Method and apparatus for synchronizing a spacecraft with a target body

Abstract

The application provides a method and device for synchronizing a spacecraft with a target celestial body, and belongs to the technical field of aerospace. The method takes the condition that the orbit plane of the spacecraft is parallel to the radiation plane of the target celestial body at a preset time as a precondition, determines the precession angle of the orbit plane of the spacecraft as the angle required for changing the orbit plane of the spacecraft from a source orbit plane to a target orbit plane, determines a reference position of orbit plane adjustment on the current orbit of the spacecraft, and then adjusts the current orbit plane of the spacecraft to the target orbit plane based on the precession angle of the source orbit plane and the reference position. The method can change the orbit plane of the spacecraft, so that the spacecraft is always synchronized with the target celestial body, thereby ensuring that the spacecraft successfully performs a detection task.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, and in particular to a method and device for synchronizing a spacecraft with a target celestial body. BACKGROUND

[0002] High-speed charged particles in the solar atmosphere form the solar wind. When the solar wind reaches the Earth, the Earth's magnetic field is impeded by the solar wind particles, forming a magnetosphere around the Earth. Detecting the solar wind and the magnetosphere helps to predict space weather, understand solar activity, and the cosmic environment, etc.

[0003] In the process of detecting the solar wind or the Earth's magnetosphere by a detection satellite orbiting the Earth, it is necessary to ensure that the detection satellite is in a stationary state relative to the Earth and is synchronized with the sun. Synchronization with the sun means that the detection surface of the detection satellite always falls within the illumination area of the sun, that is, the orbital plane of the satellite must always be directed towards the sun.

[0004] Currently, to ensure that the satellite is synchronized with the sun, the attitude of the detection satellite can be adjusted, for example, the detection direction of the satellite can be adjusted so that the orbital plane of the satellite is always directed towards the sun. However, when the satellite runs into the Earth's shadow area (referred to as the shadow area, which is the area on the Earth where the sunlight is blocked), the satellite is no longer synchronized with the sun, that is, when the satellite is in the shadow area, the satellite cannot perform the detection task. SUMMARY

[0005] The present application provides a method and device for synchronizing a spacecraft with a target celestial body, by changing the orbital plane of the spacecraft, so that the spacecraft is always synchronized with the target celestial body, to ensure that the spacecraft can successfully perform the detection task.

[0006] To achieve the above purpose, the present application adopts the following technical solutions:

[0007] In a first aspect, the present application provides a method for synchronizing a spacecraft with a target celestial body, the spacecraft orbits a central celestial body, and the central celestial body orbits the target celestial body, the method comprising: determining the precession angle of the source orbital plane of the spacecraft at a predetermined time when the spacecraft is synchronized with the target celestial body. The spacecraft is synchronized with the target celestial body when the orbital plane of the spacecraft is parallel to the radiation plane of the target celestial body; the orbital plane after the source orbital plane is rotated by the precession angle is the target orbital plane. On the current orbit of the spacecraft, a reference position for performing the orbital plane adjustment is determined. Based on the reference position and the precession angle of the source orbital plane of the spacecraft, the motion state of the spacecraft is adjusted to adjust the current orbital plane of the spacecraft to the target orbital plane.

[0008] The method for synchronizing the spacecraft with the target celestial body can adjust the current orbit plane of the spacecraft to be parallel to the radiation plane of the target celestial body, so that the orbit plane of the spacecraft is always in the radiation area of the target celestial body, that is, the spacecraft can be always synchronized with the target celestial body, so as to ensure that the spacecraft successfully performs a detection task.

[0009] In an implementation form of the first aspect, the reference position for performing the orbit plane adjustment is determined on the current orbit of the spacecraft, comprising: determining the ascending node and the descending node of the current orbit plane of the spacecraft and the target orbit plane. The reference position is determined as the point closest to the apogee of the current orbit among the ascending node and the descending node.

[0010] In an implementation form of the first aspect, the method further comprises: optimizing the precession angle of the source orbit plane of the spacecraft.

[0011] In an implementation form of the first aspect, the method further comprises: generating a w-e phase diagram according to the multiple groups of candidate orbit elements of the spacecraft; wherein w is the argument of perigee, and e is the eccentricity; the w-e phase diagram is a curve diagram for describing the change relationship between w and e. The eccentricity corresponding to a target curve in the w-e phase diagram is determined as the eccentricity of the source orbit of the spacecraft; the target curve is a curve with the minimum range of eccentricity in the w-e phase diagram.

[0012] In an implementation form of the first aspect, the motion state of the spacecraft is adjusted based on the reference position and the precession angle of the source orbit plane of the spacecraft, comprising: determining the velocity increment of the spacecraft based on the precession angle of the source orbit plane of the spacecraft. When the spacecraft runs to the reference position on the current orbit, an impulse is applied to the spacecraft based on the velocity increment, so as to adjust the current orbit plane of the spacecraft to the target orbit plane.

[0013] In an implementation form of the first aspect, the motion state of the spacecraft is adjusted based on the reference position and the precession angle of the source orbit plane of the spacecraft, comprising: determining the start position and the end position of the motion state adjustment based on the reference position and the precession angle of the source orbit plane of the spacecraft. When the spacecraft runs to the start position on the current orbit, a continuous thrust is applied to the spacecraft until the spacecraft runs to the end position, so as to adjust the current orbit plane of the spacecraft to the target orbit plane.

[0014] In a second aspect, the present application provides a device for synchronizing a spacecraft with a target celestial body, comprising a precession angle determination module, a reference position determination module, and an orbital plane adjustment module. The precession angle determination module is configured to determine a precession angle of a source orbital plane of the spacecraft at a preset time when the spacecraft is synchronized with the target celestial body. The spacecraft is synchronized with the target celestial body when an orbital plane of the spacecraft is parallel to a radiation plane of the target celestial body. The target orbital plane is the orbital plane after the source orbital plane is rotated by the precession angle. The reference position determination module is configured to determine a reference position for performing the orbital plane adjustment on a current orbit of the spacecraft. The orbital plane adjustment module is configured to adjust a motion state of the spacecraft based on the reference position and the precession angle of the source orbital plane of the spacecraft, so as to adjust the current orbital plane of the spacecraft to the target orbital plane.

[0015] In an implementation form of the second aspect, the reference position determination module is specifically configured to determine ascending nodes and descending nodes of the current orbital plane of the spacecraft and the target orbital plane. The reference position is determined as a node closest to an apogee of the current orbit among the ascending nodes and the descending nodes.

[0016] In an implementation form of the second aspect, the orbital plane adjustment module is specifically configured to determine a velocity increment of the spacecraft based on the precession angle of the source orbital plane of the spacecraft. When the spacecraft runs to the reference position on the current orbit, the orbital plane adjustment module is configured to apply a pulse to the spacecraft based on the velocity increment, so as to adjust the current orbital plane of the spacecraft to the target orbital plane.

[0017] In an implementation form of the second aspect, the device further comprises an orbit determination module. The orbit determination module is configured to determine an orbit of the spacecraft around a central celestial body.

