on the phase modulation control model, wherein the nonlinear programming variables include: multiple orbital parameters; and a first adjusting module, used to adjust the orbital parameters based on the multiple constraints for phase modulation control, so that the phase difference between the visiting spacecraft and the target launch window during rendezvous and docking is equal to the target phase difference.

[0023] According to another aspect of the present invention, a computer program product is also provided, including a non-volatile computer-readable storage medium storing a computer program, wherein the computer program, when executed by a processor, implements the spacecraft rendezvous and docking method based on step-by-step orbit raising and phasing control as described above.

[0024] According to another aspect of the present invention, an electronic device is also provided, including one or more processors and a memory, wherein the memory is used to store one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors cause the one or more processors to implement any of the above-described spacecraft rendezvous and docking methods based on step-by-step orbital phasing control.

[0025] In this invention, based on a preset launch window range, the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window is determined. Based on a determined date, a phase error reference is determined. Based on the coplanar phase difference, the phase error reference, and the target phase difference, the target launch window is determined from the preset launch window range. Based on a step-by-step orbital phasing control strategy, the visiting spacecraft is launched within the target launch window. This solves the technical problem in related technologies where the launch window cannot be accurately determined in spacecraft rendezvous and docking missions, resulting in low phasing control efficiency.

[0026] In this invention, based on a preset launch window range, the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window is determined. Then, based on a determined date, a phase error benchmark is calculated. This benchmark comprehensively considers the phase effects caused by prediction errors and collision avoidance control. Subsequently, by analyzing the coplanar phase difference, the phase error benchmark, and the target phase difference, the target launch window is accurately selected from the preset launch window range. This ensures that the visiting spacecraft launched under this window can efficiently complete phase control in an orbit-ascending mode, avoiding unnecessary orbit-descending operations of the target spacecraft, effectively saving propellant, and ensuring its long-term stable operation. This achieves the technical effect of improving the efficiency of orbit-ascending phase control. Attached Figure Description

[0027] The accompanying drawings, which are included to provide a further understanding of the invention and form part of this invention, illustrate exemplary embodiments of the invention and are used to explain the invention, but do not constitute an undue limitation of the invention. In the drawings:

[0028] Figure 1 This is a flowchart of an optional spacecraft rendezvous and docking method based on step-by-step orbit raising and phase adjustment control according to an embodiment of the present invention;

[0029] Figure 2 This is a schematic diagram of an optional spacecraft rendezvous and docking device based on step-by-step orbital ascent and phase adjustment control according to an embodiment of the present invention;

[0030] Figure 3 This is a hardware structure block diagram of an electronic device (or mobile device) for a spacecraft rendezvous and docking method based on step-by-step orbital raising and phase adjustment control, according to an embodiment of the present invention. Detailed Implementation

[0031] To enable those skilled in the art to better understand the present invention, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort should fall within the scope of protection of the present invention.

[0032] It should be noted that the terms "first," "second," etc., used in this invention are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate so that the embodiments of the invention described herein can be implemented in orders other than those illustrated or described herein. Furthermore, the terms "comprising" and "having," and any variations thereof, are intended to cover non-exclusive inclusion; for example, a process, method, system, product, or apparatus that comprises a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0033] It should be noted that all related information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data used for analysis, stored data, and displayed data) collected and involved in this invention are information and data authorized by the user or fully authorized by all parties. Furthermore, the collection, storage, use, processing, transmission, provision, disclosure, and application of this data comply with the relevant laws, regulations, and standards of the relevant regions, necessary confidentiality measures have been taken, and it does not violate public order and good morals. Corresponding operation entry points are provided for users to choose to authorize or refuse. For example, this system has an interface with relevant users or organizations. Before obtaining relevant information, a request to obtain the information needs to be sent to the aforementioned user or organization through the interface, and the relevant information is obtained only after receiving consent from the aforementioned user or organization.

[0034] For near-Earth target spacecraft, in order to cope with the attenuation effect of atmospheric drag on the orbit, it needs to consume a large amount of propellant to maintain operation in the normal operating orbit and implement orbit maintenance control. During the operation of the target spacecraft, phasing control is required to cooperate with other visiting spacecraft to complete rendezvous and docking. The optimization space between the target spacecraft's orbit maintenance and phasing control requirements is determined by the launch window. Appropriate launch window planning can effectively reduce the probability of orbit phasing control, give full play to the orbit raising capability of orbit raising phasing, reduce propellant consumption, and achieve the target orbit phasing mission while completing orbit raising and maintenance.

[0035] Based on this, and addressing the phasing control requirements of visiting spacecraft, this invention proposes a method for designing a spacecraft launch window based on a step-by-step orbit-raising phasing control strategy, incorporating phase deviation redundancy design under the influence of orbit prediction errors and collision avoidance. According to the characteristics of orbit prediction errors gradually converging over time and the gradual weakening of collision avoidance effects, a step-by-step orbit-raising phasing control method is designed. This method achieves precise control of the phasing target parameters through a multi-stage, multi-step approach, significantly improving the probability of orbit-raising phasing control. Combined with multi-pulse step-by-step orbit-raising phasing, precise control of the phasing target parameters is achieved. Simulation results show that the launch window selected by this method can effectively cope with errors caused by the space environment in rendezvous and docking applications, completing the target spacecraft phasing control in an orbit-raising mode and avoiding propellant waste caused by the target spacecraft descending to a lower orbit.

[0036] The present invention will now be described in detail with reference to various embodiments.

[0037] Example 1

[0038] According to an embodiment of the present invention, an embodiment of a spacecraft rendezvous and docking method based on step-by-step orbital ascent and phasing control is provided. It should be noted that the steps shown in the flowchart in the accompanying drawings can be executed in a computer system such as a set of computer-executable instructions. Furthermore, although a logical order is shown in the flowchart, in some cases, the steps shown or described may be executed in a different order than that shown here.

[0039] Figure 1 This is a flowchart of an optional spacecraft rendezvous and docking method based on step-by-step orbit raising and phasing control according to an embodiment of the present invention, as shown below. Figure 1 As shown, the method includes the following steps:

[0040] Step S101: Based on the preset launch window range, determine the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window. The preset launch window range includes: multiple launch windows; the launch window is the date of launching the visiting spacecraft; and the visiting spacecraft is the spacecraft that rendezvous and docks with the target spacecraft.

[0041] In this embodiment of the invention, a series of possible launch time intervals (i.e., preset launch window ranges) can be predetermined based on engineering missions and space environment conditions. These intervals take into account various factors, such as solar illumination angle, sunlight suppression angle, space station operation arrangements, etc., and aim to provide multiple options for the launch of visiting spacecraft.

[0042] In this embodiment of the invention, the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window within a preset launch window range can be calculated. This coplanar phase difference is obtained by using the right ascension of the ascending node of the virtually coplanar visiting spacecraft as the aiming parameter, correcting the coplanar timing, and eliminating the orbital plane correction control amount after the visiting spacecraft enters orbit.

[0043] Here, the launch window is the date for launching the visiting spacecraft; the visiting spacecraft is the spacecraft that will rendezvous and dock with the target spacecraft; the coplanar phase difference is the difference in orbital position between the target and visiting spacecraft at the coplanar moment, expressed in angles, and is an important parameter for determining the launch timing and phasing control strategy. In orbital mechanics, the coplanar moment refers to the moment when the orbital planes of the target and visiting spacecraft coincide, at which point the orbital inclinations and right ascension of the ascending node of the two spacecraft are the same.

[0044] Step S102: Based on the determined date, determine the phase error benchmark, wherein the determined date refers to the date used to determine the launch time of the visiting spacecraft; the phase error benchmark includes at least: prediction error and collision avoidance effect.

[0045] In this embodiment of the invention, a phase error benchmark based on prediction and hazard avoidance can be constructed according to space environment errors and the probability of approaching hazardous targets in space. This benchmark takes into account the phase error under the influence of prediction errors and avoidance control. Here, prediction error refers to the uncertainty of orbit prediction, which gradually decreases over time; the impact of collision avoidance is the orbital adjustment required to avoid hazardous targets in space, such as debris and satellites, which increases linearly over time.

[0046] In this embodiment of the invention, since the timing of the launch window determination (i.e. the determination date, which is the date used to determine the time of launch of the visiting spacecraft) is different, the resulting prediction error and collision avoidance effect are also different. Therefore, the phase error benchmark can be determined based on the determination date.

[0047] Step S103: Based on the coplanar phase difference, phase error reference and target phase difference, determine the target launch window from the preset launch window range, wherein the target phase difference is the phase difference when the visiting spacecraft rendezvous and docks with the target spacecraft.

[0048] In this embodiment of the invention, the ideal phase difference (i.e., the target phase difference) when the visiting spacecraft and the target spacecraft rendezvous and dock can be determined first. The target phase difference is the ideal value that the actual phase needs to be adjusted to when designing the phase modulation control strategy in order to meet the conditions for successful rendezvous and docking.

