Method and device for controlling deep space spacecraft

Abstract

The invention provides a control method and device for a deep space spacecraft, belongs to the technical field of cosmic navigation, and can improve the control precision of the deep space spacecraft, thereby better reducing the fuel consumption of the deep space spacecraft during the long-term execution of a scientific exploration task. According to the method, orbit deviation between an orbit where the deep space spacecraft runs and a nominal orbit is divided into a divergent component and a non-divergent component of the deep space spacecraft, and a maneuvering response curve, indicating the non-divergent component, of the deep space spacecraft is obtained through analysis according to orbit change after orbit maneuvering of the deep space spacecraft. And determining a final value of the orbit maneuver of the deep space spacecraft capable of inhibiting the divergent component and the non-divergent component by combining the initial value of the orbit maneuver of the divergent component of the deep space spacecraft and the maneuver response curve of the deep space spacecraft.

Description

Technical Field

[0001] The present invention relates to the technical field of space navigation, and particularly to a control method and device for a deep space spacecraft. Background Art

[0002] A deep space spacecraft refers to an unmanned spacecraft that performs scientific exploration tasks in the solar system space outside the Earth's orbit (such as the lunar orbit, planetary orbits, asteroid orbits, comet orbits, or interstellar space). During the process of a deep space spacecraft performing scientific exploration tasks, due to the perturbation of the solar system space on the deep space spacecraft, the deep space spacecraft deviates from its nominal orbit (such as a libration point orbit, etc.). Therefore, it is necessary to control the deep space spacecraft so that the deep space spacecraft can overcome the perturbation and ensure that the deep space spacecraft can operate on the nominal orbit to complete the scientific exploration tasks.

[0003] However, the existing control methods for deep space spacecrafts have low control accuracy for deep space spacecrafts, resulting in large fuel consumption of deep space spacecrafts during long-term scientific exploration tasks. Summary of the Invention

[0004] The present invention provides a control method and device for a deep space spacecraft, which can improve the control accuracy of the deep space spacecraft, thereby preferably reducing the fuel consumption of the deep space spacecraft during long-term scientific exploration tasks.

[0005] To achieve the above object, the present invention adopts the following technical solutions:

[0006] In a first aspect, the present invention provides a control method for a deep space spacecraft, including: determining a maneuver response curve of the deep space spacecraft based on the orbit change of the deep space spacecraft after orbital maneuver; wherein, there is a deviation between the orbit on which the deep space spacecraft operates and the nominal orbit of the deep space spacecraft; the maneuver response curve includes non-divergent components of the deep space spacecraft after orbital maneuver at each position on the orbit on which the deep space spacecraft operates; the non-divergent component of the deep space spacecraft is the deviation between the orbit on which the deep space spacecraft operates and the nominal orbit that neither exponentially diverges nor converges; the divergent component of the deep space spacecraft is the deviation between the orbit on which the deep space spacecraft operates and the nominal orbit that exponentially diverges. When the deep space spacecraft deviates from its nominal orbit during the navigation process, determining an initial value of the orbital maneuver of the deep space spacecraft according to the orbit on which the deep space spacecraft operates, the nominal orbit of the deep space spacecraft, and the divergent component of the deep space spacecraft. Then, determining a final value of the orbital maneuver of the deep space spacecraft according to the maneuver response curve of the deep space spacecraft and the initial value of the orbital maneuver of the deep space spacecraft; the final value of the orbital maneuver guides the deep space spacecraft to its nominal orbit.

[0007] In the control method of a deep - space spacecraft provided by the present invention, the orbital deviation between the orbit on which the deep - space spacecraft operates and the nominal orbit is divided into the divergence component and the non - divergence component of the deep - space spacecraft. And through the orbital change after the deep - space spacecraft performs an orbital maneuver, the maneuver response curve of the deep - space spacecraft indicating the non - divergence component is analyzed. Then, during the navigation of the spacecraft, according to the divergence component of the deep - space spacecraft, the orbit on which the deep - space spacecraft operates, and the nominal orbit of the deep - space spacecraft, the initial value of the orbital maneuver capable of suppressing the divergence component of the deep - space spacecraft is determined. Finally, combining the initial value of the orbital maneuver and the maneuver response curve of the deep - space spacecraft, the final value of the orbital maneuver of the deep - space spacecraft is determined. The final value of the orbital maneuver of the deep - space spacecraft can not only suppress the divergence component of the deep - space spacecraft but also suppress the non - divergence component of the deep - space spacecraft. The final value of the orbital maneuver determined by the above method guides the deep - space spacecraft to the nominal orbit of the deep - space spacecraft, which can improve the control accuracy of the deep - space spacecraft, thereby better reducing the fuel consumption of the deep - space spacecraft during the long - term execution of scientific exploration missions.

[0008] In one implementation manner of the first aspect, the maneuver response curve of the deep - space spacecraft satisfies the following formula;

[0009]

[0010] where, Δa i represents the change amount of the i - th mode in the non - divergence component of the deep - space spacecraft after the orbital maneuver, i represents the i - th mode of the deep - space spacecraft, a 1,max represents the maneuver threshold of the mode vector of the deep - space spacecraft on the first mode after the orbital maneuver, |Π v,1 | represents the modulus of the projection vector of the mode vector of the deep - space spacecraft on the first mode after the orbital maneuver, Π v,1 T represents the transpose of the projection vector of the mode vector of the deep - space spacecraft on the first mode after the orbital maneuver, Π v,i represents the projection vector of the mode vector of the deep - space spacecraft on the i - th mode after the orbital maneuver.

[0011] In one implementation manner of the first aspect, determining the final value of the orbital maneuver of the deep - space spacecraft includes:

[0012] Adopting the golden section method, select the optimal correction coefficient for suppressing the divergence component of the deep - space spacecraft within the value range of the correction coefficient of the deep - space spacecraft;

[0013] Calculate the maneuver overshoot coefficient of the deep - space spacecraft, and the maneuver overshoot coefficient of the deep - space spacecraft satisfies the following formula;

[0014] δ v =|δ v |·ηδ

[0015] Among them, δ v represents the maneuver overshoot coefficient of the deep space spacecraft, and |δ v | represents the modulus of the maneuver overshoot coefficient of the deep space spacecraft, and η δ represents the sign of the maneuver overshoot coefficient of the deep space spacecraft, and η δ = η v ·η MRC ·η α ; η v represents the direction of the previous orbital maneuver of the deep space spacecraft, and η MRC represents the sign of the component to be controlled of the deep space spacecraft; η α represents the adjustment direction of the component to be controlled of the deep space spacecraft for the current orbital maneuver; the component to be controlled is the mode with the largest value among the non-divergent components of the deep space spacecraft;

[0016] Calculate the final value of the orbital maneuver of the deep space spacecraft, and the final value of the orbital maneuver of the deep space spacecraft satisfies the following formula;

[0017] Δv = (1 + κ + δ v )Δv ori

[0018] Among them, Δv represents the final value of the orbital maneuver of the deep space spacecraft, κ represents the optimal correction coefficient of the divergent component of the deep space spacecraft, and Δv ori represents the initial value of the orbital maneuver of the deep space spacecraft.