[0018] In an implementation form of the second aspect, the orbit determination module is specifically configured to generate a w-e phase diagram according to a plurality of groups of candidate orbital elements. e is an eccentricity, and w is an argument of perigee. The w-e phase diagram is a curve diagram for describing a change relationship between e and w. An eccentricity corresponding to a target curve in the w-e phase diagram is determined as the eccentricity of the source orbit of the spacecraft. The target curve is a curve with a minimum range of eccentricities in the w-e phase diagram.

[0019] In an implementation form of the second aspect, the device further comprises an optimization module. The optimization module is configured to optimize the precession angle of the source orbital plane of the spacecraft.

[0020] In an implementation form of the first aspect and the second aspect, the precession angle of the source orbital plane of the spacecraft satisfies:

[0021] α t0 =ω c t0

[0022] wherein t0 represents the preset time, and α represents the precession angle of the source orbital plane of the spacecraft. t0represents the precession angle of the source orbit plane of the spacecraft from the initial time to the time t0, ω c represents the angular velocity of the central celestial body revolving around the target celestial body.

[0023] In an implementation form of the first and second aspects, the orbit of the spacecraft revolving around the central celestial body is a frozen orbit.

[0024] In an implementation form of the first and second aspects, the target celestial body is the sun, and the central celestial body is the earth.

[0025] In a third aspect, the present application provides an electronic device, comprising a processor and a memory coupled to the processor; the memory is configured to store computer instructions, when the electronic device is running, the processor executes the computer instructions stored in the memory, so that the electronic device executes the method as described in the first aspect or any implementation form thereof.

[0026] In a fourth aspect, the present application provides a computer readable storage medium, comprising computer program instructions, when the computer program instructions are executed by a computer, the computer executes the method as described in the first aspect or any implementation form thereof.

[0027] In a fifth aspect, the present application provides a computer program product, comprising computer program instructions, when the computer program instructions are executed on a computer, the computer executes the method as described in the first aspect or any implementation form thereof.

[0028] The technical effects of the second to fifth aspects and their possible implementation forms correspond to the description of the technical effects of the first aspect and its possible implementation forms, which will not be repeated here. BRIEF DESCRIPTION OF DRAWINGS

[0029] Figure 1 is an orbit root number diagram provided by the embodiments of the present application;

[0030] Figure 2 is a satellite orbit diagram provided by the embodiments of the present application;

[0031] Figure 3 is one of the method diagrams for synchronizing a spacecraft with a target celestial body provided by the embodiments of the present application;

[0032] Figure 4 is a w-e phase diagram provided by the embodiments of the present application;

[0033] Figure 5 is another method diagram for synchronizing a spacecraft with a target celestial body provided by the embodiments of the present application;

[0034] Figure 6is a schematic diagram of a node coordinate system S provided by an embodiment of the present application n and S Tar .

[0035] Figure 7 is a schematic diagram of a method for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0036] Figure 8 is a schematic diagram of a position relationship between a current orbit plane of a satellite and a target orbit plane provided by an embodiment of the present application;

[0037] Figure 9 is a schematic diagram of a method for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0038] Figure 10 is a simulation result diagram of a method for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0039] Figure 11 is a schematic diagram of a method for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0040] Figure 12 is a simulation result diagram of a method for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0041] Figure 13 is a schematic diagram of a device for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0042] Figure 14 is a schematic diagram of a device for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application;

[0043] Figure 15 is a schematic diagram of a device for synchronizing a spacecraft with a target celestial body provided by an embodiment of the present application. DETAILED DESCRIPTION

[0044] In the specification and claims of the present application, the terms "first" and "second" are used to distinguish different objects, and are not used to describe a specific order of the objects.

[0045] In the embodiments of the present application, "and / or" represents a relationship between objects, for example, A and / or B can represent the following three cases: A exists alone, B exists alone, and A and B exist together.

[0046] In the embodiments of the present application, the word "exemplary" or "for example" is used to mean serving as an example, instance, or illustration. Any embodiment or design described in this application as "exemplary" or "for example" should not be construed as being preferred or advantageous over other embodiments or designs. Rather, use of the words "exemplary" or "for example" is intended to present concepts in a concrete manner.

[0047] In the description of the present application, unless otherwise specified, the meaning of "multiple groups" is two groups or more. For example, the number of multiple groups of candidate orbit elements refers to two groups or more of candidate orbit elements.

[0048] The method and device provided in the embodiments of the present application are used to adjust the motion state of a spacecraft, thereby changing the orbital plane of the spacecraft, so that the spacecraft is synchronized with a target celestial body. In other words, the spacecraft is synchronized with the target celestial body when the orbital plane of the spacecraft is parallel to the radiation plane of the target celestial body, that is, the detection plane of the spacecraft always falls into the radiation area of the target celestial body, that is, the orbital plane of the spacecraft always faces the radiation plane of the target celestial body.

[0049] It should be understood that, in the embodiments of the present application, the spacecraft (such as a satellite or the like) moves around a central celestial body (such as a celestial body, for example, the Earth), the central celestial body moves around a target celestial body (such as a celestial body, for example, the sun, the moon, a nebula, or the like), and the orbit of the spacecraft moving around the central celestial body is a frozen orbit.

[0050] The technical terms involved in the embodiments of the present application will be introduced first as follows.

[0051] 1. Orbit element

[0052] The orbit element can be simply understood as some parameters of the running orbit of the spacecraft, for example, the parameters of the running orbit of the satellite moving around the central celestial body, the Earth. The orbit element can be used to describe the motion state of the spacecraft in space. For example, for a satellite, through the orbit element of the satellite, the position, orbit shape, motion speed, and the like of the satellite can be accurately described.

[0053] In the embodiments of the present application, each group of candidate orbit elements can be (a, e, i, Ω, ω, θ), where a represents the semi-major axis, e represents the eccentricity, i represents the orbit inclination, θ represents the true anomaly, Ω represents the right ascension of the ascending node, and ω represents the argument of perigee.

[0054] The semi-major axis a refers to half of the major axis of the satellite orbit, and the semi-major axis is used to describe the size of the satellite orbit.

[0055] The eccentricity e refers to the eccentricity of the satellite orbit, which is used to measure the degree of deviation of the orbit.

[0056] The orbit inclination i refers to the dihedral angle between the satellite orbit and the Earth equatorial plane (e.g. the orbit inclination i in FIG. 1). Figure 1

[0057] The true anomaly θ refers to the angle swept by the satellite's radial (the straight line connecting the Earth and the satellite) when the satellite moves along the orbit from the perigee (the point on the satellite orbit closest to the Earth).

[0058] Referring to FIG. 1, Figure 1 the right ascension of the ascending node Ω refers to the angle in the Earth equatorial plane from the vernal equinox V to the ascending node B (the point where the satellite orbit crosses the equator from south to north).

[0059] Continuing to refer to FIG. 1, Figure 1 the argument of the perigee ω refers to the angle in the satellite orbit plane from the ascending node B to the perigee P.