[0049] In this embodiment of the invention, the optimal target launch window can be determined from the preset launch window range based on the coplanar phase difference, the phase error reference, and the target phase difference. For example, by comparing and analyzing the coplanar phase difference, the phase error reference, and the target phase difference, it can be determined which launch window can meet the requirements of phase modulation control. That is, if the launch is carried out under this window, the phase difference between the visiting spacecraft and the target spacecraft can reach or approach the target phase difference after phase modulation control.

[0050] Step S104: Based on the step-by-step orbital raising and phasing control strategy, launch the visiting spacecraft within the target launch window. The step-by-step orbital raising and phasing control strategy is used to control the orbital parameters of the target spacecraft so that the visiting spacecraft can rendezvous and dock with the target launch window.

[0051] In this embodiment of the invention, the visiting spacecraft can be launched at a coplanar moment within the target launch window. Based on a multi-stage, multi-step phasing control strategy (i.e., a step-by-step orbit raising and phasing control strategy, which is a redundant design to address prediction errors and avoid control effects, can improve the orbit raising interval of phasing control and reduce the probability of orbit lowering during the target orbit control process, thereby saving the propellant consumption of the target spacecraft), the phase difference between the target spacecraft and the visiting spacecraft is gradually reduced through phased orbit raising and adjustment until the target phase difference is reached.

[0052] In this embodiment of the invention, the step-by-step orbit raising and phasing control strategy optimizes the number of startups, startup times, and phase adjustment amounts to ensure that the orbital parameters of the target spacecraft gradually approach the design target after each phasing control, while reserving sufficient orbit raising space for the next phasing control to avoid orbit lowering and phasing, thus saving propellant.

[0053] In this embodiment of the invention, a multi-constraint launch window design method based on step-by-step orbital ascent and phasing control is established according to the requirements of long-term on-orbit target spacecraft orbit maintenance control and visiting spacecraft rendezvous and docking phase adjustment control. First, based on space environment errors and the probability of approaching hazardous targets in space, a phase error benchmark is constructed, primarily based on prediction and hazardous event avoidance. Then, based on the error benchmark, considering the target orbit and ascent control requirements, the optimal launch window is selected within the launch window selection range to meet the target spacecraft's orbital ascent and phasing control requirements under the influence of prediction errors and collision avoidance.

[0054] In summary, based on a preset launch window range, the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window can be determined. Then, based on a determined date, a phase error benchmark is calculated. This benchmark comprehensively considers the phase effects caused by prediction errors and collision avoidance control. Subsequently, by analyzing the coplanar phase difference, the phase error benchmark, and the target phase difference, the target launch window can be accurately selected from the preset launch window range. This ensures that the visiting spacecraft launched within this window can efficiently complete phase control in an ascent mode, avoiding unnecessary ascent operations for the target spacecraft, effectively saving propellant, and ensuring its long-term stable operation. This achieves the technical effect of improving the efficiency of ascent phase control, thereby solving the technical problem in related technologies where the launch window cannot be accurately determined in spacecraft rendezvous and docking missions, resulting in low phase control efficiency.

[0055] To improve the accuracy of determining the coplanar phase difference, in the spacecraft rendezvous and docking method based on step-by-step orbital ascent and phase adjustment control provided in Embodiment 1 of this application, the coplanar moment is determined, wherein the coplanar moment is the moment of launching the visiting spacecraft; for each launch window, the first phase of the visiting spacecraft at the coplanar moment of the launch window and the second phase of the target spacecraft at the coplanar moment of the launch window are determined; based on the first phase and the second phase, the coplanar phase difference corresponding to each launch window is determined.

[0056] In this embodiment of the invention, during the orbital rendezvous process, the visiting spacecraft consumes a large amount of propellant to correct its orbital plane. Therefore, the visiting spacecraft needs to be launched into an orbit with the same orbital inclination and right ascension of the ascending node as the target spacecraft, i.e., coplanar orbits. Based on the latitude of the launch site, selecting a suitable launch azimuth angle can achieve the same orbital inclination between the visiting and target spacecraft. However, achieving the same right ascension of the ascending node requires a launch window that meets constraints such as solar illumination angle and solar suppression angle. The time when the visiting spacecraft's orbital plane intersects with that of the target spacecraft (i.e., the coplanar moment) is chosen as its orbital insertion time to achieve coplanar orbital insertion.

[0057] Therefore, to achieve rendezvous and docking between a target spacecraft and a visiting spacecraft in low Earth orbit, and to reduce propellant consumption caused by orbital plane corrections, it is necessary to ensure that their orbits are coplanar. This means that the orbital inclinations and right ascension of the ascending nodes of the target and visiting spacecraft are completely identical. To avoid or minimize orbital plane correction control after orbit insertion, the coplanarity time can be determined first—that is, finding the launch time of the visiting spacecraft that guarantees orbital coplanarity. This time depends not only on the current orbital state of the target spacecraft but also on the launch site latitude, space environment, and other mission constraints (such as solar infiltration angle and solar suppression angle).

[0058] Then, for each launch window, it is necessary to first determine the first phase u of the visiting spacecraft at the coplanar moment of the launch window.chs and the second phase u of the target spacecraft at the coplanar moment of the launch window obj Phase refers to the position of a spacecraft in its orbit, that is, the angular position of the spacecraft relative to the ascending intersection point of the orbit.

[0059] In this embodiment of the invention, let the right ascension of the ascending node of the nominal orbital instantaneous true equatorial coordinate system of the visiting spacecraft be Ω. Gchs At a certain moment t, the right ascension of the ascending node of the target spacecraft is Ω. obj Then the right ascension of the ascending node of the visiting spacecraft at that moment is Ω. chs Ω chs =f(Ω) Gchs ,t), where f(Ω) Gchs ,t) is the right ascension of the ascending node of the J2000.0 inertial coordinate system obtained by performing a series of transformations such as precession, nutation, and polar motion on the ascending node of the instantaneous true equatorial coordinate system.

[0060] Let ΔΩ = Ω obj -Ω chs When ΔΩ = 0° or ΔΩ = 180°, the rendezvous and docking orbit of the target spacecraft and the orbit of the visiting spacecraft are coplanar. This moment t is defined as the coplanarity moment between the target spacecraft's rendezvous and docking orbit and the orbit of the visiting spacecraft, denoted as t0. pln Where ΔΩ = 0° means the target spacecraft and the visiting spacecraft have coplanar orbits with the same flight direction, and ΔΩ = 180° means the target spacecraft and the visiting spacecraft have coplanar orbits with opposite flight directions. Rendezvous and docking typically requires coplanar orbits with the same flight direction.

[0061] At time t, the orbits are coplanar pln The orbital phase difference (i.e., coplanar phase difference) between the target spacecraft and the visiting spacecraft is: du = u obj -u chs , where u obj To determine the latitude argument or phase (i.e., the second phase) of the target spacecraft at the moment of the visiting spacecraft's orbital insertion (i.e., the coplanar moment), u chs The latitude argument or phase (i.e., the first phase) of the visiting spacecraft at the moment of its orbital insertion.

[0062] It should be noted that after the target spacecraft's orbit is determined, the coplanarity of the orbits determines that the launch time of the visiting spacecraft's orbit insertion is a series of points in time, rather than a time interval. The coplanarity of the target spacecraft and the visiting spacecraft's orbit insertion determined by these points in time is called "nominal coplanarity".

[0063] In this embodiment, the need for orbital plane corrections is effectively reduced, thereby decreasing propellant consumption. Through precise phase control, it is ensured that the phase difference between the two spacecraft at the coplanar moment meets the requirements of the rendezvous and docking mission within the predetermined launch window. Furthermore, considering the uncertainties of the future space environment, the robustness of the control strategy is enhanced. Thus, not only is the selection of the launch window optimized, but the efficiency and accuracy of phase control are also improved.

[0064] To improve the accuracy of determining the coplanar moment, in the spacecraft rendezvous and docking method based on step-by-step ascending phase control provided in Embodiment 1 of this application, the first ascending node right ascension of the visiting spacecraft in the instantaneous true equatorial coordinate system is determined; the first ascending node right ascension is transformed to obtain the second ascending node right ascension of the visiting spacecraft in the inertial coordinate system; virtual coplanar conditions are determined, wherein the virtual coplanar conditions refer to the condition that the virtual target plane of the target spacecraft's orbit is coplanar with the orbital plane of the visiting spacecraft's entry into orbit, and the virtual target plane refers to the plane considering the ascending node drift difference; based on the virtual coplanar conditions and the second ascending node right ascension, the coplanar moment is determined.

[0065] In this embodiment of the invention, since the visiting spacecraft needs to track the target spacecraft via long-range guidance, the visiting spacecraft's orbital insertion must be lower than that of the target spacecraft. Due to the difference in their orbital altitudes, the ascending node drift also differs; the visiting spacecraft's lower orbit results in a slightly larger westward drift rate at the ascending node. When the target spacecraft's orbit and the visiting spacecraft's orbital insertion are "nominally coplanar," the initial difference in right ascension of the ascending node is 0. The longer the long-range rendezvous process, the greater the difference in ascending node drift, and the greater the difference in the orbital plane at the endpoint of the long-range guidance, resulting in a larger control amount required for orbital plane correction. To minimize the amount of plane correction control, the target spacecraft's orbital plane should be consistent with the "virtual target plane" considering ascending node drift difference compensation, i.e., "virtually coplanar."

