Deep Space Exploration Rendezvous Guidance Method, Device, Computing Equipment and Storage Medium

Abstract

The present invention relates to the field of rendezvous guidance technology, and particularly relates to a rendezvous guidance method, device, computing device and storage medium for deep space exploration. Among them, the method is applied to the on-board guidance system of a tracking spacecraft, and includes: when the tracking spacecraft reaches the orbital position from the Earth to space, guiding according to the guidance instructions of the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbital position; performing autonomous remote guidance based on the pre-acquired orbital information of the remote anchor point and the nominal value of the near anchor point, so that the tracking spacecraft reaches the near anchor point from the remote anchor point; performing autonomous near guidance based on the nominal value of the near anchor point and the orbital information of the target spacecraft, so that the tracking spacecraft performs autonomous rendezvous and docking with the target spacecraft. In this solution, by inserting a remote anchor point and a near anchor point, the tracking spacecraft can directly dock with the target spacecraft from far to near after being launched into orbit from the Earth, which can greatly reduce the waste of time and fuel.

Description

Technical Field

[0001] The embodiments of the present invention relate to the technical field of rendezvous guidance, and particularly to a rendezvous guidance method, device, computing device and storage medium for deep space exploration. Background Art

[0002] With the continuous development of aerospace technology, various deep space exploration missions have emerged, such as Mars exploration missions, lunar exploration missions, etc. In deep space exploration missions, a tracking spacecraft launched from the Earth's surface often needs to perform a rendezvous and docking with a target spacecraft operating above a target celestial body (i.e., the celestial body to be explored).

[0003] However, after the tracking spacecraft separates from the rocket and enters the space orbit, the traditional rendezvous guidance method for deep space exploration is to guide the tracking spacecraft to land on the target celestial body through ground guidance, and then the tracking spacecraft takes off from the surface of the target celestial body and continuously approaches the target spacecraft located above the target celestial body to perform a rendezvous and docking. This traditional rendezvous guidance method for deep space exploration, which first guides to the surface of the target celestial body and then ascends from the surface of the target celestial body, is very time-consuming and fuel-consuming.

[0004] Therefore, there is an urgent need for a new rendezvous guidance method for deep space exploration. Summary of the Invention

[0005] In order to solve the problem that the traditional rendezvous guidance method for deep space exploration is relatively time-consuming and fuel-consuming, the embodiments of the present invention provide a rendezvous guidance method, device, computing device and storage medium for deep space exploration.

[0006] In a first aspect, the embodiments of the present invention provide a rendezvous guidance method for deep space exploration, which is applied to an on-board guidance system of a tracking spacecraft. The method includes:

[0007] When the tracking spacecraft reaches the orbital position from the Earth to space, it is guided according to the guidance command of the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbital position; wherein, the guidance command is determined according to the pre-determined orbital position and the orbital information of the remote anchor point;

[0008] Based on the pre-acquired orbital information of the remote anchor point and the nominal value of the near anchor point, autonomous remote guidance is performed to enable the tracking spacecraft to reach the near anchor point from the remote anchor point;

[0009] Based on the nominal value of the near anchor point and the orbital information of the target spacecraft, autonomous near guidance is performed to enable the tracking spacecraft to perform an autonomous rendezvous and docking with the target spacecraft; wherein, the target spacecraft orbits around the target celestial body.

[0010] Second aspect, an embodiment of the present invention further provides a rendezvous guidance device for deep space exploration, which is arranged in the on-board guidance system of a tracking spacecraft. The device includes:

[0011] A first guidance unit, configured to guide according to a guidance instruction of a ground control system when the tracking spacecraft reaches the orbital position from the Earth to space, so that the tracking spacecraft reaches a remote anchor point from the orbital position; wherein, the guidance instruction is determined according to a pre-determined orbital position and orbital information of the remote anchor point;

[0012] A second guidance unit, configured to perform autonomous remote guidance based on pre-acquired orbital information of the remote anchor point and nominal values of a near-range anchor point, so that the tracking spacecraft reaches the near-range anchor point from the remote anchor point;

[0013] A third guidance unit, configured to perform autonomous near-range guidance based on the nominal values of the near-range anchor point and orbital information of a target spacecraft, so that the tracking spacecraft performs autonomous rendezvous and docking with the target spacecraft; wherein, the target spacecraft orbits a target celestial body.