[0060] The orbit semi-major axis a and the eccentricity e are used to describe the size and shape of the orbit; the orbit inclination i, the right ascension of the ascending node Ω, and the argument of the perigee ω are used to describe the position of the satellite orbit; and the true anomaly θ is used to describe the specific position of the satellite on the orbit.

[0061] 2. Frozen orbit

[0062] A frozen orbit is an orbit in which the arch line of the orbit is basically unchanged. The arch line is a straight line connecting two apses on the orbit, and the long axis of the orbit is the arch line. When the arch line is basically unchanged, the orbit of the satellite is basically unchanged. Generally, a frozen orbit can maintain stable performance in a long-term evolution process. Specifically, when the orbit semi-major axis a, the eccentricity e, the orbit inclination i, and the argument of the perigee ω are basically unchanged, the arch line is basically unchanged.

[0063] In some cases, the satellite needs to keep the spatial position of the satellite relative to the Earth static in order to successfully perform a solar wind detection task or a geomagnetic layer detection task (it should be understood that the satellite performs the detection task by moving around the Earth). When the orbit of the satellite moving around the Earth does not change (i.e., the orbit of the satellite is a frozen orbit), the satellite can keep relative static with the Earth.

[0064] The solar wind refers to the flow of high-speed charged particles in the solar atmosphere, mainly composed of plasma ejected from the solar corona. The geomagnetic layer refers to the cavity structure surrounding the Earth formed when the Earth's magnetic field is hindered by solar wind particles when the solar wind reaches the Earth.

[0065] In practical applications, for example, the satellite rotates around the Earth, the apogee height r a The apogee height r a may be set according to the height of the magnetic layer top of the geomagnetic field, and then the orbit semi-major axis a is calculated by the formula a = r​a The orbit semi-major axis a can be determined according to the formula (1+e). The orbit inclination i can be set according to experience, for example, set near 63.43° (i.e. a common critical inclination). The pericenter amplitude ω can also be set according to experience, for example, set as 90° or 270°. The eccentricity e cannot be directly set according to experience and / or the type of the detection task performed by the satellite.

[0066] Therefore, in the case that the orbit semi-major axis a, the orbit inclination i and the pericenter amplitude ω are pre-set, only the orbit eccentricity e that makes the orbit of the satellite stable needs to be further determined, so as to design the frozen orbit.

[0067] In the embodiments of the present application, in the process of performing the solar wind detection task or the geomagnetic layer detection task, the satellite not only needs to keep the spatial position relative to the earth static through the frozen orbit, but also needs to keep synchronization with the sun (i.e. the detection surface of the spacecraft always falls into the radiation region of the target celestial body) in the process of performing the above detection task.

[0068] In the above case, in order to realize the detection of the solar wind or the geomagnetic layer, it is required that Figure 2 the apogee of the satellite orbit is higher than the magnetopause of the geomagnetic field, and the perigee of the satellite is located inside the geomagnetic layer. However, when the satellite is in the earth shadow area (i.e. the area on the earth where the sunlight is blocked) in the process of moving around the earth, Figure 2 the detection surface of the satellite cannot be irradiated by the sun, which will result in that the satellite cannot detect the solar wind or the geomagnetic layer.

[0069] In order to solve the above problem, the embodiments of the present application provide a method and device for keeping a spacecraft synchronized with a target celestial body, which takes the orbit plane of the spacecraft parallel to the radiation plane of the target celestial body at a preset time as a prerequisite condition (the prerequisite condition is the condition for keeping the spacecraft synchronized with the target celestial body), determines the precession angle of the orbit plane of the spacecraft as the angle required for the orbit plane of the spacecraft to change from a source orbit plane to a target orbit plane (i.e. the radiation plane of the target celestial body at the preset time), determines a reference position of the orbit plane adjustment on the current orbit of the spacecraft, and adjusts the current orbit plane of the spacecraft to the target orbit plane based on the precession angle of the source orbit plane and the reference position. It can be seen that the method can adjust the current orbit plane of the spacecraft to be parallel to the radiation plane of the target celestial body, so that the orbit plane of the spacecraft is always located in the radiation region of the target celestial body, i.e. the spacecraft can always be synchronized with the target celestial body, so as to ensure that the spacecraft successfully performs the detection task.

[0070] Exemplarily, the method for synchronizing the spacecraft with the target celestial body provided by the embodiments of the present application can be executed by an electronic device with processing function, for example, the electronic device can be a computer, a server, etc. Taking the computer as an example, the hardware part of the computer can include a processor, a memory, a network interface, a user interface, a communication bus, etc.

[0071] Exemplarily, the processor is configured to determine the precession angle of the source orbital plane of the spacecraft, the reference position of the orbital plane adjustment, and adjust the motion state of the spacecraft. The processor can include a central processing unit (CPU) or other processors, and the processor can be single-core or multi-core, for example, the processor can include multiple CPUs.

[0072] The memory is configured to store computer instructions and related data, for example, to store the precession angle of the source orbital plane of the spacecraft, the reference position, the velocity increment, the start position and the end position of the motion state adjustment, etc. The memory can be a random access memory (RAM), a read only memory (ROM), an erasable programmable read only memory (EPROM), a flash memory, or an optical memory, a magnetic disk storage medium or other magnetic storage device, or any other medium capable of storing program codes or data accessible by a computer. Optionally, the memory can be integrated in the processor, and the memory can also be independent of the processor.

[0073] The network interface is configured to enable the computer to communicate with other devices or communication networks, and the network interface can be a transceiver with transceiving function. Optionally, the network interface can include a standard wired interface, a wireless interface (such as a WI-FI interface, a Bluetooth interface, a 5G interface).

[0074] The communication bus is configured to enable the connection and communication between different components, for example, the above-mentioned processor, memory, network interface and user interface can be interconnected through the communication bus.

[0075] The user interface can include a display screen and an input unit (such as a keyboard), and optionally, the user interface can also include a standard wired interface and a wireless interface.

[0076] Those skilled in the art can understand that the above-mentioned computer can also include more or fewer components, or combine certain components, or different component arrangements, which are not limited in the embodiments of the present application.

[0077] The execution process of the method for synchronizing the spacecraft with the target celestial body provided in the embodiments of the present application can include two stages. In the first stage, the source orbit (i.e., the frozen orbit) of the spacecraft moving around the central celestial body is designed. In the second stage, the orbital plane of the orbit is adjusted so that the spacecraft is synchronized with the target celestial body during the movement of the spacecraft around the central celestial body on the source orbit.

[0078] Optionally, the target celestial body in the embodiments of the present application is the sun, and the central celestial body is the earth. Thus, one application scenario of the method for synchronizing the spacecraft with the target celestial body provided in the embodiments of the present application is that the spacecraft is synchronized with the sun during the movement of the spacecraft around the earth, so as to successfully perform the task of detecting the solar wind or the geomagnetic layer.

[0079] In combination with Figure 3 The first stage is described in detail as follows. The method for determining the frozen orbit of the spacecraft includes S1-S2.