[0066] In this embodiment of the invention, the right ascension Ω of the first ascending node of the visiting spacecraft in the instantaneous true equatorial coordinate system can be determined first. Gchs Here, the instantaneous true equatorial coordinate system is a coordinate system that changes with the Earth's rotation and can directly reflect the position of the spacecraft's orbit in the Earth coordinate system. The right ascension of the ascending node defines the orientation of the spacecraft's orbital plane relative to the equatorial plane, that is, the longitude angle from the equatorial north point to the ascending node on the equatorial plane.

[0067] Then, by transforming the right ascension of the first ascending node, we obtain the right ascension Ω of the second ascending node of the visiting spacecraft in the inertial coordinate system. chs Ω chs =f(Ω) Gchs ,t), where f(Ω) GchsThe right ascension of the ascending node in the J2000.0 inertial coordinate system is obtained by performing a series of transformations on the ascending node of the instantaneous true equatorial coordinate system, including precession, nutation, and polar motion. Here, the J2000.0 inertial coordinate system is a fixed coordinate system in space, independent of Earth's rotation. Therefore, the right ascension of the ascending node described in this inertial coordinate system more accurately reflects the spacecraft's positional relationship during long-term orbital operation, helping to eliminate the influence of Earth's own motion. This transformation process involves complex astronomical and orbital mechanics calculations, including corrections for precession, nutation, and polar motion, ensuring the accuracy of the second ascending node's right ascension.

[0068] Next, the virtual coplanarity condition is determined, which means that the orbital planes of the target spacecraft and the visiting spacecraft coincide in space, taking into account the influence of the ascending node drift difference. The ascending node drift difference is a phenomenon caused by the different effects of atmospheric drag on spacecraft at different orbital altitudes, resulting in a gradual change in the position of the ascending node over time. By calculating the ascending node drift difference between the target and visiting spacecraft, the difference in their orbital planes at a future moment can be predicted, thus setting the virtual coplanarity condition and providing a basis for subsequent launch window planning and phasing control. Then, based on the virtual coplanarity condition and the right ascension of the second ascending node, the coplanarity moment can be determined. This moment refers to the precise time when the virtual target plane of the target spacecraft and the orbital plane of the visiting spacecraft enter orbit are coplanar. Through precise calculations, the moment that satisfies the virtual coplanarity condition can be found, which is the optimal alignment point of the two spacecraft orbital planes after considering the difference in ascending node drift. The selection of this coplanarity moment is crucial for subsequent phasing control, determining the strategy and efficiency of phasing control, helping to reduce propellant consumption for orbital plane correction, optimizing spacecraft orbit maintenance, and ensuring a successful rendezvous and docking mission.

[0069] Specifically, we can set the endpoint time of long-distance guidance (i.e., the coplanar time), and the drift difference between the ascending nodes of the target spacecraft and the visiting spacecraft as δΩ. Then, the right ascension of the ascending node of the target spacecraft's "virtual target plane" at the time of the visiting spacecraft's orbit insertion is:

[0070] Therefore, the condition for the "virtual target plane" in the target spacecraft's orbit to be coplanar with the orbital plane of the visiting spacecraft (i.e., the virtual coplanar condition) is as follows: Therefore, the calculated coplanar time is (That is, the calculated time when ΔΩ = 0° or ΔΩ = 180°), at this coplanar moment, the orbital phase difference between the target spacecraft and the visiting spacecraft is: du = u obj -u chs This phase difference is the actual amount corrected by the phase modulation control.

[0071] In this embodiment, by taking into account the influence of the ascending node drift difference, the alignment of the future orbital plane can be predicted and controlled more accurately, avoiding unnecessary orbital plane corrections, saving propellant resources, and ensuring the orbital circularity of the target spacecraft. This provides a more economical and robust rendezvous and docking solution for spacecraft that operate in orbit for a long time.

[0072] To improve the accuracy of determining the phase error benchmark, in the spacecraft rendezvous and docking method based on step-by-step orbital elevation and phase adjustment control provided in Embodiment 1 of this application, a preset range in which the determined date is located is determined, and the number of avoidance control operations is determined based on the preset range; the collision avoidance impact amount during each avoidance control operation is determined based on the target spacecraft's average daily orbits and the target spacecraft's operating speed; and the phase error benchmark is determined based on the prediction error and the collision avoidance impact amount during each avoidance control operation.

[0073] In this embodiment of the invention, the determined date refers to the point in time used for planning and implementing the launch mission. It is typically any day within a specific time interval before the launch date, for example, T1 represents the timing of the launch window determination (i.e., the determined date). A preset range within which this determined date falls can be determined first. This preset range is a time period defined based on factors such as mission requirements, the accuracy of space environment forecasts, and the effectiveness of collision avoidance strategies. For example, it could be a range from 60 days to 40 days before launch, or from 40 days to 7 days before launch. After determining this range, the number of avoidance control operations can be determined based on this preset range; that is, how many collision avoidance control operations the target spacecraft might need to perform before launching the visiting spacecraft. The number of avoidance control operations directly affects the magnitude of future phase errors, and is therefore a key parameter. Specifically, this number can be estimated by statistically analyzing historical data, analyzing the distribution of space debris, and predicting potential collision risks.

[0074] Then, the average number of orbits N per day of the target spacecraft can be used as a basis. day The target spacecraft's operating speed V is used to determine the collision avoidance effect U during each avoidance control operation. ri , among which, U ri This represents the phase effect of the i-th collision avoidance on the launch time.

[0075] Then, based on the prediction error dU and the collision avoidance influence U during each avoidance control, ri The phase error reference σU is determined. Here, the forecast error refers to the deviation between the predicted orbit and the actual orbit position caused by uncertainties in atmospheric density models, equivalent area calculations, etc.

[0076] Taking into account both forecast errors and collision avoidance effects, the design error baseline for launch T1 days prior is as follows:

[0077]

[0078] Where dU represents the phase prediction error for day T1; n is the total number of collision avoidance attempts within day T1.

[0079] For example, let T1 represent the timing of determining the launch window and T0 represent the launch window. If the launch window of the visiting spacecraft is determined 1 day in advance, the phase prediction error dU is calculated 1 day in advance based on the orbital adjustment and phasing control requirements, and the phase error reference σU is calculated in conjunction with collision avoidance.

[0080] (1) If 60 > T1 >= 40, the number of avoidance control operations n = 2, and the effects of avoidance control on the phase are as follows:

[0081]

[0082] Among them, U r1 This indicates the phase effect of the first collision avoidance maneuver on the launch time; U r2 This indicates the phase effect of the second collision avoidance maneuver on the launch timing; N day V represents the average number of orbits the target spacecraft completes per day; V represents the target spacecraft's orbital speed; Δv i This represents the evasion control quantity for each collision avoidance, where i = 1, 2.

[0083] Then the T1-day phase error reference σU is:

[0084]

[0085] (2) If 40 > T1 >= 7, the number of avoidance control operations n = 1, and the effects of avoidance control on the phase are as follows:

[0086]

[0087] Then the T1-day phase error reference σU is:

[0088]

[0089] In this embodiment of the invention, the main direction of trajectory prediction error is considered to be the trajectory flight direction. The other two directions are more accurately predicted than the flight direction, but normal control will cause the orbital plane to deviate from the designed trajectory. Therefore, when designing avoidance strategies, radial distance is generally used as the avoidance target quantity. When there are many dangerous targets and the radial avoidance cost is high, an avoidance method combining flight direction and radial distance is adopted.

[0090] Avoidance control quantity Δv i The approximate formula for the influence of the phase is: Among them, U riLet N represent the phase impact of the i-th collision avoidance maneuver on the launch time, N represent the difference in the number of orbits between the avoidance control and launch time, and V represent the target spacecraft's velocity. Taking a near-Earth spacecraft at an altitude of 393 km as an example, which orbits approximately 15.5 times per day, a 0.5 m / s collision avoidance control maneuver is implemented 20 days before the visiting spacecraft's launch. The impact on the target's phase is approximately 20.6 degrees after 20 days, approximately 31 degrees after 30 days, and approximately 50 degrees after 50 days. The phase impact of the avoidance control maneuver increases linearly with time. Therefore, the impact of avoidance control must be considered when designing the phase error.

[0091] In this embodiment, by meticulously dividing the preset range and calculating the number of avoidance control operations, future changes in the orbital environment can be predicted, thereby accurately setting the phase error benchmark. This benchmark guides the selection of launch windows and the design of phasing strategies, fully considering the uncertainties of the space environment and the inevitability of collision avoidance. It ensures that the target spacecraft and the visiting spacecraft can successfully complete the rendezvous and docking mission at the correct phase, eccentricity, and orbital altitude, while minimizing propellant consumption and providing effective support for the long-term on-orbit operation of the spacecraft.

[0092] To improve the accuracy of forecast error determination, in the spacecraft rendezvous and docking method based on step-by-step orbital ascent and phasing control provided in Embodiment 1 of this application, before determining the phase error benchmark based on the forecast error and the collision avoidance impact during each avoidance control, an atmospheric drag model is constructed. The atmospheric drag model includes at least: atmospheric density variable, atmospheric drag coefficient, equivalent area variable, and spacecraft motion vector. Based on the magnitude of spacecraft perturbation factors, the atmospheric drag model is used to determine the forecast error.