[0014] Third aspect, an embodiment of the present invention further provides a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, the method described in any embodiment of this specification is implemented.

[0015] Fourth aspect, an embodiment of the present invention further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed on a computer, the computer is made to execute the method described in any embodiment of this specification.

[0016] An embodiment of the present invention provides a rendezvous guidance method, device, computing device, and storage medium for deep space exploration, which are applied to the on-board guidance system of a tracking spacecraft. First, when the tracking spacecraft reaches the orbital position from the Earth to space, the on-board guidance system is guided according to the guidance instructions sent by the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbital position; wherein, the guidance instructions are determined according to the pre-determined orbital information of the orbital position and the remote anchor point; then, the on-board guidance system performs autonomous remote guidance based on the pre-acquired orbital information of the remote anchor point and the nominal value of the near-range anchor point, so that the tracking spacecraft reaches the near-range anchor point from the remote anchor point; finally, the on-board guidance system performs autonomous near-range guidance based on the nominal value of the near-range anchor point and the orbital information of the target spacecraft, so that the tracking spacecraft performs autonomous rendezvous and docking with the target spacecraft orbiting around the target celestial body. In this solution, by inserting a remote anchor point and a near-range anchor point, the tracking spacecraft can directly approach the target spacecraft step by step from far to near after being launched into orbit from the Earth and dock with it. Compared with the traditional method of first guiding to the surface of the target celestial body and then ascending from the surface of the target celestial body for rendezvous and docking, it can greatly reduce the waste of time and fuel. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0018] Figure 1 is a flowchart of a rendezvous guidance method for deep space exploration provided by an embodiment of the present invention;

[0019] Figure 2 is a schematic diagram of the spatial position of the near-range anchor point provided by an embodiment of the present invention;

[0020] Figure 3 is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0021] Figure 4 is a structural diagram of a rendezvous guidance device for deep space exploration provided by an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0022] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the scope of protection of the present invention.

[0023] As mentioned above, deep space exploration missions refer to exploration activities that break away from the Earth's gravity and enter the solar system space and cosmic space, such as Mars exploration missions, lunar exploration missions, etc. However, after the tracking spacecraft separates from the rocket and enters the space orbit, the traditional rendezvous guidance method for deep space exploration is to use ground guidance to make the tracking spacecraft land on the target celestial body, and then the tracking spacecraft takes off from the surface of the target celestial body and continuously approaches the target spacecraft located above the target celestial body for rendezvous and docking. This traditional deep space exploration rendezvous guidance method of first guiding to the surface of the target celestial body and then ascending from the surface of the target celestial body is very time-consuming and fuel-consuming.

[0024] To solve the above technical problems, the inventor can consider designing a remote anchor point and a near anchor point between the orbit insertion position and the target spacecraft, so that the tracking spacecraft can directly approach the target spacecraft step by step from far to near after being launched from the Earth into orbit, that is, first from the orbit insertion position to the remote anchor point, then from the remote anchor point to the near anchor point, and finally from the near anchor point to the target spacecraft to dock with it, so as to optimize time and fuel consumption.

[0025] The following describes the specific implementation of the above concept.

[0026] Please refer to Figure 1 , an embodiment of the present invention provides a rendezvous guidance method for deep space exploration, which is applied to the on-board guidance system of a tracking spacecraft. The method includes:

[0027] Step 100, when the tracking spacecraft reaches the orbit insertion position from the Earth to space, it is guided according to the guidance command of the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbit insertion position; wherein, the guidance command is determined according to the pre-determined orbit information of the orbit insertion position and the remote anchor point;

[0028] Step 102, based on the pre-acquired orbit information of the remote anchor point and the nominal value of the near anchor point, perform autonomous remote guidance, so that the tracking spacecraft reaches the near anchor point from the remote anchor point;

[0029] Step 104, based on the nominal value of the near anchor point and the orbit information of the target spacecraft, perform autonomous near guidance, so that the tracking spacecraft performs autonomous rendezvous and docking with the target spacecraft; wherein, the target spacecraft orbits around the target celestial body.