[0080] S1, generating a w-e phase diagram according to a plurality of sets of candidate orbit elements of the spacecraft.

[0081] In the w-e phase diagram, w is the argument of perigee, and e is the eccentricity. The w-e phase diagram is a curve diagram for describing the change relationship between w and e.

[0082] Each set of candidate orbit elements includes an orbit semi-major axis a, an eccentricity e, an orbit inclination i, a true anomaly θ, a right ascension of the ascending node Ω, and an argument of pericenter ω.

[0083] It should be understood that each set of candidate orbit elements corresponds to a satellite orbit, and each set of candidate orbit elements also corresponds to a curve in the w-e phase diagram for describing the change relationship between e and w. That is, the w-e phase diagram includes a plurality of curves, and each curve is used to describe the change relationship between w and e in the orbit elements of a satellite orbit.

[0084] Optionally, S1 includes S11-S12.

[0085] S11, for each set of candidate orbit elements of the plurality of sets of candidate orbit elements, generating orbit elements at a plurality of time points starting from each set of candidate orbit elements.

[0086] In one implementation, a plurality of eccentricities are selected as candidate eccentricities (also referred to as initial eccentricities) from an eccentricity set in a given eccentricity value range, and each candidate eccentricity is combined with preset other orbit elements (i.e., the orbit semi-major axis a, the orbit inclination i, the true anomaly θ, the right ascension of the ascending node Ω, and the argument of pericenter ω) to obtain a plurality of sets of candidate orbit elements. The candidate orbit elements are a kind of average value measurement.

[0087] Optionally, the average orbit element is converted into the instantaneous orbit element by the mean-to-instant conversion on each of the plurality of sets of candidate orbit elements, to obtain a plurality of sets of instantaneous candidate orbit elements; then, the plurality of sets of instantaneous candidate orbit elements are input into a simulation software (for example, a satellite simulation software Satellite Tool Kit, STK for short), and the integrator in the simulation software is used to generate the instantaneous orbit elements at a plurality of time points within a period of time starting from each set of candidate orbit elements, that is, to obtain the orbit elements at the plurality of time points starting from each set of candidate orbit elements.

[0088] For example, the simulation software used to generate the orbit elements at the plurality of time points can be the satellite simulation software STK, and the integrator in the satellite simulation software STK can be a high-precision orbit propagator (HPOP) or a long orbit propagator (LOP). Of course, the simulation software used to generate the orbit elements at the plurality of time points can also be other simulation software with similar functions.

[0089] Optionally, the period of time starting from each set of candidate orbit elements can be 5 years or 8 years, which is specifically set according to actual needs, and the embodiment of the present application does not make further limitation on the period of time.

[0090] S12, generating a curve corresponding to each set of candidate orbit elements in the w-e phase diagram according to the orbit elements at the plurality of time points, to obtain the w-e phase diagram.

[0091] For example, the process of drawing a curve in the w-e phase diagram includes:

[0092] First, the instantaneous orbit elements at the plurality of time points are converted into the mean orbit elements at the plurality of time points.

[0093] Secondly, the eccentricity e and the pericenter argument ω in the mean orbit elements at the plurality of time points are obtained.

[0094] In the embodiment of the present application, the w-e phase diagram is a diagram describing the transformation relationship between e and ω, which is drawn in a coordinate system with the eccentricity e as the vertical coordinate and the pericenter argument ω as the horizontal coordinate. The pericenter argument ω and the eccentricity e at one time point calculated above are taken as a point (ω, e) in the coordinate system, for example, a point in the coordinate system is denoted as wherein j represents the number of candidate orbit roots of the jth group, and i represents the ith time point in a period of time starting from the number of candidate orbit roots of the jth group. The multiple time points correspond to multiple points in the coordinate system, and the multiple points can draw a curve in the w-e phase diagram. A schematic diagram of the w-e phase diagram is shown in Figure 4

[0095] S2, determining the eccentricity corresponding to the target curve in the w-e phase diagram as the eccentricity of the source orbit of the spacecraft, the target curve being the curve with the minimum range of eccentricity in the w-e phase diagram.

[0096] It can be understood that the smaller the change in eccentricity e (i.e., the smaller the fluctuation in eccentricity), the better the stability of the satellite orbit described by the curve in the w-e phase diagram, and the curve with the minimum change in eccentricity e in the w-e phase diagram can be used as the frozen orbit. Therefore, in the w-e phase diagram obtained in S1, the curve with the minimum range of eccentricity (i.e., the difference between the maximum and minimum values of eccentricity) is taken as the target curve, and the eccentricity corresponding to the candidate orbit root in the target curve is taken as the eccentricity of the frozen orbit of the spacecraft.

[0097] After the eccentricity of the frozen orbit is determined, the eccentricity is combined with the preset other orbit roots (i.e., the orbit semi-major axis a, the orbit inclination i, the true anomaly θ, the ascending node right ascension Ω, and the pericenter amplitude angle ω) to obtain the orbit roots of the source orbit of the spacecraft.

[0098] In the second phase, as shown in Figure 5 the spacecraft is based on the orbit roots designed above, and to synchronize the spacecraft with the target celestial body, the method for synchronizing the spacecraft with the target celestial body provided by the present application comprises S101-S103.

[0099] S101, determining the precession angle of the source orbit plane of the spacecraft when the spacecraft is synchronized with the target celestial body at a preset time.

[0100] It should be understood that the spacecraft is synchronized with the target celestial body when the orbit plane of the spacecraft is parallel to the radiation plane of the target celestial body (i.e., the orbit plane of the spacecraft is always directed towards the radiation plane of the target celestial body); the orbit plane of the spacecraft after the source orbit plane is rotated by the precession angle is the target orbit plane, that is, the orbit plane of the spacecraft is rotated to the target orbit plane by the precession angle, so that the spacecraft can be synchronized with the target celestial body.

[0101] ​In this embodiment, the orbital plane of the spacecraft's source orbit (i.e., frozen orbit) designed in the first stage is used as the spacecraft's source orbital plane. The orbital plane obtained by rotating the spacecraft's source orbital plane around the angular momentum vector of the central celestial body orbiting the target celestial body as the axis of rotation and the angular velocity of the central celestial body orbiting the target celestial body as the angular velocity of rotation is used as the radiation surface of the target celestial body.

[0102] Given a preset time, the angle of rotation of the radiating surface of the target celestial body from the initial time to the preset time is determined as the precession angle of the spacecraft's source orbital plane when the spacecraft is synchronized with the target celestial body. At the preset time, the spacecraft's orbital plane is called the spacecraft's target orbital plane.

[0103] Optionally, the duration from the current time to the preset time can be one orbital period of the spacecraft or two orbital periods, and this application embodiment does not limit it.

[0104] To ensure that the spacecraft is synchronized with the target celestial body, the precession angle of the spacecraft's source orbital plane should satisfy the following:

[0105]

[0106] Where t0 represents the preset time. ω represents the precession angle of the spacecraft's source orbital plane from the initial moment to time t0. c It represents the angular velocity of the central celestial body revolving around the target celestial body.