[0019] In an implementation manner of the first aspect, the initial value of the orbital maneuver of the deep space spacecraft satisfies the following formula;

[0020] (δX + Δv ori )·Π = 0

[0021] Among them, δX represents the orbital deviation between the orbit on which the deep space spacecraft operates and the nominal orbit, Δv ori represents the initial value of the orbital maneuver of the deep space spacecraft, and Π represents the projection vector of the modal vector of the deep space spacecraft.

[0022] In an implementation manner of the first aspect, the nominal orbit is a libration point orbit.

[0023] Second aspect, the present invention provides a control device for a deep space spacecraft, including a response curve determination module, an initial value determination module, and a final value determination module. The response curve determination module is configured to determine the maneuver response curve of the deep space spacecraft based on the orbital change of the deep space spacecraft after orbital maneuver; wherein, there is a deviation between the orbit on which the deep space spacecraft operates and the nominal orbit of the deep space spacecraft; the maneuver response curve includes the non-divergent component of the deep space spacecraft after orbital maneuver at each position of the orbit on which the deep space spacecraft operates; the non-divergent component of the deep space spacecraft is the deviation between the orbit on which the deep space spacecraft operates and the nominal orbit that neither exponentially diverges nor converges; the divergent component of the deep space spacecraft is the deviation between the orbit on which the deep space spacecraft operates and the nominal orbit that exponentially diverges. The initial value determination module is configured to determine the initial value of the orbital maneuver of the deep space spacecraft according to the orbit on which the deep space spacecraft operates, the nominal orbit of the deep space spacecraft, and the divergent component of the deep space spacecraft when the deep space spacecraft deviates from the nominal orbit of the deep space spacecraft during the navigation process. The final value determination module is configured to determine the final value of the orbital maneuver of the deep space spacecraft according to the maneuver response curve of the deep space spacecraft and the initial value of the orbital maneuver of the deep space spacecraft; the final value of the orbital maneuver guides the deep space spacecraft to the nominal orbit.

[0024] Third aspect, the present invention 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 runs, the processor executes the computer instructions stored in the memory so that the electronic device executes the method described in the first aspect or any one of its implementation manners above.

[0025] Fourth aspect, the present invention provides a computer-readable storage medium, including computer program instructions, and when the computer program instructions are executed by a computer, the computer is caused to execute the method described in the first aspect or any one of its implementation manners above.

[0026] Fifth aspect, the present invention provides a computer program product, including computer program instructions, and when the computer program instructions run on a computer, the computer is caused to execute the method described in the first aspect or any one of its implementation manners above.

[0027] For the technical effects corresponding to the second aspect to the fifth aspect and their possible implementation manners, reference can be made to the description of the technical effects of the first aspect and its possible implementation manners above, and details are not repeated here. Description of the Drawings

[0028] Figure 1 is one of the schematic diagrams of a control method for a deep space spacecraft provided by an embodiment of the present application;

[0029] Figure 2 It is a schematic diagram of the x-y plane of the rotating coordinate system provided by an embodiment of the present application;

[0030] Figure 3 It is a schematic diagram of the maneuver response curve of a deep space spacecraft provided by an embodiment of the present application;

[0031] Figure 4 It is the second schematic diagram of a control method for a deep space spacecraft provided by an embodiment of the present application;

[0032] Figure 5 It is a schematic structural diagram of a control device for a deep space spacecraft provided by an embodiment of the present application. Detailed implementation manners

[0033] In the description of the present invention, if there are terms such as "first" and "second" in the specification and claims, they are used to distinguish different objects, rather than to describe a specific order of the objects.

[0034] In the embodiments of the present application, "and / or" represents the relationship between objects. For example, A and / or B can represent the following three situations: A exists alone, B exists alone, and A and B exist simultaneously.

[0035] In the embodiments of the present application, words such as "exemplary" or "for example" are used to give examples, illustrations or explanations. Any embodiment or design solution described as "exemplary" or "for example" in the embodiments of the present application should not be construed as being more preferred or having more advantages than other embodiments or design solutions. Exactly speaking, using words such as "exemplary" or "for example" aims to present relevant concepts in a specific manner.

[0036] In the description of the present invention, unless otherwise specified, the meaning of "multiple times" refers to two or more times. For example, multiple orbital maneuvers refer to two or more orbital maneuvers.

[0037] The method and device provided by the embodiments of the present application relate to the orbital control of deep space spacecraft, and can be used to control the deep space spacecraft to perform multiple orbital maneuvers to suppress the deep space spacecraft from deviating from the nominal orbit during the process of performing scientific exploration tasks, so as to ensure that the deep space spacecraft can operate on the nominal orbit to complete the scientific exploration tasks.

[0038] It can be understood that during the navigation of the deep space spacecraft, due to the influence of various disturbances (or called perturbations) in the space environment, there is an orbital deviation between the orbit on which the deep space spacecraft operates and the nominal orbit of the deep space spacecraft. In this case, orbital maneuvers are performed on the deep space spacecraft to guide the deep space spacecraft back to the nominal orbit through the orbital maneuvers.

[0039] To solve the problem in the background art that the existing control method for deep-space spacecraft has low control accuracy for deep-space spacecraft, resulting in large fuel consumption during long-term scientific exploration missions, the embodiments of the present application provide a control method and device for deep-space spacecraft. The orbital deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit is divided into the divergence component and the non-divergence component of the deep-space spacecraft. And through the orbital change after the deep-space spacecraft performs an orbital maneuver, the maneuver response curve of the deep-space spacecraft indicating the non-divergence component is analyzed. Then, in combination with the initial value of the orbital maneuver that can suppress the divergence component of the deep-space spacecraft and the maneuver response curve of the deep-space spacecraft, the final value of the orbital maneuver of the deep-space spacecraft that can suppress the divergence component and the non-divergence component is determined. The final value of the orbital maneuver determined by the above method guides the deep-space spacecraft to the nominal orbit of the deep-space spacecraft, which can improve the control accuracy of the deep-space spacecraft, thereby better reducing the fuel consumption of the deep-space spacecraft during long-term scientific exploration missions.

[0040] Exemplarily, a control method for a deep-space spacecraft provided by an embodiment of the present invention can be executed by an electronic device with processing capabilities. For example, the electronic device can be a computer, a server, etc. Taking the electronic device as a 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.