[0030] In the embodiment of the present invention, first, when the tracking spacecraft reaches the orbital position in space from the Earth, the on-board guidance system of the tracking spacecraft is guided according to the guidance command sent by the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbital position; wherein, the guidance command is determined according to the pre-determined orbital information of the orbital position and the remote anchor point; then, the on-board guidance system performs autonomous remote guidance based on the pre-acquired orbital information of the remote anchor point and the nominal value of the near anchor point, so that the tracking spacecraft reaches the near anchor point from the remote anchor point; finally, the on-board guidance system performs autonomous near guidance based on the nominal value of the near anchor point and the orbital information of the target spacecraft, so that the tracking spacecraft performs autonomous rendezvous and docking with the target spacecraft orbiting around the target celestial body. In this solution, by inserting the remote anchor point and the near anchor point, the tracking spacecraft can directly approach the target spacecraft step by step from far to near after being launched into orbit from the Earth and dock with it. Compared with the traditional method of first guiding to the surface of the target celestial body and then ascending from the surface of the target celestial body for rendezvous and docking, it can greatly reduce the waste of time and fuel.

[0031] Regarding step 100:

[0032] In the embodiment of the present invention, the rendezvous guidance process of deep space exploration is divided into three stages. From the orbital position to the remote anchor point is the ground navigation guidance stage, from the remote anchor point to the near anchor point is the on-board autonomous remote guidance stage, and from the near anchor point to the target spacecraft is the on-board autonomous near guidance stage. This step 100 is the ground navigation guidance stage, and this stage relies on ground orbit determination means. In order to improve the orbit measurement accuracy, the orbit determination time before each orbit maneuver is generally not less than 2 near celestial body orbital periods. The ground control system comprehensively calculates according to the pre-determined orbital information of the orbital position and the remote anchor point, fully considering the orbital semi-major axis element, eccentricity element, orbital inclination element, right ascension of the ascending node element, and the phase angle element of the tracking spacecraft relative to the target spacecraft, and then sends an orbit maneuver guidance command to the on-board guidance system of the tracking spacecraft. The on-board guidance system makes the tracking spacecraft reach the remote anchor point from the orbital position according to the guidance command.

[0033] In some embodiments, the orbital information of the remote anchor point is determined by the following method:

[0034] Based on the mission constraints and the capture orbit, determine the first constraint model;

[0035] Based on the time constraint from the remote anchor point to the near anchor point, determine the second constraint model;

[0036] Based on the effective measurement field of view of the on-board remote sensor, determine the third constraint model;

[0037] Determine the fourth constraint model based on the precision handshaking condition of ground guidance and on-orbit autonomous guidance;

[0038] Based on the first constraint model, the second constraint model, the third constraint model, and the fourth constraint model, determine the orbital information of the remote anchor point.

[0039] The following will separately explain each constraint model.

[0040] First constraint model: Based on the mission constraints and the capture orbit, the distance hp from the pericenter of the orbit where the remote anchor point is located to the center of the target celestial body can be determined.

[0041] In some embodiments, the second constraint model is:

[0042]

[0043] N 1 =∈[2, N max

[0044] In the formula, N 1 and N 2 are respectively the number of phase adjustment pulses and the number of precise adjustment pulses required from the remote anchor point to the near anchor point, t 1 and t 2 are respectively the orbital periods of the orbits where the phase adjustment pulses and the precise adjustment pulses are located, N max is the maximum value of the number of phase adjustment pulses, and t max is the time constraint.

[0045] In this embodiment, based on the time constraint t max from the remote anchor point to the near anchor point, the number of phase adjustment pulses N 1 required from the remote anchor point to the near anchor point can be determined, which is the second constraint model.

[0046] In some embodiments, the third constraint model is:

[0047]

[0048]

[0049] In the formula, S 1 is the spatial orientation of the remote anchor point relative to the target spacecraft, a T is the semi-major axis of the orbit of the target spacecraft, hp and ha are respectively the distance from the pericenter of the orbit where the remote anchor point is located to the center of the target celestial body and the distance from the apocenter of the orbit where the remote anchor point is located to the center of the target celestial body, θ is the phase angle difference of the remote anchor point relative to the target spacecraft, δ a is the effective measurement field of view of the on-orbit remote sensor, a c ​is the semi-major axis of the orbit where the remote anchor point is located, N 1 and N 2 are respectively the number of phase adjustment pulses and the number of precise adjustment pulses required from the remote anchor point to the near anchor point.