[0107] In one implementation, taking a spacecraft as a satellite, the central celestial body as Earth, and the target celestial body as the Sun as an example, the orbital elements of the satellite's source orbit are simulated to simulate the satellite's motion in the source orbit, and a nodal coordinate system S is established to simulate the source orbit plane of the satellite. n And the nodal coordinate system S used to simulate the sunlit surface Tar .

[0108] like Figure 6 As shown, the satellite in the simulation experiment is referred to as the reference satellite, and the center of the reference satellite is taken as the nodal coordinate system S. n The origin (point O) n ), point O n The direction pointing towards the perigee of the satellite's source orbit is used as the nodal coordinate system S. n x n The positive direction of the axis, z n The positive direction of the axis is consistent with the direction of the normal vector to the satellite's source orbital plane. n The positive direction of the axis, x n The positive direction of the axis and z n The positive directions of the axes form a right-handed coordinate system. The nodal coordinate system S n xn -y n The plane is the plane where the satellite's source orbital plane is located.

[0109] Continue to refer to Figure 6 With the Earth's orbital angular momentum vector H sun With the axis of rotation as the angular velocity ω of the Earth's revolution around the sun. sun Rotate the above nodal coordinate system S n Then the nodal coordinate system S can be obtained. Tar Node coordinate system S Tar The three axes are x Tar axis, y Tar axis and z Tar Axis, nodal coordinate system S Tar The origin (point O) Tar ) and coordinate system S n The origin (point O) n The same, the nodal coordinate system S Tar x Tar -y Tar The flat surface is the side of the Earth that is exposed to the sun.

[0110] Based on the above description, when the satellite and the sun are synchronized at a preset time, the nodal coordinate system S Tar x Tar -y Tar If the plane is parallel to the satellite's orbital plane (i.e., the target orbital plane), then the nodal coordinate system S Tar x Tar axis, y Tar axis and z Tar The vector of the axis satisfies:

[0111] x Tar =cosα t ·H sun x n +sinα t (H sun ×x n )+(x n ·H sun (1-cosα) t )H sun

[0112] y Tar =cosα t vH sun y n +sinα t (H sun ×y n )+(y n ·H sun (1-cosα) t )H sun

[0113] z Tar =cosα t ·H sun z n +sinα t (H sun ×z n )+(z n ·H sun (1-cosα) t )H sun Formula (1)

[0114] Where, α t Let α be the rotation angle of the sunlit surface. t Satisfy: α t =ω sun t.

[0115] Given a preset time t0, the precession angle of the spacecraft's source orbital plane is ω. sun At time t0, the target orbital plane of the spacecraft and the nodal coordinate system S Tar x n -y n Planes are parallel.

[0116] S102. Determine the reference position for performing orbital plane adjustments on the spacecraft's current orbit.

[0117] In one implementation, the reference position is the starting position for performing orbital plane adjustments, and the spacecraft begins adjusting its orbital plane when it reaches that reference position.

[0118] In another implementation, the starting and ending positions of the track surface adjustment can be determined based on the reference position.

[0119] Optionally, combined Figure 5 ,like Figure 7 As shown, S102 includes S1021-S1022.

[0120] S1021. Determine the ascending and descending nodes of the spacecraft's current orbital plane and the target orbital plane.

[0121] In this embodiment, the current orbital plane of the spacecraft is constructed by the current orbital elements of the spacecraft, and the ascending and descending nodes of the current orbital plane and the target orbital plane are obtained by simulation.

[0122] Understandably, the ascending node refers to the point where the spacecraft's current orbital plane intersects with the target orbital plane when the spacecraft is moving from south to north; the descending node refers to the point where the spacecraft's current orbital plane intersects with the target orbital plane when the spacecraft is moving from north to south.

[0123] Alternatively, the ascending node and the descending node of the current orbit plane and the target orbit plane can be simulated based on an ephemeris model in the SKT software. The ephemeris model in the SKT software takes the six orbital elements of the current orbit plane and the six orbital elements of the target orbit plane at the initial time as inputs, and obtains the true anomaly of the ascending node and the true anomaly of the descending node of the spacecraft's current orbit plane and the target orbit plane based on the inputs.

[0124] In an implementation manner, the spacecraft is a satellite, the central celestial body is the Earth, and the target celestial body is the Sun. The orbital elements of the current orbit plane of the satellite are simulated, the motion state of the satellite on the current orbit is simulated, and a VNC coordinate system for simulating the current orbit plane of the satellite is established.

[0125] As shown in Figure 8 , the satellite in the simulation experiment is referred to as a maneuvering satellite, the center of the maneuvering satellite is taken as the origin (point O sat ) of the VNC coordinate system, the positive direction of the x-axis of the VNC coordinate system is the direction of the velocity vector of the maneuvering satellite, the positive direction of the z-axis is consistent with the direction of the normal vector of the current orbit plane of the satellite, and the positive direction of the y-axis, the positive direction of the x-axis, and the positive direction of the z-axis form a right-handed coordinate system. The x-y plane of the VNC coordinate system is the plane in which the current orbit plane of the satellite is located.

[0126] Using the simulation software, the x Tar -y Tar plane of the node coordinate system S Tar at the above-mentioned t0time is taken as input, and the ascending node and the descending node of the current orbit plane and the target orbit plane of the spacecraft can be output, and the simulation results are referred to Figure 8 .

[0127] S1022、the point closest to the apogee of the current orbit among the ascending and descending nodes is determined as the reference position.

[0128] It can be understood that the kinetic energy and the potential energy of the spacecraft change with the position of the spacecraft on the orbit during the operation of the spacecraft on the orbit. During the movement of the spacecraft from the perigee on the orbit to the apogee on the orbit, the distance between the spacecraft and the central celestial body becomes larger, the kinetic energy of the spacecraft becomes smaller, and the potential energy becomes larger. When the spacecraft runs to the apogee on the orbit, the kinetic energy of the spacecraft is the smallest, and the potential energy is the largest. Since the kinetic energy of the spacecraft is positively correlated with the speed of the spacecraft, when the spacecraft runs to the apogee on the orbit, the kinetic energy of the spacecraft is the smallest, and the speed of the spacecraft is also the smallest.

[0129] In the embodiments of the present application, in order to make the orbit of the spacecraft be a frozen orbit, when the orbital plane of the spacecraft is adjusted, only the direction of the speed of the spacecraft is changed, and the speed of the spacecraft is not changed. It can be understood that the less the speed of the spacecraft is, the less fuel is required to change the direction of the speed of the spacecraft, and therefore, the closer the reference position is to the apogee, the smaller the speed of the spacecraft at the reference position is, and the less fuel is required to adjust the orbital plane according to the reference position. In conclusion, selecting the point closest to the apogee of the current orbit from among the ascending node and the descending node as the reference position can save the fuel required to perform the adjustment of the orbital plane.