[0041] Among them, the processor is used to control the electronic device to execute relevant processing and calculation tasks. For example, determining the maneuver response curve of the deep-space spacecraft, determining the initial value and the final value of the orbital maneuver of the deep-space spacecraft, etc. The processor can include a central processing unit (CPU) or other processors. The processor can be single-core or multi-core. For example, the processor can include multiple CPUs.

[0042] The memory is used to store computer instructions and related data. For example, storing the maneuver response curve of the deep-space spacecraft, the orbit on which the deep-space spacecraft operates, the nominal orbit, the initial value of the orbital maneuver, and the final value of the orbital maneuver. 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 any other magnetic storage device, or any other medium capable of storing program code or data that can be accessed by a computer. Optionally, the memory can be integrated in the processor, and the memory can also be independent of the processor.

[0043] The network interface is used for the computer to communicate with other devices or communication networks. The network interface can be a transceiver with transceiver functions. Optionally, the network interface can include a standard wired interface or a wireless interface (such as a Wi-Fi interface, a Bluetooth interface, or a 5G interface).

[0044] The communication bus is used to achieve connection and communication between different components. For example, the processor, memory, network interface, and user interface mentioned above can be interconnected through the communication bus.

[0045] The user interface may include a display screen and an input unit (such as a keyboard). Optionally, the user interface may also include a standard wired interface and a wireless interface.

[0046] Those skilled in the art will appreciate that the above-mentioned computer may also include more or fewer components, or a combination of certain components, or different arrangements of components, which is not limited in the embodiments of the present application.

[0047] Before describing in detail a control method and device for a deep space spacecraft provided in an embodiment of the present application, some professional terms involved in the embodiment of the present application are explained below.

[0048] 1. Nominal track

[0049] The nominal orbit refers to the ideal orbit of a deep-space spacecraft, predetermined based on mission design and theoretical calculations. The nominal orbit serves as the reference orbit for deep-space spacecraft mission planning, describing the orbital parameters that a deep-space spacecraft should follow under undisturbed conditions (e.g., ignoring perturbation forces). In this embodiment, the nominal orbit of a deep-space spacecraft is a libration point orbit.

[0050] 2. Libration Point and Libration Point Orbit

[0051] In celestial mechanics, libration points (also known as Lagrange points) are unique gravitational equilibrium points between two massive celestial bodies (such as the Earth and the Sun, or the Earth and the Moon). At these points, a small object (such as a satellite or asteroid) can remain stationary relative to the two massive bodies. Currently, there are five such points: L1, L2, L3, L4, and L5.

[0052] Lagrange Point Orbits refer to the special orbits in which a spacecraft moves near the Lagrange points (L1 - L5). Since the Lagrange points themselves are gravitational equilibrium points, the motion of the spacecraft near these points is governed by complex dynamics, with diverse orbit types, and unique stability and application value. Common types of Lagrange point orbits include Periodic Orbits, Quasi-Periodic Orbits, and Transfer Orbits.

[0053] It should be noted that in the embodiments of this application, the nominal orbit of the deep-space spacecraft is a periodic orbit among the Lagrange point orbits, such as the Halo orbit.

[0054] 3. Circular Restricted Three-Body Problem

[0055] In orbital mechanics and celestial mechanics, the Circular Restricted Three-Body Problem (CRTBP for short) is a special simplified model of the Three-Body Problem, used to study the motion laws of small celestial bodies under the gravitational action of two large-mass celestial bodies.

[0056] CRTBP introduces the following three key restrictive conditions on the basis of the classical three-body problem:

[0057] I. Restricted: Only study the motion of small celestial bodies in the gravitational fields of the two large celestial bodies, and the motion of the two large celestial bodies is not affected by the small celestial body; the mass of the third celestial body (such as a spacecraft, an asteroid) is much smaller than that of the two main celestial bodies (such as the Earth-Moon, Sun-Earth), so its gravity can be ignored for the motion of the two main celestial bodies.

[0058] II. Circular: The two main celestial bodies move in uniform circular motion (Keplerian circular orbits) around their common center of mass, and the orbital angular velocity is constant.

[0059] III. Planar: Usually assume that the motion occurs in the plane of the orbits of the two large celestial bodies (i.e., a two-dimensional problem), but it can also be extended to the three-dimensional case.

[0060] 4. Floquet Mode Method

[0061] The Floquet modal method is an important mathematical tool for studying the stability and dynamic characteristics of periodically time-varying linear systems (such as systems for controlling deep-space spacecraft), and is widely used in the dynamic analysis of libration point orbits (such as Halo orbits and Lissajous orbits). This method is based on the Floquet theory, and by solving the periodic state transition matrix of the system, it reveals the stability, modal structure, and control strategy of the orbit.

[0062] Optionally, as Figure 1 shown, a control method for a deep-space spacecraft provided by an embodiment of the present application includes S101-S104.

[0063] S101. Determine the maneuver response curve of the deep-space spacecraft based on the orbit change of the deep-space spacecraft after orbital maneuver.

[0064] In one implementation, the above S101 includes S1011-S1013.

[0065] S1011. Calculate the orbit deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit of the deep-space spacecraft.

[0066] It should be understood that in the process of determining the maneuver response curve of the above deep-space spacecraft, in order to simulate the perturbation conditions (such as gravitational perturbation generated by unknown celestial bodies, periodic perturbation generated by the deep-space spacecraft during periodic motion, tracking error introduced by the navigation system, and maneuver error generated during maneuver, etc.) suffered by the deep-space spacecraft in the actual space environment, it is assumed that there is a deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit of the deep-space spacecraft.

[0067] The solution process of the orbit deviation between the orbit on which the above deep-space spacecraft operates and the nominal orbit of the deep-space spacecraft is described below.

[0068] Step 1. Construct the normalized motion equation of the deep-space spacecraft based on the CRTBP framework. The normalized motion equation of the deep-space spacecraft satisfies:

[0069] where x, y, and z respectively represent the position components of the deep-space spacecraft on the x-axis, y-axis, and z-axis of the rotating coordinate system, respectively represent the velocity components of the deep-space spacecraft on the x-axis, y-axis, and z-axis of the rotating coordinate system, respectively represent the acceleration components of the deep-space spacecraft on the x-axis, y-axis, and z-axis of the rotating coordinate system, and U represents the pseudo-potential function of the deep-space spacecraft under CRTBP.

[0070] The above pseudo-potential function U satisfies:

[0071]

[0072] r1 = [(x - μ) 2 + y 2 + z 2 12 , r2 = [(x - μ + 1) 2 + y 2 + z 2 12

[0073] Wherein, μ represents the mass constant in CRTBP, μ = secondary body mass / (primary body mass + secondary body mass), r1 represents the distance from the deep - space spacecraft to the primary body, and r2 represents the distance from the deep - space spacecraft to the secondary body.