[0050] In this embodiment, after the remote anchor point, the on-board full-autonomous remote guidance and control are completed. Then, the measurement range characteristics of the remote sensor used in the remote guidance stage pose requirements for the orbital elements. Based on the effective measurement field of view of the on-board remote sensor, the phase angle difference θ between the remote anchor point and the target spacecraft, the distance hp from the perigee of the orbit where the remote anchor point is located to the center of the target celestial body, and the distance ha from the apogee to the center of the target celestial body can be determined, which is the third constraint model.

[0051] Fourth constraint model: Based on the accuracy handshake condition between ground guidance and on-board autonomous guidance, determine the number of phase adjustment pulses N 1 and the distance ha from the apogee of the orbit where the remote anchor point is located to the center of the target celestial body.

[0052] In this embodiment, during the ground navigation and guidance stage, the tracking spacecraft approaches the remote anchor point. According to the ground orbit determination accuracy, ground guidance strategy, and pulse execution accuracy, the accuracy range of the remote anchor point exit deviation can be determined as C 2 ; after the remote anchor point, it turns into the on-board autonomous remote guidance and control stage. The accuracy range of the entrance deviation that the on-board autonomous remote guidance and control stage can adapt to is C 1 , and the accuracies of the three stages of the ground navigation and guidance stage, the on-board autonomous remote guidance and control stage, and the on-board autonomous near guidance and control stage increase step by step. Therefore, the accuracy handshake condition between ground guidance and on-board autonomous guidance is C1 > C2, and C1 is positively correlated with the number of phase adjustment pulses N 1 and the distance ha from the apogee of the orbit where the remote anchor point is located to the center of the target celestial body. Therefore, based on the accuracy handshake condition between ground guidance and on-board autonomous guidance, the number of phase adjustment pulses N 1 and the distance ha from the apogee of the orbit where the remote anchor point is located to the center of the target celestial body can be constrained, which is the fourth constraint model.

[0053] After determining the four constraint models, first for ha, hp, N 1 , N 2 , t 1 and t 2Initialize the initial values and continuously adjust the initial values of these six parameters until the first constraint model, the second constraint model, the third constraint model, and the fourth constraint model are satisfied simultaneously. The orbital information of the remote anchor point can be obtained, that is, the remote anchor point satisfies the task constraints, time constraints, the effective measurement field of view of the on-board remote sensor, and the accuracy handshake conditions of ground guidance and on-board autonomous guidance, which can improve the safety and reliability of ground navigation and guidance and on-board autonomous remote guidance and control.

[0054] Regarding step 102:

[0055] During the on-board autonomous remote guidance and control phase, it depends on the on-board absolute orbit autonomous measurement means and adopts a combined guidance method with N 1 phase adjustment pulses for N 2 orbits and precise adjustment pulses for N

[0056] orbits to enable the tracking spacecraft to reach the near-range anchor point from the remote anchor point. The phase adjustment pulse guidance takes the relative phase information between the remote anchor point and the near-range anchor point as the input condition to complete the integrated adjustment of the orbital altitude and relative phase; the precise adjustment pulse can complete the joint adjustment of the orbital semi-major axis eccentricity and the argument of perigee based on the orbital information of the remote anchor point and the nominal value of the near-range anchor point. Figure 2 In some embodiments, the schematic diagram of the spatial position of the near-range anchor point in

[0057] can be referred to. The nominal value of the near-range anchor point is determined in the following manner:

[0058] Based on the effective measurement field of view of the on-board near-range sensor, determine the relative distance in the track direction between the near-range anchor point and the target spacecraft;

[0059] Based on the volume size of the target celestial body, determine whether to ignore the position-velocity coupling term;