[0130] Optionally, after the precession angle of the source orbital plane of the spacecraft is determined through S101, the precession angle can be further optimized by using the following method, and the reference position determined by the target orbital plane corresponding to the precession angle and the current orbital plane of the satellite is taken as the optimized reference position through S102.

[0131] In an implementation manner, still taking the spacecraft as a satellite, the central celestial body as the Earth, and the target celestial body as the Sun as an example, the precession angle is optimized by using a sequential quadratic programming algorithm.

[0132] Specifically, the optimization increment of the precession angle of the source orbital plane of the satellite is defined as a0, the speed increment of the spacecraft is denoted as Δv, a0 is taken as the optimization variable, the speed increment Δv (the speed increment required to be applied to the satellite to adjust the current orbital plane of the satellite to the target orbital plane) is taken as the performance index of optimization, and the precession angle of the source orbital plane of the satellite is determined in the case that the speed increment is minimum under the constraint conditions of the first constraint condition and the second constraint condition.

[0133] The first constraint condition is that |θ b -θ a |<10°,

[0134] Wherein, θ b is the true anomaly at the reference position determined based on a0, and θ a is the true anomaly at the apogee of the current orbit of the satellite.

[0135] The second constraint condition is that Δβ=Δβ sun ,

[0136] Wherein, Δβ is the angle between the target orbital plane and the current orbital plane of the satellite based on a0, and Δβ sun is the angle of rotation of the sun's illumination surface from the current time to the preset time.

[0137] Based on the above, the optimization variable X of the sequential quadratic programming algorithm satisfies X=a0, the performance index J satisfies J=||Δv||→min, and the constraint condition satisfies: Wherein, θa = 180°.

[0138] After optimization, the precession angle of the satellite's source orbital plane satisfies:

[0139] The velocity increment Δv satisfies:

[0140] Δv = v + -v - Formula (2)

[0141] Wherein, v + represents the velocity vector of the satellite on the target orbital plane, v - represents the velocity vector of the satellite on the current orbital plane (a known quantity).

[0142] In the embodiment of the application, the velocity vector of the satellite includes a radial velocity vector and a tangential velocity vector, wherein the radial velocity vector points to the center of the earth, and the tangential velocity vector is along the tangent direction of the satellite orbit. Then the velocity vector v + of the satellite on the target orbital plane satisfies:

[0143]

[0144] Wherein, represents the radial velocity vector of the satellite on the target orbital plane, represents the tangential velocity vector of the satellite on the target orbital plane, represents the radial velocity vector of the satellite on the current orbital plane (obtained by the motion state of the spacecraft on the orbit at the current time), represents the tangential velocity vector of the satellite on the current orbital plane (obtained by the motion state of the spacecraft on the orbit at the current time), n + represents the normal vector of the target orbital plane, r - represents the position vector of the satellite on the target orbital plane (obtained by the position of the spacecraft on the orbit at the current time), and ||*|| represents the modulus value.

[0145] From the above content of the embodiment of the application, the x Tar -y Tar plane of the node coordinate system S Tar at the preset time is parallel to the target orbital plane, so the vector of the z Tar axis of the node coordinate system S Tar at the preset time is the same as the normal vector n + of the target orbital, that is, n + =z Tar .Therefore, by substituting formula (1) into formula (3), the following is obtained:

[0146]

[0147] Further, substituting the above formula and into formula (2), the following formula is obtained:

[0148]

[0149] From formula (4), it can be seen that the velocity increment applied to the satellite when adjusting the satellite orbit plane only changes the direction of the tangential velocity of the satellite, without changing the size of the tangential velocity of the satellite, nor changing the direction and size of the radial velocity of the satellite.

[0150] S103, based on the reference position and the precession angle of the source orbit plane of the spacecraft, adjusting the motion state of the spacecraft to adjust the current orbit plane of the spacecraft to the target orbit plane.

[0151] In an implementation manner, a chemical propulsion system is adopted to adjust the current orbit plane of the spacecraft to the target orbit plane through pulse maneuver, in combination with Figure 7 As shown in Figure 9 S103 includes S1031A-S1032A.

[0152] It can be understood that pulse maneuver refers to changing the velocity of the spacecraft under the action of a single pulse, so that the orbit of the spacecraft changes. Pulse maneuver is characterized by short action time, and can achieve instantaneous change of the velocity of the spacecraft, thereby achieving instantaneous change of the orbit of the spacecraft.

[0153] S1031A, based on the precession angle of the source orbit plane of the spacecraft, determining the velocity increment of the spacecraft.

[0154] The velocity increment of the spacecraft refers to the velocity increment applied to the spacecraft to adjust the current orbit of the spacecraft to the target orbit plane. The method for determining the velocity increment of the spacecraft can refer to the calculation method of the velocity increment Δv in S1022 of the embodiment of the present application.

[0155] S1032A, when the spacecraft runs to the reference position on the current orbit, applying a pulse to the spacecraft based on the velocity increment to adjust the current orbit plane of the spacecraft to the target orbit plane.

[0156] For example, when the spacecraft is a satellite, the central celestial body is the Earth, and the target celestial body is the Sun, the fuel required to adjust the current orbit of the satellite to the target orbit plane is calculated according to the mass of the satellite and the velocity increment, when the satellite runs to the reference position on the current orbit, a pulse is generated by burning fuel, and then the current orbit plane of the satellite is adjusted to the target orbit plane based on pulse maneuver.

[0157] The formula for calculating the fuel required to adjust the current orbit of the satellite to the target orbit plane is:

[0158]

[0159] Wherein, Δm represents the fuel mass required for adjusting the current orbit plane of the satellite to the target orbit plane, m0 represents the total mass of the satellite, I sp1 represents the specific impulse of the chemical propulsion system of the satellite, the specific impulse refers to the thrust generated by the propulsion system per second for consuming unit mass of fuel, g0 represents the standard gravity acceleration of the earth sea level.

[0160] Figure 10 For the schematic diagram of the simulation result of adjusting the orbit plane of the spacecraft through the pulse maneuver, the simulation result is the variation of the angle between the sun vector (i.e. the sun illumination plane) and the normal vector of the target orbit plane (i.e. the z Tar axis of the nodal coordinate system S Tar ) within 30 days of the satellite running on the orbit, the variation of the angle between the sun vector and the normal vector of the initial orbit plane (i.e. the z n axis of the nodal coordinate system S n ) and the variation of the angle between the sun vector and the normal vector of the current orbit plane of the satellite (i.e. the z axis of the NVC coordinate system). It should be noted that in the simulation process, the orbit plane is adjusted once in each orbit period, and the orbit plane is adjusted to the target orbit plane.

[0161] As can be seen from Figure 10 , within 30 days, the angle between the sun vector and the normal vector of the initial orbit plane gradually increases, which indicates that the source orbit plane of the satellite gradually deviates from the sun illumination plane during this process; the angle between the sun vector and the normal vector of the target orbit plane is almost zero, which indicates that the x Tar -y Tar plane of the nodal coordinate system S Tar has a good simulation effect on the sun illumination plane; the angle between the sun vector and the orbit plane gradually increases (within the range of 0-5°) in an orbit period, and after the orbit plane is adjusted to the target orbit plane, the angle between the sun vector and the orbit plane of the satellite almost returns to zero.