[0093] In this embodiment of the invention, the accuracy of orbit and phase prediction is of great significance for strategy formulation and performance design for flight control and rendezvous / docking missions. The perturbation effects considered in orbit determination and prediction mainly include Earth mass perturbation, non-spherical gravitational perturbation, atmospheric drag perturbation, lunar and solar gravitational perturbation, and solar radiation pressure perturbation. Atmospheric drag, apart from Earth's non-spherical perturbation, is the most significant perturbation affecting the accuracy of spacecraft orbit determination and prediction.

[0094] According to the theory of free molecular flow, atmospheric molecules undergo various motions after colliding with the spacecraft surface, including retention, diffuse scattering, and specular reflection. The atmospheric drag model is as follows:

[0095]

[0096] Among them, F D Where ρ is atmospheric drag, v is the atmospheric density at the spacecraft's location, Cd is the atmospheric drag coefficient, S is the equivalent area of ​​atmospheric drag on the spacecraft in the velocity direction, and u is the atmospheric drag coefficient. vThis is the unit velocity vector of the spacecraft relative to the atmosphere.

[0097] As can be seen from the atmospheric drag model, the key to high-precision orbit prediction lies in determining the model parameters such as ρ, S, and Cd, and the prediction strategy. Given the difficulty in determining the exact value of Cd, the equivalent area can be calculated first. Then, Cd can be treated as an unknown quantity and solved together with the spacecraft's motion vector in the orbit calculation. The solved Cd, to some extent, compensates for the errors in atmospheric drag caused by model errors in atmospheric density and calculation errors in the equivalent area, resulting in a better fit between the dynamic model and the observation data.

[0098] However, atmospheric density is highly random, and various influencing factors are quite complex. Most current atmospheric density models are semi-empirical models, with errors ranging from approximately 10% to 20%.

[0099] Table 1 shows the magnitude of perturbation factors for low-Earth orbit spacecraft.

[0100] Table 1

[0101]

[0102] Table 1 presents the perturbation magnitudes of the main perturbation factors affecting the orbital motion of low-Earth orbit spacecraft. It can be seen that, apart from atmospheric density, the mathematical models for other perturbation factors affecting spacecraft orbital motion are relatively complete. The uncertainty of atmospheric density is the most critical factor affecting the accuracy of low-Earth orbit spacecraft orbit prediction.

[0103] In designing the phase prediction error, the Cd calculated under different space environments and equivalent areas in the orbit calculation is used as a benchmark. The calculated Cd values ​​for each stage are used for prediction. Taking into account the errors of the space environment and aerodynamic characteristics, a Cd coefficient deviation of 20% is set as the simulation condition to calculate the medium- and long-term prediction error. That is, the prediction error can be determined based on the magnitude of spacecraft perturbation factors using an atmospheric drag model.

[0104] In this embodiment, by meticulously constructing an atmospheric drag model and accurately estimating forecast errors, launch windows can be planned earlier and more rationally. This ensures that the target spacecraft and the visiting spacecraft rendezvous and dock under optimal orbital parameters, while minimizing propellant consumption and providing effective support for the long-term on-orbit operation of the spacecraft. Thus, not only is the safety and success rate of space missions improved, but the cost of space missions is also reduced.

[0105] To improve the accuracy of determining the target launch window, in the spacecraft rendezvous and docking method based on step-by-step orbital ascent and phase adjustment control provided in Embodiment 1 of this application, the sum of the phase error reference and the target phase difference is determined to obtain the comparison phase difference; candidate coplanar phase differences greater than the comparison phase difference are selected from all coplanar phase differences to obtain a set of candidate coplanar phase differences; the absolute difference between each candidate coplanar phase difference and the comparison phase difference is determined; the candidate coplanar phase difference indicated by the smallest absolute difference is determined as the target coplanar phase difference, and the launch window indicated by the target coplanar phase difference is determined as the target launch window.

[0106] In this embodiment of the invention, the specific launch date is determined in advance based on the selectable range of launch dates. The main basis for this determination is the coplanar phase difference between the target spacecraft and the visiting spacecraft within the predetermined launch window. The launch window is designed based on a step-by-step orbital phasing control strategy. The selected specific launch date must allow for orbital phasing control space in the phase difference at the coplanar moment of the two targets. That is, when selecting the launch window for the visiting spacecraft, uncertain influencing factors such as orbit prediction errors and collision avoidance control must be comprehensively considered. A date with a phase adjustment amount greater than the error benchmark is selected as the first launch window to reduce the probability of implementing orbital phasing.

[0107] In this embodiment of the invention, the process for determining the target launch window is as follows:

[0108] (1) Define the leading edge T and range (i days) of the selectable launch window dates;

[0109] (2) Calculate the daily coplanar phase difference U within the selected launch window range. i ;

[0110] (3) Based on the coplanar phase difference U i Target phase difference U t The phase error reference σU is used to select the transmission window.

[0111] The design principles for the launch window are as follows:

[0112] U i -(U t +σU)>0, Δu=min(U i -(U t +σU));

[0113] The date T0 that satisfies the condition that the coplanar phase difference is greater than the sum of the phase differences between the day's error reference and the target, and that has the smallest difference from the sum of the two, will be used as the main window (i.e. the target launch window).

[0114] The launch window is determined according to this principle. The phase prediction error and the amount of phase adjustment for collision avoidance are larger than the actual flight process. During the phase adjustment process, the probability of orbit de-orbiting is relatively small.

[0115] It should be noted that it is not recommended to determine the launch date too early. If the launch date is determined too early, the orbit prediction deviation and collision avoidance will be significantly affected. If phase adjustment control is carried out according to the full-ascent orbit, a large amount of phase reserve is required, resulting in more control operations, larger speed increments, and the rendezvous and docking orbit height exceeding the limit.

[0116] Specifically, the sum of the phase error reference and the target phase difference U can be determined first. t +σU, this value is called the comparison phase difference. Here, the phase error benchmark, after considering prediction errors and collision avoidance control effects under a defined launch window, represents the phase difference reserve between the target spacecraft and the visiting spacecraft, ensuring the effectiveness and flexibility of the phase adjustment control strategy. The target phase difference is the ideal phase difference when the visiting and target spacecraft complete rendezvous and docking, and is the target value for designing the phase adjustment control strategy. The comparison phase difference obtained by adding these two values ​​represents the minimum phase adjustment space that the target spacecraft must reserve to accommodate the phase adjustment requirements of the visiting spacecraft after it enters orbit, under the selected launch window.

[0117] Then, candidate coplanar phase differences greater than the comparison phase difference are selected from all coplanar phase differences to form a candidate coplanar phase difference set. Coplanar phase difference refers to the difference in orbital position between the target spacecraft and the visiting spacecraft at coplanar moments under different launch windows. During the selection process, only those coplanar phase differences with a reserved phase adjustment amount greater than the comparison phase difference are retained. This ensures that the target spacecraft has sufficient phase adjustment space when performing phasing control to cope with uncertainties in the space environment, such as changes in atmospheric density and collision avoidance requirements, while ensuring a successful entry into the designed rendezvous and docking orbit. Next, the absolute difference between each candidate coplanar phase difference and the comparison phase difference can be determined. This absolute difference reflects the difference between the actual phase adjustment amount of the target spacecraft and the minimum required phase adjustment amount under the candidate launch window. By calculating these differences, the efficiency and propellant consumption of the target spacecraft's phasing control strategy under each candidate launch window can be quantified. Finally, the candidate coplanar phase difference indicated by the minimum absolute difference is determined as the target coplanar phase difference, and based on this target coplanar phase difference, the indicated launch window is determined as the target launch window. Choosing a coplanar phase difference with the smallest absolute difference means selecting a launch window that minimizes propellant consumption and optimizes the phasing control strategy. This decision-making process fully considers space environment prediction errors, collision avoidance control parameters, and mission requirements, ensuring that the selection of the launch window not only meets the requirements of coplanarity and phase alignment but also minimizes propellant consumption, thereby improving the mission's economy and feasibility.

[0118] In this embodiment, the launch window selection is effectively optimized, ensuring phase synchronization between the target spacecraft and the visiting spacecraft during the rendezvous and docking mission. Simultaneously, propellant consumption is significantly reduced, improving the efficiency and economy of the space mission. This approach is not only applicable to specific space station missions but can also be widely applied to the long-term on-orbit operation of large spacecraft requiring frequent rendezvous and docking. It provides a new solution for spacecraft orbit maintenance and resource management, playing a crucial role in promoting the development of space technology. By minimizing propellant consumption during phasing control, the on-orbit lifespan of the spacecraft can be extended, the frequency of resupply missions reduced, the overall cost of space exploration and operation lowered, and the flexibility and adaptability of space missions improved.