[0060] In this embodiment, as Figure 2As shown, O is the target celestial body. When the deep space exploration mission is a lunar exploration mission, O is the moon; when the deep space exploration mission is a Mars exploration mission, O is Mars. Here, the target celestial body is not limited. P is the target spacecraft, which can be a spacecraft such as a space station, and A is the short-range anchor point. Among them, the relative distance x in the trace direction between the short-range anchor point and the target spacecraft can be determined based on the effective measurement field of view of the on-board short-range sensor. Since the volume and density of the target celestial body are different, the position and velocity coupling effects are also different. Therefore, before determining the radial relative position, trace velocity, and radial velocity of the short-range anchor point, it is necessary to determine whether to ignore the position-velocity coupling term. For example, in a Mars exploration mission, due to the relatively large volume and density of Mars, the position-velocity coupling term is relatively smaller, and at this time, the position-velocity coupling term can be ignored. While in a lunar exploration mission, since the moon is a small celestial body, the position-velocity coupling term needs to be considered to determine the radial relative position z, trace velocity, and radial velocity of the short-range anchor point. According to the volume of the target celestial body, it is determined whether to ignore the position-velocity coupling term. If it can be ignored, then the nominal value of the short-range anchor point can be quickly calculated.

[0061] In some embodiments, when the judgment result is not to ignore, the radial relative position, trace velocity, and radial velocity are calculated by the following formulas:

[0062]

[0063]

[0064]

[0065] In the formula, x is the relative distance in the trace direction between the short-range anchor point and the target spacecraft, z is the radial relative position of the short-range anchor point, r T is the distance between the target spacecraft and the center of the sphere of the target celestial body, r C is the distance between the orbit where the short-range anchor point is located and the center of the sphere of the target celestial body, μ is the celestial body gravitational coefficient, vx is the trace velocity of the short-range anchor point, and vz is the radial velocity of the short-range anchor point.

[0066] In some embodiments, when the judgment result is to ignore, the radial relative position, trace velocity, and radial velocity are calculated by the following formulas:

[0067]

[0068]

[0069] vz = 0

[0070] In the formula, x is the relative distance in the trace direction between the short-range anchor point and the target spacecraft, z is the radial relative position of the short-range anchor point, r T$r$ is the distance between the target spacecraft and the center of the celestial body, $\mu$ is the gravitational coefficient of the celestial body, $v_x$ is the trace velocity of the near-range anchor point, and $v_z$ is the radial velocity of the near-range anchor point.

[0071] In this embodiment, fully considering the field of view index and measurement accuracy index of the on-board near-range sensor, the near-range anchor points that meet the mission safety and reliability can be selected.

[0072] Regarding step 104:

[0073] The on-board autonomous near-range guidance and control stage relies on on-board relative orbit autonomous measurement means, including Hohmann transfer, CW guidance, line-of-sight guidance, and golden section phase plane control algorithm. Based on the nominal values of the near-range anchor points and the orbital information of the target spacecraft, the near-range autonomous rendezvous and docking mission is completed, realizing the rendezvous process from far to near in relative distance and from rough to precise in control.

[0074] As Figure 3 、 Figure 4 shown, the embodiment of the present invention provides a rendezvous guidance device for deep space exploration. The device embodiment can be implemented by software, or by hardware or a combination of software and hardware. From the hardware level, as Figure 3 shown, it is a hardware architecture diagram of a computing device where the rendezvous guidance device for deep space exploration provided by the embodiment of the present invention is located. In addition to Figure 3 the processor, memory, network interface, and non-volatile memory shown, the computing device where the device is located in the embodiment usually may also include other hardware, such as a forwarding chip responsible for processing packets, etc. Taking software implementation as an example, as Figure 4 shown, as a logically meaningful device, it is formed by the CPU of its computing device reading the corresponding computer program in the non-volatile memory into the memory for running. A rendezvous guidance device for deep space exploration provided by this embodiment is set in the on-board guidance system of the tracking spacecraft. The device includes:

[0075] The first guidance unit 401 is used to, when the tracking spacecraft reaches the orbital position from the earth to space, guide according to the guidance command of the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbital position; wherein, the guidance command is determined according to the pre-determined orbital information of the orbital position and the remote anchor point;

[0076] The second guidance unit 402 is used to perform autonomous remote guidance based on the pre-acquired orbital information of the remote anchor point and the nominal values of the near-range anchor points, so that the tracking spacecraft reaches the near-range anchor point from the remote anchor point;

[0077] The third guidance unit 403 is configured to perform autonomous short-range guidance based on the nominal values of the short-range anchor points and the orbital information of the target spacecraft, so as to enable the tracking spacecraft to perform autonomous rendezvous and docking with the target spacecraft; wherein, the target spacecraft orbits around the target celestial body.