[0074] The above - mentioned rotating coordinate system is constructed based on CRTBP. The origin of the above - mentioned rotating coordinate system is set at the centroid of the primary body and the secondary body (for example, the sun and the earth, with the sun as the primary body and the earth as the secondary body). The positive direction of the x - axis of the above - mentioned rotating coordinate system points from the centroid to this celestial body. The positive direction of the z - axis is perpendicular to the rotating plane formed by the primary and secondary bodies, and the positive direction of the y - axis follows the right - hand rule (Right - Hand Rule) with respect to the spatial direction relationship with the x - axis and the z - axis. Figure 2 is the x - y plane of the rotating coordinate system. Referring to Figure 2 It can be known that point P is the coordinate point of the deep - space spacecraft in the rotating coordinate system. The coordinate value of point P is (x, y, z). The coordinate value of the primary body m1 in the rotating coordinate system is (-μ, 0, 0), and the coordinate value of the secondary body m2 in the rotating coordinate system is (1 - μ, 0, 0).

[0075] Step 2: Solve the above - mentioned orbit deviation.

[0076] Let be the state vector of the deep - space spacecraft. Let X(t) represent the spacecraft trajectory composed of all state points at different times. Then the normalized motion equation of the deep - space spacecraft can be transformed into the following formula.

[0077]

[0078] Let represent the nominal orbit of the deep - space spacecraft, and X(t) represent the orbit on which the deep - space spacecraft operates (i.e., the actual orbit of the deep - space spacecraft). It should be noted that the above X(t) is obtained by integrating based on the orbit initial value For example, under the earth - moon CRTBP, the initial value example of the halo orbit is given as follows: X0 = [0.8233842785, 0, - 0.0133820227695511, - 2.2×10 -16 , 0.1292541343, 1.6×10 -15 ​​, the normalized period of the orbit is T = 2.7442213061.

[0079] Then the orbit deviation δX(t) between the orbit on which the deep - space spacecraft is operating at the current moment and the nominal orbit is Then δX(t) satisfies the following formula.

[0080]

[0081] Among them, A(t) is the Jacobian matrix of f(X). A(t) satisfies the following formula. Then δX(t) can be solved through the above formula (1).

[0082]

[0083] S1012. Perform modal decomposition on the above - mentioned orbit deviation to obtain the divergence component and non - divergence component of the deep - space spacecraft.

[0084] In the embodiment of the present application, the divergence component of the deep - space spacecraft is the deviation with exponential divergence among the orbit deviations between the orbit on which the deep - space spacecraft is operating and the nominal orbit. The non - divergence component of the deep - space spacecraft is the deviation that neither exponentially diverges nor converges among the orbit deviations between the orbit on which the deep - space spacecraft is operating and the nominal orbit.

[0085] Since multiple modal components are obtained after performing modal decomposition on the above - mentioned orbit deviation, correspondingly, the divergence component of the deep - space spacecraft is the modal component with exponential divergence among the above - mentioned multiple modal components, and the non - divergence component of the deep - space spacecraft is the modal component that neither exponentially diverges nor converges among the above - mentioned multiple modal components.

[0086] The process of performing modal decomposition on the above - mentioned orbit deviation will be described below.

[0087] Step 1. Perform modal decomposition on the orbit deviation to obtain the unit modal vectors of multiple modal components.

[0088] The unit modal vectors of the above - mentioned multiple modal components satisfy the following formula.

[0089]

[0090] Among them, t represents the current moment, δX(t) represents the orbit deviation between the orbit on which the deep - space spacecraft is operating at the current moment and the nominal orbit. δx, δy, and δz respectively represent the deviation values between the actual position of the deep - space spacecraft at the current moment and the corresponding position of the deep - space spacecraft on the nominal orbit on the x - axis, y - axis, and z - axis of the rotating coordinate system. And They respectively represent the deviation values between the actual velocity of the deep-space spacecraft at the current moment and the corresponding velocity of the deep-space spacecraft on the nominal orbit on the x-axis, y-axis, and z-axis of the rotating coordinate system. α i represents the i-th modal component (which can also be called the deviation component) obtained by performing modal decomposition on the above orbit deviation δX, and i = 1, 2, …, 6. The set of modal components obtained by performing modal decomposition on the orbit deviation is {α1, α2, α3, α4, α5, α6}. represents the unit modal vector of the i-th modal component obtained by performing modal decomposition on the above orbit deviation δX.

[0091] Step 2: Calculate the modal vector E(t).

[0092] Furthermore, after obtaining the above formula (2), the modal vector E(t) obtained by performing modal decomposition on the above orbit deviation δX(t) satisfies the following formula.

[0093]

[0094] Among them, t represents the current moment, and 0 represents the initial moment.

[0095] The above E(t) = [e1(t), e2(t), e3(t), e4(t), e5(t), e6(t)]. Φ(t, 0) is the deviation value between the state of the deep-space spacecraft at the current moment and the state of the deep-space spacecraft at the initial moment, which is calculated by the state deviation transition matrix Φ(t2, t1) of the deep-space spacecraft. J R is the first intermediate parameter, and S is the second intermediate parameter.

[0096] Continue to explain the parameters in the above formula (3). For any two different moments t1, t2, the above state deviation transition matrix Φ(t2, t1) satisfies: δX(t2) = Φ(t2, t1)δX(t1). Then, in the above formula (4), substituting t1 = t, t2 = 0 into the above state deviation transition matrix, Φ(t, 0) can be obtained.

[0097] The above first intermediate parameter J R and the above second intermediate parameter S are solved according to the monodromy matrix C of the deep-space spacecraft. Specifically, perform eigenvalue decomposition on the monodromy matrix C of the deep-space spacecraft to obtain the eigenvalue diagonal matrix D and the corresponding eigenmatrix V, satisfying C·V = V·D. Perform real Jordan decomposition on the real matrix (V, D) pair, and correspondingly obtain the diagonal form matrix (S, J C ), to obtain the second intermediate parameter S. Then take the matrix logarithm of J C , that is, J R = ln(J C ), to obtain the first intermediate parameter JR 。

[0098] The single-value matrix C of the above deep-space spacecraft is obtained by solving the state deviation transfer matrix Φ(t2, t1) of the deep-space spacecraft. The single-value matrix C of the deep-space spacecraft satisfies: C = Φ(t + T peri , t) = Φ(T peri , 0). Wherein, T peri represents the orbital period of the orbit in which the deep-space spacecraft operates.