[0119] To accurately construct the phase modulation control model, in the spacecraft rendezvous and docking method based on step-by-step orbital elevation phase modulation control provided in Embodiment 1 of this application, before launching the visiting spacecraft within the target launch window based on the step-by-step orbital elevation phase modulation control strategy, a first relationship between orbital phase, orbital angular velocity, and flight time is determined, wherein the orbital phase is determined based on the orbital semi-major axis; based on the first relationship, a second relationship between orbital phase change and orbital semi-major axis change is determined; based on the second relationship and the relationship between orbital semi-major axis change and velocity increment, a third relationship between orbital phase change and velocity increment is determined; based on the first, second, and third relationships, a phase modulation control model is constructed.

[0120] In this embodiment of the invention, the relationship between the spacecraft's orbital phase u and its orbital angular velocity n and flight time t (i.e., the first relationship) is: u = n·t, where, a is the semi-major axis of the orbit, and μ is the Earth's gravitational constant.

[0121] The orbital phase change is: Δu=Δn·t+n·Δt, where Δn·t is an isochronous change and n·Δt is a time-varying change. This formula shows that the changes in orbital parameters generated by orbital control and the timing of orbital control can both be used as planning variables for orbital phase adjustment.

[0122] Considering isochronous variations, the relationship between the spacecraft's orbital phase change Δu and the semi-major axis change Δa (i.e., the second relationship) is as follows:

[0123] Therefore, the change Δa in the semi-major axis of the orbit required for the orbital phase adjustment Δu is: Where, Δu=u obj -u chs , which is the orbital phase difference between the target spacecraft and the visiting spacecraft at the moment when their orbits are in the same plane.

[0124] Based on the relationship between the change in the semi-major axis of the spacecraft's orbit Δa and the velocity increment Δv, the relationship between Δu and Δv (i.e., the third relationship) is obtained as follows:

[0125]

[0126] Where e is the orbital eccentricity, f is the true anomaly angle, and p = a(1-e) 2 () is the semi-blind diameter.

[0127] In this embodiment of the invention, a phasing control model can be constructed based on the first, second, and third relationships determined above. This model integrates the mathematical relationships between orbital phase, orbital angular velocity, flight time, orbital semi-major axis variation, and velocity increment, providing theoretical support for the implementation of a step-by-step orbital phasing control strategy. Through the construction of this model, adjustments to the spacecraft's orbital parameters can be planned to achieve the desired orbital phase target, while also considering the economical use of propellant and the precision of phasing control.

[0128] In this embodiment of the invention, there are two track phasing control methods: one is track raising phasing, and the other is track lowering phasing. When the semi-major axis of the track increases, the track angular velocity decreases, and the track phase angle decreases. Therefore, track raising phasing is suitable for reducing the track phase angle. When the semi-major axis of the track decreases, the track angular velocity increases, and the track phase angle increases. Therefore, track lowering phasing is suitable for increasing the track phase angle.

[0129] It should be noted that the phase angle range for spacecraft orbit phasing is 0-360°. If phasing is performed within a phase angle range of ±180° using the velocity increment optimization control method, there are two control modes: orbit raising and orbit lowering. To meet the orbit circularity requirements after orbit phasing control, when using orbit lowering phasing, orbit control is preferentially performed at perigee; when using orbit raising phasing, orbit control is preferentially performed at apogee.

[0130] In this embodiment, by constructing a detailed phasing control model, changes in orbital phase can be predicted and controlled, thereby optimizing the selection of the launch window, reducing propellant consumption, and ensuring orbital synchronization between the target spacecraft and the visiting spacecraft during rendezvous and docking missions. This phasing control strategy based on orbital parameter variations is not only applicable to large spacecraft operating in orbit for extended periods, but also effectively addresses uncertainties in the space environment, such as changes in atmospheric drag and collision avoidance requirements.

[0131] To achieve precise rendezvous and docking between the target spacecraft and the visiting spacecraft, the spacecraft rendezvous and docking method based on step-by-step orbital phasing control provided in Embodiment 1 of this application determines multiple constraints for phasing control. These constraints include at least: phase difference constraints, eccentricity constraints, and semi-major axis constraints. Based on the phasing control model, nonlinear programming variables are determined, including multiple orbital parameters. Based on the multiple constraints for phasing control, the orbital parameters are adjusted so that the phase difference between the visiting spacecraft and the target launch window during rendezvous and docking is equal to the target phase difference.

[0132] In this embodiment of the invention, the phase control for rendezvous and docking not only constrains the phase relationship at coplanar moments but also imposes requirements on the semi-major axis and eccentricity. Furthermore, due to the continuous attenuation effect of the thin atmosphere in near-Earth space on spacecraft orbits, spacecraft operating in orbit for extended periods require periodic orbital elevation to maintain operation at the target orbital altitude. Therefore, to conserve propellant and ensure long-term stable operation of the target spacecraft, orbit elevation and phase control become a strong constraint in engineering implementation.

[0133] In this embodiment of the invention, the step-by-step phase adjustment control achieves precise correction of the phase adjustment target parameters through a phased, multi-step, and dynamically optimized approach, balancing the requirements for track lifting, control accuracy, and adaptability. The design parameters can be continuously adjusted over time to meet the constraints of the terminal semi-major axis, eccentricity, and phase. The variables in this nonlinear programming are:

[0134] X = [t1, t2, t3...t] i ,Δv1,Δv2,Δv3,....Δv i ] T ;

[0135] Among them, t i Δv represents the i-th phase modulation control time. i This represents the speed increment during the i-th phase modulation control, where i represents the number of start-ups (i.e., the number of phase modulation control operations).

[0136] The planned terminal conditions (i.e., multiple constraints for phase modulation control) are as follows:

[0137] Terminal phase difference: U = U tar , among which, U tar This indicates the target phase difference.

[0138] Terminal eccentricity: 0 < e < e max , where e max This represents the maximum eccentricity.

[0139] Terminal semi-major axis: a min <a<a max , where amin a represents the least semi-major axis. max This indicates the largest semi-major axis.

[0140] In this embodiment of the invention, considering that the determining factor of the velocity increment among the variables is the phase adjustment amount, the variables of the nonlinear programming can be adjusted according to the relationship between the phase adjustment amount Δu and the velocity increment Δv:

[0141] X = [t1, t2, t3...t] i ,Δu1,Δu2,Δu3,....Δu i ] T ;

[0142] Where, Δu i This represents the phase adjustment amount for the i-th phase modulation control.

[0143] Due to the constraints of the rail-raising phase adjustment control, a rail-raising control amount needs to be reserved for subsequent phase adjustment control during each phase adjustment control. That is, the reserved phase adjustment amount must be greater than the error reference σu for subsequent phase changes. i The intermediate constraints of nonlinear programming are expressed as functions:

[0144]

[0145] Where du is the initial phase adjustment from the current forecast to the coplanar time, and t i Let Δu be the time of the i-th phase modulation control. i Let Δu be the phase adjustment amount for the i-th control, and n represent the total number of control operations. k This represents the phase adjustment amount for each operation, where k represents the k-th control, and σu k This represents the phase error reference after the k-th control. During the step-by-step phase adjustment process of the orbital ascent, the phase reserve for the last step is 0, while the phase reserves for the remaining steps are not less than the control time t. i The corresponding phase prediction error benchmark σU is determined by optimizing the number of power-on cycles i and the power-on time t. i Phase adjustment amount Δu i The orbital parameters are used to target the final phase-modulated target semi-major axis a, eccentricity e, and target phase difference U.

[0146] During the implementation of the project, taking into account the impact of collision warning and the accuracy of phase control, the optimal time for the start-up of the last phase control is 2 to 3 days before the launch of the visiting spacecraft. Other variables are optimized according to the phase control target. The final optimization target is to minimize the number of orbit control operations or optimize the velocity increment.

[0147] In this embodiment, by clearly defining the multiple constraints of phase modulation control, the phase modulation control problem is transformed into a nonlinear programming problem. By dynamically optimizing the orbital parameters, not only can the precise phase alignment between the target spacecraft and the visiting spacecraft be ensured during rendezvous and docking, but the economical use of propellant and the long-term stability of the orbit can also be fully considered during the process, avoiding unnecessary orbit reduction operations, thereby saving propellant resources and extending the on-orbit life of the spacecraft.

[0148] The following simulation verification will be used to verify the effectiveness of the spacecraft rendezvous and docking scheme based on the step-by-step orbital elevation and phase adjustment control strategy proposed in this embodiment.

[0149] (1) Window design and phase modulation strategy verification.

[0150] 1) Track input.

[0151] The simulated orbital parameters used in the trial calculation are shown in Table 2, with the orbital epoch being approximately -50 days from the leading edge of the window.

[0152] Table 2

[0153]

[0154] The orbital insertion parameters of the visiting spacecraft and the phasing target parameters of the target spacecraft are shown in Table 3.

[0155] Table 3

[0156]

[0157]

[0158] Based on the window design, the comprehensive phase reserve and target phase value corresponding to different phase control dates were calculated, and the results are shown in Table 4. The relationship between the comprehensive reserved phase and the target phase is: Comprehensive reserved phase + Nominal phase difference for rapid rendezvous and docking (11°) = Target phase.

[0159] Table 4

[0160]

[0161] 2) Select the launch window.