[0162] Since the angle between the orbit plane of the satellite and the sun illumination plane is within the range of 0-5°, the detection surface of the satellite can be within the sun illumination range, so that the satellite can perform the detection task, therefore, the method of realizing the synchronization of the satellite and the sun through the pulse maneuver can make the satellite always within the sun illumination range, thereby realizing the synchronization of the satellite and the sun.

[0163] In another implementation manner, an electric propulsion system is adopted, and the current orbit plane of the spacecraft is adjusted to the target orbit plane through small-thrust maneuver, combined Figure 7 , as shown in Figure 11 , S103 includes S1031B-S1032B.

[0164] It can be understood that the low-thrust maneuver is a process of changing the speed of a spacecraft moving on an orbit by continuously changing the thrust to change the orbit of the spacecraft. The low-thrust maneuver is characterized by a small thrust and a small speed increment, and a thrust needs to be continuously applied for a period of time to change the speed of the spacecraft.

[0165] S1031B, based on the reference position and the argument of the source orbit plane of the spacecraft, determining a starting position and an ending position of the orbit plane adjustment.

[0166] Optionally, when the spacecraft is a satellite, the central celestial body is the Earth, and the target celestial body is the Sun, the mean anomaly θ of the reference position obtained in S102 is adjusted to θ b As an optimization initial value, in combination with the argument of the source orbit plane of the satellite, a low-thrust duration for adjusting the current orbit plane of the satellite to the target orbit plane is searched based on a particle swarm optimization (PSO) algorithm under the condition that the fuel consumption Δm is minimum. An initial time t in The position of the satellite on the orbit is the starting position of the motion state adjustment, and an ending time t f The position of the satellite on the orbit is the ending position of the motion state adjustment. The fuel consumption Δm refers to the fuel required for adjusting the current orbit plane of the satellite to the target orbit plane.

[0167] The optimization variable X of the above-mentioned particle swarm optimization algorithm satisfies: The performance index J satisfies: J = Δm→min; and the constraint condition satisfies: Wherein, θ represents the mean anomaly of the position of the satellite on the orbit at the initial time of the low-thrust duration, θ represents the mean anomaly of the position of the satellite on the orbit at the ending time of the low-thrust duration, θ opt θ represents the mean anomaly of the reference position, and Δβ represents the angle between the target orbit plane and the current orbit plane of the satellite, Δβ sun represents the angle of the sun illumination surface turning from the current time to the preset time, i.e., the angle of the sun vector turning from the current time to the preset time.

[0168] The fuel consumption Δm satisfies:

[0169]

[0170] Wherein, I sp2 represents the specific impulse of the electric propulsion system of the satellite, T max represents the maximum thrust of the electric propulsion system, and u = (T maxu) y, y represents a vector of the y-axis of the VNC coordinate system, u represents a small thrust vector, u represents a normalized small thrust, and u e [0, 1].

[0171] The size of the small thrust is generally the maximum thrust of the electric thrust system, or can be the size of the small thrust after adjusting the gas pressure of the thruster, or can be slightly smaller than the maximum thrust, and the size of the small thrust is not limited in the embodiments of the application.

[0172] S1032B, when the spacecraft runs to the starting position on the current orbit, a continuous thrust is applied to the spacecraft until the spacecraft runs to the ending position, so as to adjust the current orbital plane of the spacecraft to the target orbital plane.

[0173] Figure 12 The simulation results of the method for achieving satellite and sun synchronization through small thrust maneuvering are shown in the schematic diagram. It can be seen that, in the simulation process, the included angle between the sun irradiation surface and the current orbit of the satellite is always within the range of 0-5°, so that the detection surface of the satellite can be within the sun irradiation range, and thus the satellite can perform a detection task. Therefore, the method for achieving satellite and sun synchronization through small thrust maneuvering can keep the satellite within the sun irradiation range, thereby achieving satellite and sun synchronization.

[0174] In some cases, the fuel required for adjusting the current orbital plane of the satellite to the target orbital plane multiple times in the future can be calculated according to the precession angle of the current orbital plane of the satellite and the source orbital plane of the satellite at multiple preset moments in the future, so as to estimate the maximum maintenance time of the satellite synchronized with the sun based on the amount of fuel carried by the satellite.

[0175] The method for achieving spacecraft and target celestial body synchronization provided in the embodiments of the application takes the orbital plane of the spacecraft at a preset moment being parallel to the radiation plane of the target celestial body as a prerequisite condition (which is the condition for the spacecraft and the target celestial body to be synchronized), determines the precession angle of the orbital plane of the spacecraft required for the orbital plane of the spacecraft to change from the source orbital plane to the target orbital plane (i.e., the radiation plane of the target celestial body at the preset moment), determines the reference position of the orbital plane adjustment on the current orbit of the spacecraft, and then adjusts the current orbital plane of the spacecraft to the target orbital plane based on the precession angle of the source orbital plane and the reference position. It can be seen that the method can adjust the current orbital plane of the spacecraft to be parallel to the radiation plane of the target celestial body, so that the orbital plane of the spacecraft is always within the radiation region of the target celestial body, i.e., the spacecraft can be always synchronized with the target celestial body, so as to ensure that the spacecraft can successfully perform a detection task.

[0176] Correspondingly, the embodiments of the application provide a device for achieving spacecraft and target celestial body synchronization, as shown in Figure 13 which includes a precession angle determination module 1201, a reference position determination module 1202, and an orbital plane adjustment module 1203.

[0177] The precession angle determination module 1201 is configured to determine a precession angle of a source orbital plane of the spacecraft when the spacecraft is synchronized with the target celestial body at a preset time.

[0178] The spacecraft is synchronized with the target celestial body when the orbital plane of the spacecraft is parallel to the radiation plane of the target celestial body. The orbital plane after the source orbital plane rotates by the precession angle is the target orbital plane. For example, the precession angle determination module 1201 is configured to implement S101 of the method for synchronizing the spacecraft with the target celestial body.

[0179] The reference position determination module 1202 is configured to determine a reference position for performing the orbital plane adjustment on the current orbit of the spacecraft. For example, the reference position determination module 1202 is configured to implement S102 of the method for synchronizing the spacecraft with the target celestial body.

[0180] The orbital plane adjustment module 1203 is configured to adjust the motion state of the spacecraft based on the reference position and the precession angle of the source orbital plane of the spacecraft, so as to adjust the current orbital plane of the spacecraft to the target orbital plane. For example, the orbital plane adjustment module 1203 is configured to implement S103 of the method for synchronizing the spacecraft with the target celestial body.

[0181] Optionally, the reference position determination module 1202 is specifically configured to determine the ascending node and the descending node of the current orbital plane of the spacecraft and the target orbital plane. The point closest to the apogee of the current orbit among the ascending node and the descending node is determined as the reference position. For example, the reference position determination module 1202 is specifically configured to implement S1021-S1022 of the method for synchronizing the spacecraft with the target celestial body.