[0099] Specifically, taking the nominal orbit of the above deep-space spacecraft as the halo orbit as an example, in the above modal vector E(t), e1(t) represents the deviation pointing in the most divergent direction of the deep-space spacecraft on the orbit it operates, e2(t) represents the deviation pointing in the fastest converging direction of the deep-space spacecraft on the orbit it operates, e3(t) and e4(t) represent the deviations pointing in the tangential direction of the orbit along which the deep-space spacecraft operates, and e5(t) and e6(t) represent the deviations generated by the periodic motion of the deep-space spacecraft on the orbit it operates.

[0100] Correspondingly, when the nominal orbit of the deep-space spacecraft is the halo orbit, the modal component α1 corresponding to e1(t) represents the exponentially divergent deviation in the orbital deviation δX(t). Its physical meaning is that the deviation in the α1 direction will exponentially diverge without performing an orbital maneuver on the deep-space spacecraft, which means that without performing an orbital maneuver on the deep-space spacecraft, the deviation between the orbit on which the spacecraft operates and the nominal orbit in the α1 direction will exponentially increase. The modal component α2 corresponding to e2(t) represents the self-converging deviation in the orbital deviation δX(t). Its physical meaning is that the deviation in the α2 direction will self-converge without performing an orbital maneuver on the deep-space spacecraft, which means that without performing an orbital maneuver on the deep-space spacecraft, the deviation between the orbit on which the spacecraft operates and the nominal orbit in the α2 direction will disappear by itself. The modal components α3, α4, α5, and α6 corresponding to e3(t), e4(t), e5(t), and e6(t) represent the deviations in the orbital deviation δX(t) that neither exponentially diverge nor converge. Its physical meaning is that without performing an orbital maneuver on the deep-space spacecraft, the deviations in the α3, α4, α5, and α6 directions still exist and accumulate over time.

[0101] It can be understood that during the process of controlling the deep-space spacecraft to navigate, there is no need to consider the influence of the self-converging deviation on the orbit of the deep-space spacecraft, that is, there is no need to suppress the above modal component α2 through an orbital maneuver. Therefore, when the nominal orbit of the deep-space spacecraft is the halo orbit, the divergent component of the deep-space spacecraft is α1, and the non-divergent components of the deep-space spacecraft are α3, α4, α5, and α6.

[0102] S1013. Simulate the non-divergent component of the deep-space spacecraft to obtain the maneuver response curve of the deep-space spacecraft.

[0103] In one implementation, the maneuver response curve of the above deep-space spacecraft satisfies the following formula.

[0104]

[0105] Where, Δa i represents the change amount of the i-th mode in the non-divergent component of the deep-space spacecraft after orbital maneuver, and i represents the i-th mode of the deep-space spacecraft. a 1,max represents the maneuver threshold of the mode vector of the deep-space spacecraft on the first mode after orbital maneuver. Optionally, a 1,max can take a value of 10 -4 , or can also be 10 -5 , which is not limited in the embodiments of the present application. |Π v,1 | represents the modulus of the projection vector of the mode vector of the deep-space spacecraft on the first mode after orbital maneuver, and Π v,1 T represents the transpose of the projection vector of the mode vector of the deep-space spacecraft on the first mode after orbital maneuver, and Π v,i represents the projection vector of the mode vector of the deep-space spacecraft on the i-th mode after orbital maneuver.

[0106] The following details the specific derivation process of the above formula (4).

[0107] Step 1. Calculate the projection vectors of the multiple mode vectors obtained by solving formula (2).

[0108] The projection vectors of the multiple mode vectors satisfy the following formula.

[0109] Π i = E(t) -1 ·ζ i = [π i,1 , π i,2 , π i,3 , π i,4 , π i,5 , π i,6 T Formula (5)

[0110] Where, Π i represents the projection vector along the unit component of the mode vector; π i,1 ~π i,6 are the 6 components of Π i ; ζ iζi is a 6-dimensional column vector with the i-th term being 1 and the rest being 0. For example, ζ4 = [0, 0, 0, 1, 0, 0] T .

[0111] Step 2: Derive the response formula satisfied by the maneuver response curve of the deep space spacecraft, that is, the above formula (4).

[0112] In one implementation, when the orbital maneuver of the deep space spacecraft is the velocity change of the deep space spacecraft, only the influence of the modal vector on the velocity of the deep space spacecraft needs to be considered. At this time, the projection vector Π along the unit component v,i = [0, 0, 0, π i,4 , π i,5 , π i,6 T .

[0113] It should be noted that since the existing control methods for deep space spacecraft (such as the Floquet modal method) always suppress the divergence components of the deep space spacecraft, and since the nominal orbit of the deep space spacecraft in the embodiments of the present application is a periodic orbit in the libration point orbit, the decomposition of the modal vector of the divergence components is only related to the dynamics near the libration point, and thus, like the periodic orbit, has periodicity. Then it can be proved that if orbital maneuvers are applied at different phases of the periodic orbit, the maneuver vectors only differ in magnitude and positive / negative, and there is no change in direction. Moreover, each orbital maneuver can not only suppress the divergence of the divergence components but also affect the non-divergence components.

[0114] On this basis, when the orbital maneuver of the deep space spacecraft acts on the deep space spacecraft, there is a fixed proportional relationship between the change amount of the deep space spacecraft in the divergence components and the change amount in the non-divergence components.

[0115] In the embodiments of the present application, when the modulus value of the divergence component a1 reaches the maneuver threshold a 1,max , it is necessary to perform an orbital maneuver on the deep space spacecraft. Therefore, when it is clear that the fixed proportional relationship between the change amount in the divergence components and the change amount in the non-divergence components is , through a 1,max and , the change amount (also called the maneuver response value) of the i-th mode in the non-divergence components of the deep space spacecraft after the orbital maneuver can be derived. Thus, the above formula (4) is derived.

[0116] ​In an application scenario of the above implementation method, by using the above formula (4) to simulate the non-divergent components of a deep-space spacecraft, the maneuver response curve of the deep-space spacecraft can be obtained. Specifically, taking the nominal orbit of the above deep-space spacecraft as a halo orbit as an example, the maneuver response curves of the non-divergent components (α3, α4, α5, and α6) of the deep-space spacecraft obtained by simulating with the above formula (4) are as Figure 3 shown.

[0117] S102. During the navigation process, when the deep-space spacecraft deviates from the nominal orbit of the deep-space spacecraft, determine the initial value of the orbital maneuver of the deep-space spacecraft according to the orbit on which the deep-space spacecraft operates, the nominal orbit of the deep-space spacecraft, and the divergent components of the deep-space spacecraft.