[0162] Assume the launch window is selected between January 10th and 15th, 2024. Approaching the launch date, the target spacecraft's average orbital altitude is approximately 384.5 km. At this altitude, the relationship between the launch window and phase is that for every day the launch window is delayed, the coplanar phase difference will increase by approximately 130°. Table 5 shows the coplanar phase differences corresponding to different launch windows.

[0163] Table 5

[0164]

[0165]

[0166] The launch window is determined based on the principle that "coplanar phase difference - target phase >", while also minimizing the increase in the orbital phase adjustment velocity. It can be seen that:

[0167] If the launch window is determined in -50 days and the target phase is 134.5°, it is recommended to choose January 15th as the first launch window.

[0168] If the launch window is determined in -45 days and the target phase is 114.0°, it is recommended to choose January 12 as the first launch window.

[0169] If the launch window is determined in -40 days and the target phase is 93.00°, it is recommended to choose January 12 as the first launch window.

[0170] If the launch window is determined in -30 days and the target phase is 58.80°, it is recommended to choose January 12 as the first launch window.

[0171] Assuming a launch window of -45 days, January 12th is selected as the first launch window. Based on this, calculations and analyses of the phasing control between the target spacecraft and the visiting spacecraft are conducted.

[0172] 3) Phase modulation strategy design.

[0173] -45 days later, January 12th was determined as the launch window, with a coplanar timing phase difference of 131.559°. A step-by-step orbital ascent and phase adjustment control strategy was adopted, and a three-pulse strategy and a two-pulse strategy were designed.

[0174] The three-pulse orbital ascent phasing control strategy is as follows: Considering propellant consumption and phase prediction accuracy, the execution time of the first pulse is scheduled between -35 and -30 days, the second pulse between -25 and -6 days, and the third pulse between -2 days. The execution times of the first and second pulses are iterated, the velocity increment consumed by the three-pulse phasing is calculated, and a solution set satisfying the constraints is selected.

[0175] Table 6 presents the calculation results for the first pulse being set at -35d.

[0176] Table 6

[0177]

[0178]

[0179] Table 7 shows the calculation results for the first pulse being set at -30d.

[0180] Table 7

[0181]

[0182] The calculation results show that for the three-pulse phase modulation strategy, the later the execution time of the first pulse, the more speed increment the first pulse consumes; when the execution time of the first pulse is determined, the later the execution time of the second pulse, the more speed increment the second pulse consumes, and the less speed increment the third pulse consumes. However, the total speed increment of the three pulses shows a trend of first decreasing and then increasing, and the total speed increment has a minimum value.

[0183] Dual-pulse orbital ascent phasing control strategy: Considering propellant consumption and phase prediction accuracy, the execution time of the first pulse is scheduled between -35 days and -6 days, and the execution time of the second pulse is scheduled between -2 days. The execution time of the first pulse is iterated, and the velocity increment consumed by the two pulses during phasing is calculated. A solution set that satisfies the constraints is then selected.

[0184] Table 8 presents the dual-pulse phase modulation control strategy.

[0185] Table 8

[0186]

[0187] The calculation results show that the later the first pulse is executed, the more velocity increment it consumes, while the second pulse consumes less velocity increment. The total velocity increment of the two pulses shows a trend of first decreasing and then increasing, and there is a minimum value in the total velocity increment.

[0188] (2) Error simulation verification.

[0189] 1) Track input.

[0190] Table 9 provides the orbital parameters.

[0191] Table 9

[0192] Parameter Description numerical values coordinate system J2000 Era (BJT) 2023-12-26T08:00:00 Semi-major axis (m) 6760548.463 Eccentricity 0.0014673 Track inclination angle (°) 41.615 Right ascension of the ascending node (°) 67.080 Argument of perigee (°) 118.594 Angle of approach (°) 60.043 Average altitude (km) 385.151 AP 8.52 F107 153.7 F107P 153.4

[0193] January 17, 2023 was selected as the launch window. The initial phase difference was calculated to be 123.2° 30 days before launch. The orbital parameters of the visiting spacecraft and the phase adjustment target parameters of the target spacecraft are shown in Table 10.

[0194] Table 10

[0195]

[0196] 2) Design of nominal phase modulation strategy.

[0197] A two-pulse step-by-step phase modulation control strategy is adopted, with the two phase modulation controls scheduled 18 days and 2 days before the launch of the visiting spacecraft, respectively.

[0198] Table 11 presents the nominal phase modulation strategies.

[0199] Table 11

[0200] Parameter name TZ07 Phase Adjustment Initial phase difference (°) 123.2 Boot time (BJT) 2023-12-29T15:55:40 Velocity increment (m / s) 2.371 Booting time 2 (BJT) 2024-01-15T17:26:43 Velocity increment 2 (m / s) 2.754 Target orbital altitude (km) 388.602 Target orbital eccentricity 0.0006

[0201] 3) Error simulation verification.

[0202] Among them, the space environment error analysis before the first phase adjustment: if the space environment forecast parameter value increases by 20% before the first phase adjustment, the phase forecast deviation mainly affects the control quantity of the first phase adjustment. The calculation results of the control parameters are shown in Table 12.

[0203] Table 12

[0204] Parameter name TZ07 Phase Adjustment Initial phase difference (°) 137.9 Boot time (BJT) 2023-12-29T15:55:19 Velocity increment (m / s) 2.696 Booting time 2 (BJT) 2024-01-15T17:23:21 Velocity increment 2 (m / s) 2.795 Target orbital altitude (km) 388.230 Target orbital eccentricity 0.000592

[0205] If the prediction deviation increases by 20% before the first phase adjustment, the initial phase difference will increase, and the amount of the first phase adjustment will also increase. Therefore, the speed increment of the first phase adjustment control will increase by approximately 0.33 m / s.

[0206] If the space environment forecast parameter values ​​are reduced by 20% before the first phase adjustment, the control parameter calculation results are shown in Table 13.

[0207] Table 13

[0208]

[0209]

[0210] If the predicted deviation before the first phase adjustment is reduced by 20%, the initial phase difference will be smaller, and the amount of the first phase adjustment will be smaller. The speed increment of the first phase adjustment control will be reduced by approximately 0.37 m / s.

[0211] Space environment error analysis after the first phase adjustment: If the space environment forecast parameter value increases by 20% after the first phase adjustment, the phase forecast deviation will have a significant impact on the second phase adjustment control. The calculation results of the control parameters are shown in Table 14.

[0212] Table 14

[0213] Parameter name TZ07 Phase Adjustment Initial phase difference (°) 123.2 Boot time (BJT) 2023-12-29T15:55:40 Velocity increment (m / s) 2.371 Booting time 2 (BJT) 2024-01-15T17:23:52 Velocity increment 2 (m / s) 4.872 Target orbital altitude (km) 391.463 Target orbital eccentricity 0.001118

[0214] If the prediction deviation increases by 20% after the first phase adjustment, the subsequent phase difference will be larger, and the second phase adjustment amount will also be larger. Therefore, the second phase adjustment control speed increment will increase by approximately 2.118 m / s.

[0215] If the space environment forecast parameter values ​​decrease by 20% after the first phase adjustment, the control parameter calculation results are shown in Table 15.

[0216] Table 15

[0217] Parameter name TZ07 Phase Adjustment Initial phase difference (°) 107.4 Boot time (BJT) 2023-12-29T15:55:40 Velocity increment (m / s) 2.371 Booting time 2 (BJT) 2024-01-15T17:29:24 Velocity increment 2 (m / s) 0.766 Target orbital altitude (km) 385.989 Target orbital eccentricity 0.000154

[0218] If the prediction deviation decreases by 20% after the first phase adjustment, the subsequent phase difference will be smaller, and the second phase adjustment amount will be smaller. Therefore, the second phase adjustment control speed increment will decrease by approximately 2 m / s.

[0219] Simulation results show that determining the launch date too early leads to significant deviations in orbit prediction and impacts on collision avoidance. Designing a phase adjustment strategy based on orbit ascent requires a large reserved phase, which can easily cause problems such as numerous control operations, large velocity increments, and exceeding rendezvous and docking orbit altitude limits. The multi-pulse orbit ascent phase adjustment control strategy has better overall adaptability and stronger redundancy for velocity increments and orbit altitudes. However, due to engineering constraints, in specific implementation, it is necessary to comprehensively select a strategy with fewer control operations and that meets mission requirements based on the launch window. The launch window based on the step-by-step orbit ascent phase adjustment control design can effectively cope with the impact of space environment deviations, ensure that the target spacecraft completes orbit ascent phase adjustment control, and avoid propellant waste.

[0220] In this embodiment of the invention, addressing the rendezvous and docking requirements of spacecraft, a method is proposed based on a step-by-step orbital ascent and phasing control strategy. This method incorporates phase deviation redundancy design for the spacecraft launch window, taking into account orbital prediction errors and the impact of collision avoidance. Based on the characteristics of orbital prediction errors gradually converging over time and the gradual weakening of collision avoidance effects, a step-by-step orbital ascent and phasing control method is designed. This method achieves precise control of the phasing target parameters through a multi-stage, multi-step approach, significantly increasing the probability of successful orbital ascent and phasing control. Combined with multi-pulse step-by-step orbital ascent and phasing, precise control of the phasing target parameters is achieved. Simulation results show that selecting the launch window based on the proposed design method can effectively address errors caused by the space environment during the rendezvous and docking preparation phase, completing the target spacecraft phasing control in ascent mode and avoiding propellant waste caused by spacecraft descent. Thus, the problem of phasing control of the target spacecraft while maintaining orbit is solved, demonstrating good adaptability. This method can be applied to the long-term on-orbit operation of target spacecraft. Based on the requirements of orbit maintenance and phasing control, the launch window can be optimized to provide orbit raising control space for the phasing of the visiting spacecraft, effectively saving propellant and providing support for its long-term on-orbit operation.