[0182] Optionally, the orbital plane adjustment module 1203 is specifically configured to determine a velocity increment of the spacecraft based on the precession angle of the source orbital plane of the spacecraft. When the spacecraft runs to the reference position on the current orbit, an impulse is applied to the spacecraft based on the velocity increment, so as to adjust the current orbital plane of the spacecraft to the target orbital plane. For example, the orbital plane adjustment module 1203 is specifically configured to implement S1031A-S1032A of the method for synchronizing the spacecraft with the target celestial body.

[0183] Optionally, the orbital plane adjustment module 1203 is specifically configured to determine a start position and an end position of the motion state adjustment based on the reference position and the precession angle of the source orbital plane of the spacecraft. When the spacecraft runs to the start position on the current orbit, a continuous thrust is applied to the spacecraft until the spacecraft runs to the end position, so as to adjust the current orbital plane of the spacecraft to the target orbital plane. For example, the orbital plane adjustment module 1203 is specifically configured to implement S1031B-S1032B of the method for synchronizing the spacecraft with the target celestial body.

[0184] In one implementation manner, the method is applied to the spacecraft in the synchronous orbit. Figure 13 For example, the method is applied to the spacecraft in the synchronous orbit.Figure 14 As shown, the aforementioned device for synchronizing the spacecraft with the target celestial body also includes an orbit determination module 1200. The orbit determination module 1200 is used to determine the orbit of the spacecraft around the central celestial body.

[0185] Optionally, the orbit determination module 1200 is specifically used to: generate a we phase diagram based on multiple sets of candidate orbital elements; where e is the eccentricity and w is the perigee argument; the we phase diagram is a curve used to describe the relationship between e and w. The eccentricity corresponding to the target curve in the we phase diagram is determined as the eccentricity of the spacecraft's source orbit; the target curve is the curve with the smallest range of eccentricities in the we phase diagram. For example, the orbit determination module 1200 is specifically used to implement S1-S2 of the above-mentioned method for synchronizing the spacecraft with the target celestial body.

[0186] In another implementation method, combined with Figure 14 ,like Figure 15 As shown, the aforementioned device for synchronizing the spacecraft with the target celestial body also includes an optimization module 1204. The optimization module 1204 is used to optimize the precession angle of the spacecraft's source orbital plane.

[0187] The modules of the aforementioned spacecraft synchronization device with the target celestial body can also be used to perform other steps in the above method embodiments. All relevant content involved in the above method embodiments can be referred to in the functional description of the corresponding functional module, and will not be repeated here.

[0188] This application also provides an electronic device, including: a processor and a memory coupled to the processor; the memory is used to store computer instructions, and when the electronic device is running, the processor executes the computer instructions stored in the memory to cause the electronic device to perform the methods in the above embodiments. The processor can implement the above-described orbit determination module 1200, precession angle determination module 1201, reference position determination module 1202, orbital plane adjustment module 1203, and optimization module 1204; the memory can also be used to store the precession angle, reference position, velocity increment, and start and end positions of motion state adjustment of the spacecraft's source orbital plane.

[0189] This application also provides a computer-readable storage medium including a computer program that, when run on a computer, performs the methods described in the above embodiments.

[0190] This application also provides a computer program product, which includes computer program instructions that, when run on a computer, execute the methods described in the above embodiments.

[0191] The various embodiments in the specification are described in a progressive manner, and the same or similar parts between the various embodiments can be mutually referred to, and each embodiment focuses on the differences from other embodiments.

[0192] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present application, and are not limited thereto; although the present application has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that the technical solutions recorded in the foregoing embodiments can still be modified, or some technical features therein can be replaced by equivalents; and these modifications or replacements do not make the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A method for synchronizing a spacecraft with a target celestial body, characterized in that, The spacecraft orbits a central celestial body, the central celestial body orbits the target celestial body, and the method includes: The precession angle of the spacecraft's source orbital plane when the spacecraft is synchronized with the target celestial body at a predetermined time; the spacecraft is synchronized with the target celestial body when its orbital plane is parallel to the radiation surface of the target celestial body; the orbital plane after the source orbital plane rotates by the precession angle is the target orbital plane; Determine the reference position for performing orbital plane adjustments on the spacecraft's current orbit; Based on the reference position and the precession angle of the spacecraft's source orbital plane, the motion state of the spacecraft is adjusted to adjust the spacecraft's current orbital plane to the target orbital plane.

2. The method as described in claim 1, characterized in that, Determine the reference position for performing orbital plane adjustments within the spacecraft's current orbit, including: Determine the ascending and descending nodes of the spacecraft's current orbital plane and the target orbital plane; The point between the ascending node and the descending node that is closest to the apogee of the current orbit is determined as the reference position.

3. The method as described in claim 1, characterized in that, The precession angle of the spacecraft's source orbital plane satisfies: a t0 =ω c t0 Where t0 represents the preset time, α t0 ω represents the precession angle of the spacecraft's source orbital plane from the initial moment to time t0. c This represents the angular velocity of the central celestial body revolving around the target celestial body.

4. The method as described in claim 1 or 3, characterized in that, The method further includes: The precession angle of the spacecraft's source orbital plane is optimized.

5. The method as described in claim 1, characterized in that, The orbit in which the spacecraft revolves around the central celestial body is a frozen orbit.

6. The method as described in claim 5, characterized in that, The method further includes: Based on multiple candidate orbital elements of the spacecraft, a we phase diagram is generated; where w is the perigee argument and e is the eccentricity; the we phase diagram is a graph used to describe the relationship between w and e. The eccentricity corresponding to the target curve in the we phase diagram is determined as the eccentricity of the spacecraft's source orbit; the target curve is the curve with the smallest range of eccentricity in the we phase diagram.

7. The method as described in claim 1, characterized in that, Adjusting the motion state of the spacecraft based on the reference position and the precession angle of the spacecraft's source orbital plane includes: The velocity increment of the spacecraft is determined based on the precession angle of the spacecraft's source orbital plane; When the spacecraft reaches the reference position on its current orbit, a pulse is applied to the spacecraft based on the velocity increment to adjust the spacecraft's current orbital plane to the target orbital plane.

8. The method as described in claim 1, characterized in that, Adjusting the motion state of the spacecraft based on the reference position and the precession angle of the spacecraft's source orbital plane includes: Based on the reference position and the precession angle of the spacecraft's source orbital plane, the starting and ending positions of the motion state adjustment are determined; When the spacecraft reaches the starting position on the current orbit, continuous thrust is applied to the spacecraft until it reaches the ending position, so as to adjust the current orbital plane of the spacecraft to the target orbital plane.

9. The method as described in claim 1, characterized in that, The target celestial body is the Sun, and the central celestial body is the Earth.

10. An electronic device, characterized in that, The device includes a processor and a memory coupled to the processor; the memory is used to store computer instructions, which, when the electronic device is running, are executed by the processor to cause the electronic device to perform the method as described in any one of claims 1 to 9.

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