[0118] As can be known from the relevant description in S101, the judgment criterion for the above deep-space spacecraft to deviate from the nominal orbit of the deep-space spacecraft is that the divergent component of the deep-space spacecraft is greater than or equal to the maneuver threshold. Taking the nominal orbit of the deep-space spacecraft as a halo orbit as an example, the above judgment criterion can be |a1| ≥ a 1,max . Then, during the navigation process of the deep-space spacecraft, continuously use formula (1) to solve the orbital deviation δX(t) between the orbit on which the deep-space spacecraft operates at the current moment and the nominal orbit, and then decompose δX(t) through formula (2) to obtain a1. When |a1| ≥ a 1,max , it is determined that the deep-space spacecraft deviates from the nominal orbit of the deep-space spacecraft, and the deep-space spacecraft needs to perform an orbital maneuver.

[0119] Optionally, the above orbital maneuver can be the control acceleration acting on the deep-space spacecraft, or the velocity change of the deep-space spacecraft, or the pulsed force acting on the deep-space spacecraft. The embodiments of the present application do not make limitations. Taking the above orbital maneuver as the velocity change of the deep-space spacecraft as an example, the initial value of the orbital maneuver of the deep-space spacecraft satisfies the following formula.

[0120] (δX + Δv ori )·Π = 0 Formula (6)

[0121] where δX represents the orbital deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit, Δv ori represents the initial value of the orbital maneuver of the deep-space spacecraft, Π represents the projection vector of the modal vector of the deep-space spacecraft, Π = [π1, π2, π3, π4, π5, π6] T and Π = E(t) -1 ·[1, 0, 0, 0, 0, 0] T . The physical meaning of the projection vector Π of the modal vector of the deep-space spacecraft is the projection of the orbital deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit on the divergent component of the deep-space spacecraft.

[0122] Specifically, when applying the above formula (6) to an x-y-z three-axis (corresponding to the x-y-z three axes of the rotating coordinate system) controller, the above formula (6) can be transformed into: Δv x π4 + Δv y π5 + Δv z π6 + α1 = 0. In the above case, the solution component that minimizes the modulus value of Δv satisfies the following formula.

[0123]

[0124] When applying the above formula (5) to an x-y two-axis (corresponding to the x-y two axes of the rotating coordinate system) controller, the above formula (5) can be transformed into: Δv x π4 + Δv y π5 + α1 = 0.

[0125] S103. Determine the final value of the orbital maneuver of the deep-space spacecraft according to the maneuver response curve of the deep-space spacecraft and the initial value of the orbital maneuver of the deep-space spacecraft.

[0126] In an application scenario, combined with Figure 1 , such as Figure 4 shown, the above S103 includes S1031 - S1033.

[0127] S1031. Adopt the golden section method to select the optimal correction coefficient that suppresses the divergence component within the value range of the correction coefficient of the deep-space spacecraft.

[0128] Optionally, the value range [κ low , κ up of the correction coefficient of the deep-space spacecraft can be [-0.03, 0.03], or can also be [-0.015, 0.015]. The embodiments of the present application do not limit the correction coefficient of the deep-space spacecraft.

[0129] The following describes the specific process of step S1031.

[0130] Step 1. Set the optimization index of the correction coefficient of the deep-space spacecraft.

[0131] The above optimization index is: the divergence component of the deep-space spacecraft after recursing for a period of time, and starting from the previous orbital maneuver as the starting time point, select the correction coefficient κ' of the deep-space spacecraft from [κ ori , κ low , κ up on the basis of Δv ori such that the magnitude of the applied orbital maneuver is (1 + κ')Δv ori . It can be understood that although the divergence component of the deep-space spacecraft always diverges at an exponential speed, the divergence speed is still different with different selections of κ'.

[0132] The mapping relationship of the above optimization index can be expressed as: F n : α 1,n refers to the divergence component of deep space spacecraft with different divergence speeds as κ' is selected differently.

[0133] Step 2: Adopt the golden section algorithm to search for the optimal κ within the range of [κ low , κ up to minimize F n and obtain the optimal correction coefficient κ for suppressing the divergence component of the deep space spacecraft. Since the golden section algorithm belongs to the commonly used technical means in this technical field, the specific process of the golden section algorithm is not further limited in the embodiments of the present application.

[0134] S1032: Calculate the maneuver overshoot coefficient of the deep space spacecraft.

[0135] The maneuver overshoot coefficient of the deep space spacecraft satisfies the following formula;

[0136] δ v = |δ v |·η δ

[0137] where δ v represents the maneuver overshoot coefficient of the deep space spacecraft, |δ v | represents the modulus of the maneuver overshoot coefficient of the deep space spacecraft, η δ represents the sign of the maneuver overshoot coefficient of the deep space spacecraft, η δ ∈{-1, 0, +1} and η δ = η v ·η MRC ·η α ; η v represents the direction of the previous orbital maneuver of the deep space spacecraft, η v ∈{-1, +1}; η MRC represents the sign of the component to be controlled of the deep space spacecraft, η MRC ∈{-1, 0, +1}; η α represents the adjustment direction of the component to be controlled of the deep space spacecraft for this orbital maneuver, η α ∈{-1, 0, +1}; the component to be controlled is the mode with the largest value among the non-divergent components of the deep space spacecraft.

[0138] It can be understood that the selection of the modulus |δ v | of the maneuver overshoot coefficient of the above deep space spacecraft is related to the considered error environment and model accuracy. The larger the error, the larger |δ v | is selected. For example, under the Earth-Moon CRTBP, it is recommended that the value is not less than 10-4 If a more reasonable determination of the value of |δ v | is required, it should be tried within a large range after selecting the three-body system and determining the error parameters, and the value at the optimal effect shall be taken as the standard.

[0139] The direction η of the last orbital maneuver of the deep-space spacecraft mentioned above v is solved as follows.

[0140] Taking the nominal orbit of the deep-space spacecraft as the halo orbit as an example, when the modulus of the divergence component α1 of the deep-space spacecraft reaches α 1,max , an orbital maneuver is applied to make α1 zero. Then, it is set that the orbital maneuver to make α1 zero from -α 1,max is positive, and vice versa, the orbital maneuver to make α1 zero from +α 1,max is negative. Therefore, when the direction of the last orbital maneuver is positive, η v = 1; when the direction of the last orbital maneuver is negative, η v = -1.

[0141] The solution process of the sign η of the component to be controlled of the above deep-space spacecraft MRC is as follows.

[0142] Step 1: Select the component to be controlled of the deep-space spacecraft.

[0143] Continuing to take the nominal orbit of the deep-space spacecraft as the halo orbit as an example, the numerical values of the non-divergence components α3, α4, α5, and α6 of the deep-space spacecraft are calculated through the above formula (2), and the non-divergence component with the largest numerical value among α3, α4, α5, and α6 is taken as the component to be controlled of the deep-space spacecraft.