[0221] The following is a detailed description with reference to another embodiment.

[0222] Example 2

[0223] The spacecraft rendezvous and docking device based on step-by-step orbit raising and phase adjustment control provided in this embodiment includes multiple implementation units, each of which corresponds to a specific implementation step in Embodiment 1 above.

[0224] Figure 2This is a schematic diagram of an optional spacecraft rendezvous and docking device based on step-by-step orbital ascent and phasing control according to an embodiment of the present invention, as shown below. Figure 2 As shown, the spacecraft rendezvous and docking device may include: a first determining unit 20, a second determining unit 21, a third determining unit 22, and a launching unit 23.

[0225] The first determining unit 20 is used to determine the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window based on a preset launch window range. The preset launch window range includes: multiple launch windows; the launch window is the date of launching the visiting spacecraft; and the visiting spacecraft is the spacecraft that will rendezvous and dock with the target spacecraft.

[0226] The second determining unit 21 is used to determine the phase error reference based on a determined date, wherein the determined date refers to the date used to determine the launch time of the visiting spacecraft; the phase error reference includes at least: prediction error and collision avoidance effect;

[0227] The third determining unit 22 is used to determine the target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference and the target phase difference, wherein the target phase difference is the phase difference when the visiting spacecraft rendezvous and docks with the target spacecraft.

[0228] Launch unit 23 is used to launch the visiting spacecraft within the target launch window based on a step-by-step orbital phasing control strategy. The step-by-step orbital phasing control strategy is used to control the orbital parameters of the target spacecraft so that the visiting spacecraft can rendezvous and dock with the target launch window.

[0229] The aforementioned spacecraft rendezvous and docking device can determine the coplanar phase difference between the target spacecraft and the visiting spacecraft for each launch window based on a preset launch window range. Then, based on a determined date, it calculates a phase error benchmark, which comprehensively considers the phase effects caused by prediction errors and collision avoidance control. Subsequently, by analyzing the coplanar phase difference, the phase error benchmark, and the target phase difference, it accurately selects the target launch window from the preset launch window range. This ensures that the visiting spacecraft launched under this window can efficiently complete phase control in an orbit-ascending mode, avoiding unnecessary orbit-descending operations of the target spacecraft, effectively saving propellant, and ensuring its long-term stable operation. This achieves the technical effect of improving the efficiency of orbit-ascending phase control, thereby solving the technical problem in related technologies where the launch window cannot be accurately determined in spacecraft rendezvous and docking missions, resulting in low phase control efficiency.

[0230] Optionally, the first determining unit includes: a first determining module for determining the coplanar time, wherein the coplanar time is the time of launching the visiting spacecraft; a second determining module for determining, for each launch window, a first phase of the visiting spacecraft at the coplanar time of the launch window and a second phase of the target spacecraft at the coplanar time of the launch window; and a third determining module for determining the coplanar phase difference corresponding to each launch window based on the first phase and the second phase.

[0231] Optionally, the first determining module includes: a first determining submodule, used to determine the right ascension of the first ascending node of the visiting spacecraft in the instantaneous true equatorial coordinate system; a first transformation submodule, used to transform the right ascension of the first ascending node to obtain the right ascension of the second ascending node of the visiting spacecraft in the inertial coordinate system; a second determining submodule, used to determine the virtual coplanarity condition, wherein the virtual coplanarity condition refers to the condition that the virtual target plane of the target spacecraft's orbit is coplanar with the orbital plane of the visiting spacecraft's entry into orbit, and the virtual target plane refers to the plane considering the drift difference of the ascending node; and a third determining submodule, used to determine the coplanarity time based on the virtual coplanarity condition and the right ascension of the second ascending node.

[0232] Optionally, the second determining unit includes: a fourth determining module, used to determine the preset range in which the determined date is located, and to determine the number of avoidance control operations based on the preset range; a fifth determining module, used to determine the collision avoidance impact amount during each avoidance control operation based on the target spacecraft's average number of orbits per day and the target spacecraft's operating speed; and a sixth determining module, used to determine the phase error benchmark based on the prediction error and the collision avoidance impact amount during each avoidance control operation.

[0233] Optionally, the spacecraft rendezvous and docking also includes: a first construction module, used to construct an atmospheric drag model before determining the phase error benchmark based on the prediction error and the collision avoidance impact amount during each avoidance control, wherein the atmospheric drag model includes at least: atmospheric density variable, atmospheric drag coefficient, equivalent area variable, and spacecraft motion vector; and a seventh determination module, used to determine the prediction error based on the magnitude of the spacecraft perturbation factor using the atmospheric drag model.

[0234] Optionally, the third determining unit includes: an eighth determining module, used to determine the sum of the phase error reference and the target phase difference to obtain a comparison phase difference; a first filtering module, used to filter candidate coplanar phase differences that are greater than the comparison phase difference from all coplanar phase differences to obtain a set of candidate coplanar phase differences; a ninth determining module, used to determine the absolute difference between each candidate coplanar phase difference and the comparison phase difference; and a tenth determining module, used to determine the candidate coplanar phase difference indicated by the smallest absolute difference as the target coplanar phase difference, and to determine the transmission window indicated by the target coplanar phase difference as the target transmission window.

[0235] Optionally, the spacecraft rendezvous and docking also includes: an eleventh determination module, used to determine the first relationship between orbital phase, orbital angular velocity, and flight time before launching the visiting spacecraft within the target launch window, based on a step-by-step orbital phasing control strategy, wherein the orbital phase is determined based on the orbital semi-major axis; a twelfth determination module, used to determine the second relationship between orbital phase change and orbital semi-major axis change based on the first relationship; a thirteenth determination module, used to determine the third relationship between orbital phase change and velocity increment based on the second relationship and the relationship between orbital semi-major axis change and velocity increment; and a second construction module, used to construct a phasing control model based on the first, second, and third relationships.

[0236] Optionally, the launch unit includes: a fourteenth determining module for determining multiple constraints of phase modulation control, wherein the multiple constraints of phase modulation control include at least: phase difference constraint, eccentricity constraint, and semi-major axis constraint; a fifteenth determining module for determining nonlinear programming variables based on the phase modulation control model, wherein the nonlinear programming variables include: multiple orbital parameters; and a first adjusting module for adjusting the orbital parameters based on the multiple constraints of phase modulation control, so that the phase difference between the visiting spacecraft and the target launch window during rendezvous and docking is equal to the target phase difference.

[0237] The aforementioned spacecraft rendezvous and docking device may also include a processor and a memory. The first determining unit 20, the second determining unit 21, the third determining unit 22, the launching unit 23, etc., are all stored in the memory as program units, and the processor executes the aforementioned program units stored in the memory to realize the corresponding functions.

[0238] The aforementioned processor contains a kernel, which retrieves the corresponding program units from memory. One or more kernels can be configured, and by adjusting kernel parameters, a step-by-step ascent and phasing control strategy can be used to launch the visiting spacecraft within the target launch window.

[0239] The aforementioned memory may include non-permanent memory in computer-readable media, such as random access memory (RAM) and / or non-volatile memory, such as read-only memory (ROM) or flash RAM, and the memory includes at least one memory chip.

[0240] The present invention also provides a computer program product, which, when executed on a data processing device, is suitable for executing an initialization program having the following method steps: determining the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window based on a preset launch window range; determining a phase error reference based on a determined date; determining a target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference, and the target phase difference; and launching the visiting spacecraft within the target launch window based on a step-by-step orbital phasing control strategy.

[0241] According to another aspect of the present invention, a computer program product is also provided, including a non-volatile computer-readable storage medium storing a computer program, wherein the computer program, when executed by a processor, implements the spacecraft rendezvous and docking method based on step-by-step orbit raising and phasing control as described above.

[0242] According to another aspect of the present invention, an electronic device is also provided, including one or more processors and a memory, wherein the memory is used to store one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors cause the one or more processors to implement the above-described spacecraft rendezvous and docking method based on step-by-step orbital ascent and phasing control.

[0243] Figure 3 This is a hardware structure block diagram of an electronic device (or mobile device) for a spacecraft rendezvous and docking method based on step-by-step orbital ascent and phasing control, according to an embodiment of the present invention. Figure 3 As shown, an electronic device may include one or more processors (e.g., Figure 3 The processors 302a, 302b, ..., 302n, etc., may include, but are not limited to, processing devices such as microprocessors (MCUs) or programmable logic devices (FPGAs), and a memory 304 for storing data. In addition, it may include: a display, an input / output interface (I / O interface), a universal serial bus (USB) port (which may be included as one of the ports in the I / O interface), a network interface, a keyboard, a power supply, and / or a camera. Those skilled in the art will understand that... Figure 3 The structure shown is for illustrative purposes only and does not limit the structure of the electronic device described above. For example, the electronic device may also include components that are more... Figure 3 The more or fewer components shown, or having the same Figure 3 The different configurations shown.