[0144] Step 2: Predict the maneuver time of this orbital maneuver

[0145] Set this orbital maneuver as the jth orbital maneuver, then the predicted value of the maneuver time of the jth orbital maneuver satisfies the following formula.

[0146]

[0147] Among them, represents the predicted value of the maneuver time of the jth orbital maneuver, T peri represents the orbital period of the orbit in which the deep-space spacecraft operates. The subscripts of the parameters related to T in the above formula all represent the number of orbital maneuvers. The embodiments of the present application do not elaborate on the parameters related to T in the above.

[0148] Step 3: Take the value of the sign η of the component to be controlled of the deep-space spacecraft MRC for the value.

[0149] At the maneuvering moment of this orbital maneuver After that, according to the maneuvering moment of this orbital maneuver Determine the phase of the deep-space spacecraft on the orbit at this moment. Take the controlled component corresponding to this phase on the maneuvering response curve of the controlled component of the deep-space spacecraft (i.e., the controlled component at the maneuvering moment of this orbital maneuver) as η MRC The judgment basis for assignment.

[0150] Specifically, first assign a value to η MRC According to the positive or negative nature of the controlled component. When the controlled component is positive, η MRC = +1. When the controlled component is negative, η MRC = -1. Then, according to the modulus value of the controlled component, finally assign a value to η MRC When the modulus value of the controlled component is less than |α 1,max | / 20, η MRC = 0. Otherwise, η MRC remains unchanged.

[0151] The following is the solution process for the adjustment direction η α of the controlled component of the deep-space spacecraft for this orbital maneuver.

[0152] Referring to steps 1 and 2 in the above solution process of η MRC Obtain the controlled component a i at the maneuvering moment of this orbital maneuver. Take the above controlled component a i as the assignment basis for the adjustment direction η α of the controlled component of the deep-space spacecraft for this orbital maneuver. Specifically, first assign a value to η i According to the positive or negative nature of the above controlled component a α . When the above controlled component a i is positive, η α = -1. When the above controlled component a i is negative, η α = +1. Then, according to the value of the controlled component a i , finally assign a value to η α . When the controlled component satisfies -α i,max <α i <+α i,max , η α = 0. Otherwise, η α remains unchanged.

[0153] S1033. Calculate the final value of the orbital maneuver of the deep-space spacecraft.

[0154] The final value of the above-mentioned orbital maneuver can guide a deep-space spacecraft onto the nominal orbit of the deep-space spacecraft. The final value of the orbital maneuver of the above-mentioned deep-space spacecraft satisfies the following formula;

[0155] Δv = (1 + κ + δ v )Δv ori

[0156] where Δv represents the final value of the orbital maneuver of the deep-space spacecraft, κ represents the optimal correction coefficient of the divergence component of the deep-space spacecraft, and Δv ori represents the initial value of the orbital maneuver of the deep-space spacecraft.

[0157] In summary, in a control method for a deep-space spacecraft provided by an embodiment of the present application, the orbital deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit is divided into the divergence component and the non-divergence component of the deep-space spacecraft, and the maneuver response curve of the deep-space spacecraft indicating the non-divergence component is obtained through the orbit change after the deep-space spacecraft performs an orbital maneuver. Then, during the navigation of the spacecraft, according to the divergence component of the deep-space spacecraft, the orbit on which the deep-space spacecraft operates, and the nominal orbit of the deep-space spacecraft, the initial value of the orbital maneuver capable of suppressing the divergence component of the deep-space spacecraft is determined. Finally, combining the initial value of the orbital maneuver and the maneuver response curve of the deep-space spacecraft, the final value of the orbital maneuver of the spacecraft is determined. The final value of the orbital maneuver determined by the above method guides the deep-space spacecraft onto the nominal orbit of the deep-space spacecraft, can improve the control accuracy of the deep-space spacecraft, and thus better reduces the fuel consumption of the deep-space spacecraft during the long-term execution of scientific exploration tasks.

[0158] Correspondingly, an embodiment of the present application provides a control device for a deep-space spacecraft, as Figure 5 shown, including a response curve determination module 501, an initial value determination module 502, and a final value determination module 503.

[0159] Among them, the response curve determination module 501 is configured to determine the maneuver response curve of the deep-space spacecraft based on the orbit change of the deep-space spacecraft after orbit maneuver; wherein, the orbit in which the deep-space spacecraft operates deviates from the nominal orbit of the deep-space spacecraft; the maneuver response curve includes the non-divergent component of the deep-space spacecraft after orbit maneuver at each position of the orbit in which the deep-space spacecraft operates; the non-divergent component of the deep-space spacecraft is the deviation between the orbit in which the deep-space spacecraft operates and the nominal orbit that neither exponentially diverges nor converges; the divergent component of the deep-space spacecraft is the deviation between the orbit in which the deep-space spacecraft operates and the nominal orbit that exponentially diverges. For example, the response curve determination module 501 is configured to implement S101 of the above control method for the deep-space spacecraft.

[0160] The initial value determination module 502 is configured to determine the initial value of the orbit maneuver of the deep-space spacecraft according to the orbit in which the deep-space spacecraft operates, the nominal orbit of the deep-space spacecraft, and the divergent component of the deep-space spacecraft when the deep-space spacecraft deviates from the nominal orbit of the deep-space spacecraft during the navigation process. For example, the initial value determination module 502 is configured to implement S102 of the above control method for the deep-space spacecraft.

[0161] The final value determination module 503 is configured to determine the final value of the orbit maneuver of the deep-space spacecraft according to the maneuver response curve of the deep-space spacecraft and the initial value of the orbit maneuver of the deep-space spacecraft; the final value of the orbit maneuver guides the deep-space spacecraft to the nominal orbit. For example, the final value determination module 503 is configured to implement S103 of the above control method for the deep-space spacecraft.

[0162] Optionally, the final value determination module 503 is specifically configured to: adopt the golden section method to select the optimal correction coefficient for suppressing the divergent component of the deep-space spacecraft within the value range of the correction coefficient of the deep-space spacecraft. Calculate the maneuver overshoot coefficient of the deep-space spacecraft. Calculate the final value of the orbit maneuver of the deep-space spacecraft. For example, the final value determination module 503 is specifically configured to implement S1031-S1033 of the above control method for the deep-space spacecraft.

[0163] Each module of the above control device for the deep-space spacecraft can also be used to execute other steps in the above method embodiments. All relevant contents involved in the above method embodiments can be cited in the function descriptions of the corresponding functional modules, which will not be elaborated here.

[0164] An embodiment of the present application further 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 runs, the processor executes the computer instructions stored in the memory, so that the electronic device executes the method in the above embodiment. Among them, the processor can implement the above response curve determination module 501, initial value determination module 502, and final value determination module 503; the above memory can also be used to store the maneuver response curve of the deep space spacecraft, the orbit on which the deep space spacecraft operates, the nominal orbit, the initial value of the orbit maneuver, and the final value of the orbit maneuver, etc.

[0165] An embodiment of the present application further provides a computer-readable storage medium, which includes a computer program that, when running on a computer, executes the method described in the above embodiment.

[0166] An embodiment of the present application further provides a computer program product, which includes computer program instructions that, when running on a computer, execute the method described in the above embodiment.

[0167] Each embodiment in this specification is described in a progressive manner. The same or similar parts among the embodiments can be referred to each other, and the key point of each embodiment is to illustrate the differences from other embodiments.

[0168] Finally, it should be noted that: the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it; although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that: they can still modify the technical solutions described in the foregoing embodiments, or perform equivalent replacements on some of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present application.

Claims

1. A control method for a deep space spacecraft, characterized in that, Including: Based on the orbital changes of the deep-space spacecraft after orbital maneuver, determining the maneuver response curve of the deep-space spacecraft; wherein, there is a deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit of the deep-space spacecraft; the maneuver response curve includes the non-divergent component of the deep-space spacecraft after orbital maneuver at each position on the orbit on which the deep-space spacecraft operates; the non-divergent component of the deep-space spacecraft is the deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit that neither exponentially diverges nor converges; the divergent component of the deep-space spacecraft is the deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit that exponentially diverges; During the navigation process, when the deep-space spacecraft deviates from the nominal orbit of the deep-space spacecraft, determining the initial value of the orbital maneuver of the deep-space spacecraft according to the orbit on which the deep-space spacecraft operates, the nominal orbit of the deep-space spacecraft, and the divergent component of the deep-space spacecraft; Determining the final value of the orbital maneuver of the deep-space spacecraft according to the maneuver response curve of the deep-space spacecraft and the initial value of the orbital maneuver of the deep-space spacecraft; the final value of the orbital maneuver guides the deep-space spacecraft to the nominal orbit of the deep-space spacecraft.

2. The method according to claim 1, characterized in that, The maneuver response curve of the deep-space spacecraft satisfies the following formula; Among them, Δa i represents the change amount of the i-th mode in the non-divergent component of the deep-space spacecraft after orbit maneuver, where i represents the i-th mode of the deep-space spacecraft, and a 1,max represents the maneuvering threshold of the mode vector of the deep-space spacecraft on the first mode after orbit maneuver, |Π v,1 | represents the modulus value of the projection vector of the mode vector of the deep-space spacecraft on the first mode after orbit maneuver, and Π v,1 T represents the transpose of the projection vector of the mode vector of the deep-space spacecraft on the first mode, and Π v,i represents the projection vector of the mode vector of the deep-space spacecraft on the i-th mode after orbit maneuver.

3. The method according to claim 1 or 2, characterized in that, The determining the final value of the orbital maneuver of the deep-space spacecraft includes: Adopting the golden section method to select the optimal correction coefficient for suppressing the divergent component of the deep-space spacecraft within the value range of the correction coefficient of the deep-space spacecraft; Calculating the maneuver overshoot coefficient of the deep-space spacecraft, and the maneuver overshoot coefficient of the deep-space spacecraft satisfies the following formula; δ v = |δ v |·η δ Among them, δ v represents the maneuver overshoot coefficient of the deep space spacecraft, |δ v | represents the modulus of the maneuver overshoot coefficient of the deep space spacecraft, η δ represents the sign of the maneuver overshoot coefficient of the deep space spacecraft, and η δ = η v ·η MRC ·η α ; η v represents the direction of the last orbital maneuver of the deep space spacecraft, η MRC represents the sign of the component to be controlled of the deep space spacecraft; η α represents the adjustment direction of the component to be controlled of the deep space spacecraft for the current orbital maneuver; the component to be controlled is the mode with the largest value among the non-divergent components of the deep space spacecraft; Calculating the final value of the orbital maneuver of the deep-space spacecraft, and the final value of the orbital maneuver of the deep-space spacecraft satisfies the following formula; Δv = (1 + κ + δ v )Δv ori Among them, Δv represents the final value of the orbital maneuver of the deep-space spacecraft, κ represents the optimal correction coefficient of the divergence component of the deep-space spacecraft, and Δv ori represents the initial value of the orbital maneuver of the deep-space spacecraft.

4. The method according to claim 1, wherein The initial value of the orbital maneuver of the deep-space spacecraft satisfies the following formula; (δX + Δv ori )·Π = 0 where δX represents the orbital deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit, and Δv ori represents the initial value of the orbital maneuver of the deep-space spacecraft, and Π represents the projection vector of the modal vector of the deep-space spacecraft.

5. The method according to claim 1, wherein The nominal orbit is a libration point orbit.

6. A control device for a deep space spacecraft, characterized in that, Including a response curve determination module, an initial value determination module, and a final value determination module; The response curve determination module is configured to determine the maneuver response curve of the deep-space spacecraft based on the orbital changes of the deep-space spacecraft after orbital maneuver; wherein, there is a deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit of the deep-space spacecraft; the maneuver response curve includes the non-divergent component of the deep-space spacecraft after orbital maneuver at each position on the orbit on which the deep-space spacecraft operates; the non-divergent component of the deep-space spacecraft is the deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit that neither exponentially diverges nor converges; the divergent component of the deep-space spacecraft is the deviation between the orbit on which the deep-space spacecraft operates and the nominal orbit that exponentially diverges; The initial value determination module is configured to, when the deep space spacecraft deviates from the nominal orbit of the deep space spacecraft during navigation, determine the initial value of the orbital maneuver of the deep space spacecraft according to the orbit on which the deep space spacecraft operates, the nominal orbit of the deep space spacecraft, and the divergence component of the deep space spacecraft; The final value determination module is configured to determine the final value of the orbital maneuver of the deep space spacecraft according to the maneuver response curve of the deep space spacecraft and the initial value of the orbital maneuver of the deep space spacecraft; the final value of the orbital maneuver guides the deep space spacecraft to the nominal orbit.

7. An electronic device, characterized in that, It includes a processor and a memory coupled to the processor; the memory is used to store computer instructions, and when the electronic device runs, the processor executes the computer instructions stored in the memory, so that the electronic device executes the method according to any one of claims 1 to 5.

8. A computer-readable storage medium, characterized in that, It includes computer program instructions, and when the computer program instructions are executed by a computer, the computer is caused to execute the method according to any one of claims 1 to 5.

9. A computer program product, characterized in that, It includes computer program instructions, and when the computer program instructions run on a computer, the computer is caused to execute the method according to any one of claims 1 to 5.

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