[0244] The sequence numbers of the above embodiments of the present invention are for descriptive purposes only and do not represent the superiority or inferiority of the embodiments.

[0245] The embodiments or examples disclosed herein are not exhaustive, but merely illustrative of some embodiments or examples, and are not intended to limit the scope of protection of this disclosure. Unless otherwise specified, each step in a particular embodiment or example can be implemented as an independent embodiment, and the steps can be arbitrarily combined. For example, a solution after removing some steps in a particular embodiment or example can also be implemented as an independent embodiment, and the order of the steps in a particular embodiment or example can be arbitrarily interchanged. Furthermore, optional methods or examples in a particular embodiment or example can be arbitrarily combined; moreover, embodiments or examples can be arbitrarily combined. For example, some or all steps of different embodiments or examples can be arbitrarily combined, and a particular embodiment or example can be arbitrarily combined with optional methods or examples of other embodiments or examples.

[0246] In the above embodiments of the present invention, the descriptions of each embodiment have different focuses. For parts not described in detail in a certain embodiment, please refer to the relevant descriptions of other embodiments.

[0247] In the several embodiments provided by this invention, it should be understood that the disclosed technical content can be implemented in other ways. The device embodiments described above are merely illustrative; for example, the division of units can be a logical functional division, and in actual implementation, there may be other division methods. For example, multiple units or components may be combined or integrated into another system, or some features may be ignored or not executed. Furthermore, the displayed or discussed mutual coupling, direct coupling, or communication connection can be through some interfaces; the indirect coupling or communication connection of units or modules can be electrical or other forms.

[0248] The units described as separate components may or may not be physically separate. The components shown as units may or may not be physical units; that is, they may be located in one place or distributed across multiple units. Some or all of the units can be selected to achieve the purpose of this embodiment according to actual needs.

[0249] Furthermore, the functional units in the various embodiments of the present invention 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.

[0250] If the integrated unit is implemented as a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present invention, in essence, or the part that contributes to the prior art, or all or part of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute all or part of the steps of the methods described in the various embodiments of the present invention. The aforementioned storage medium includes various media capable of storing program code, such as USB flash drives, read-only memory (ROM), random access memory (RAM), portable hard drives, magnetic disks, or optical disks.

[0251] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A spacecraft rendezvous and docking method based on step-by-step orbit raising and phasing control, characterized in that, include: Based on a preset launch window range, the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window is determined. The preset launch window range includes: multiple launch windows; the launch window is the date of launching the visiting spacecraft; and the visiting spacecraft is the spacecraft that will rendezvous and dock with the target spacecraft. A phase error benchmark is determined based on a specific date, wherein the specific date refers to the date used to determine the launch time of the visiting spacecraft; the phase error benchmark includes at least: prediction error and collision avoidance impact. Based on the coplanar phase difference, the phase error reference, and the target phase difference, the target launch window is determined from the preset launch window range, wherein the target phase difference is the phase difference when the visiting spacecraft rendezvous and docks with the target spacecraft; Based on a step-by-step orbital ascent and phasing control strategy, the visiting spacecraft is launched within the target launch window. The step-by-step orbital ascent and phasing control strategy is used to control the orbital parameters of the target spacecraft so that the visiting spacecraft can rendezvous and dock with the target launch window.

2. The spacecraft rendezvous and docking method according to claim 1, characterized in that, The steps for determining the coplanar phase difference between the target spacecraft and the visiting spacecraft for each launch window, based on a preset launch window range, include: Determine the coplanar moment, wherein the coplanar moment is the moment when the visiting spacecraft is launched; For each launch window, a first phase of the visiting spacecraft at the coplanar time of the launch window and a second phase of the target spacecraft at the coplanar time of the launch window are determined. Based on the first phase and the second phase, the coplanar phase difference corresponding to each of the transmission windows is determined.

3. The spacecraft rendezvous and docking method according to claim 2, characterized in that, The steps to determine the coplanar moment include: Determine the right ascension of the first ascending node of the visiting spacecraft in the instantaneous true equatorial coordinate system; By transforming the right ascension of the first ascending node, the right ascension of the second ascending node of the visiting spacecraft in the inertial coordinate system is obtained; Determine the virtual coplanarity condition, wherein the virtual coplanarity condition refers to the condition that the virtual target plane of the target spacecraft's orbit is coplanar with the orbital plane of the visiting spacecraft's entry into orbit, and the virtual target plane refers to the plane that takes into account the difference in ascending node drift. The coplanarity time is determined based on the virtual coplanarity condition and the right ascension of the second ascending node.

4. The spacecraft rendezvous and docking method according to claim 1, characterized in that, The steps for determining the phase error reference based on a given date include: Determine a preset range within which the determined date falls, and based on the preset range, determine the number of times to circumvent control. Based on the target spacecraft’s average number of orbits per day and the target spacecraft’s operating speed, the collision avoidance impact amount is determined for each avoidance control operation. The phase error benchmark is determined based on the prediction error and the collision avoidance impact amount during each avoidance control.

5. The spacecraft rendezvous and docking method according to claim 4, characterized in that, Before determining the phase error reference based on the prediction error and the collision avoidance impact amount at each avoidance control, the method further includes: Construct an atmospheric drag model, wherein the atmospheric drag model includes at least: atmospheric density variable, atmospheric drag coefficient, equivalent area variable, and spacecraft motion vector; The prediction error is determined using the atmospheric drag model based on the magnitude of spacecraft perturbation factors.

6. The spacecraft rendezvous and docking method according to claim 1, characterized in that, The step of determining the target transmission window from the preset transmission window range based on the coplanar phase difference, the phase error reference, and the target phase difference includes: The sum of the phase error reference and the target phase difference is determined to obtain the comparison phase difference; From all the said coplanar phase differences, candidate coplanar phase differences that are greater than the compared phase difference are selected to obtain a set of candidate coplanar phase differences; Determine the absolute difference between each of the candidate coplanar phase differences and the compared phase differences; The candidate coplanar phase difference indicated by the minimum absolute difference is determined as the target coplanar phase difference, and the transmission window indicated by the target coplanar phase difference is determined as the target transmission window.

7. The spacecraft rendezvous and docking method according to claim 1, characterized in that, Prior to launching the visiting spacecraft within the target launch window, based on a step-by-step orbital ascent and phasing control strategy, the following is also included: A first relationship is established between orbital phase and orbital angular velocity and time of flight, wherein the orbital phase is determined based on the orbital semi-major axis; Based on the first relationship, a second relationship between orbital phase change and orbital semi-major axis change is determined; Based on the second relationship and the relationship between the orbital semi-major axis change and the velocity increment, a third relationship between the orbital phase change and the velocity increment is determined. A phase modulation control model is constructed based on the first relationship, the second relationship, and the third relationship.

8. The spacecraft rendezvous and docking method according to claim 7, characterized in that, Based on a step-by-step orbital ascent and phasing control strategy, the steps for launching the visiting spacecraft within the target launch window include: Determine multiple constraints for phase modulation control, wherein the multiple constraints for phase modulation control include at least: phase difference constraint, eccentricity constraint, and semi-major axis constraint. Based on the phase modulation control model, nonlinear programming variables are determined, wherein the nonlinear programming variables include: multiple track parameters; Based on the phase control multiple constraints, the orbital parameters are adjusted so that the phase difference between the visiting spacecraft and the target launch window during rendezvous and docking is equal to the target phase difference.

9. A spacecraft rendezvous and docking device based on step-by-step orbit raising and phasing control, characterized in that, include: The first determining unit is configured to determine the coplanar phase difference between the target spacecraft and the visiting spacecraft corresponding to each launch window based on a preset launch window range, wherein the preset launch window range includes: multiple launch windows; the launch window is the date of launching the visiting spacecraft; and the visiting spacecraft is the spacecraft that will rendezvous and dock with the target spacecraft. The second determining unit is used to determine a phase error reference based on a determined date, wherein the determined date refers to the date used to determine the launch time of the visiting spacecraft; the phase error reference includes at least: prediction error and collision avoidance effect; The third determining unit is used to determine the target launch window from the preset launch window range based on the coplanar phase difference, the phase error reference and the target phase difference, wherein the target phase difference is the phase difference when the visiting spacecraft rendezvous and docks with the target spacecraft; The launch unit is used to launch the visiting spacecraft in the target launch window based on a step-by-step orbital ascent and phasing control strategy, wherein the step-by-step orbital ascent and phasing control strategy is used to control the orbital parameters of the target spacecraft so that the visiting spacecraft can rendezvous and dock with the target launch window.

10. An electronic device, characterized in that, It includes one or more processors and a memory, the memory being used to store one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors cause the one or more processors to implement the spacecraft rendezvous and docking method based on step-by-step orbital ascent and phasing control as described in any one of claims 1 to 8.

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