[0078] In an embodiment of the present invention, it further includes a calculation unit 404 configured to calculate the orbital information of the long-range anchor points and the nominal values of the short-range anchor points in the first guidance unit 401, the second guidance unit 402, and the third guidance unit 403;

[0079] When calculating the orbital information of the long-range anchor points, the calculation unit 404 is configured to execute:

[0080] Determine the first constraint model based on the mission constraints and the capture orbit;

[0081] Determine the second constraint model based on the time constraint from the long-range anchor point to the short-range anchor point;

[0082] Determine the third constraint model based on the effective measurement field of view of the on-board long-range sensor;

[0083] Determine the fourth constraint model based on the accuracy handshake condition between ground guidance and on-board autonomous guidance;

[0084] Determine the orbital information of the long-range anchor points based on the first constraint model, the second constraint model, the third constraint model, and the fourth constraint model.

[0085] In an embodiment of the present invention, the second constraint model in the calculation unit 404 is:

[0086]

[0087] N 1 =∈[2, N max

[0088] In the formula, N 1 and N 2 are respectively the number of phase adjustment pulses and the number of precise adjustment pulses required from the long-range anchor point to the short-range anchor point, t 1 and t 2 are respectively the orbital periods of the orbits where the phase adjustment pulses and the precise adjustment pulses are located, N max is the maximum value of the number of phase adjustment pulses, and t max is the time constraint.

[0089] In an embodiment of the present invention, the third constraint model in the calculation unit 404 is:

[0090]

[0091] ​

[0092] In the formula, S 1 is the spatial azimuth of the remote anchor relative to the target spacecraft, a T is the semi-major axis of the target spacecraft's orbit, hp and ha are respectively the distances from the perigee of the orbit where the remote anchor is located to the center of the target celestial body and from the apogee to the center of the target celestial body, θ is the phase angle difference of the remote anchor relative to the target spacecraft, δ a is the effective measurement field of view of the on-board remote sensor, a c is the semi-major axis of the orbit where the remote anchor is located, N 1 and N 2 are respectively the number of phase adjustment pulses and the number of precise adjustment pulses required from the remote anchor to the near anchor.

[0093] In an embodiment of the present invention, when calculating the nominal value of the near anchor, the calculation unit 404 is used to execute:

[0094] Based on the effective measurement field of view of the on-board near sensor, determine the relative distance in the track direction between the near anchor and the target spacecraft;

[0095] Based on the volume size of the target celestial body, determine whether to ignore the position-velocity coupling term;

[0096] Based on the judgment result and the relative distance in the track direction, determine the radial relative position, track velocity, and radial velocity of the near anchor.

[0097] In an embodiment of the present invention, when the judgment result is not to ignore, the calculation unit 404 calculates the radial relative position, track velocity, and radial velocity through the following formula:

[0098]

[0099]

[0100]

[0101] In the formula, x is the relative distance in the track direction between the near anchor and the target spacecraft, z is the radial relative position of the near anchor, r T is the distance between the target spacecraft and the center of the target celestial body, r C is the distance between the orbit where the near anchor is located and the center of the target celestial body, μ is the celestial body gravitational coefficient, vx is the track velocity of the near anchor, and vz is the radial velocity of the near anchor.

[0102] In an embodiment of the present invention, when the judgment result is to ignore, the calculation unit 404 calculates the radial relative position, track velocity, and radial velocity through the following formula:

[0103]

[0104]

[0105] vz = 0

[0106] where x is the relative distance in the trace direction between the short-range anchor point and the target spacecraft, z is the relative radial position of the short-range anchor point, r T is the distance between the target spacecraft and the center of the celestial body, μ is the celestial gravitational coefficient, vx is the velocity in the trace direction of the short-range anchor point, and vz is the radial velocity of the short-range anchor point.

[0107] It can be understood that the structure illustrated in the embodiments of the present invention does not constitute a specific limitation on a deep space exploration rendezvous guidance device. In other embodiments of the present invention, a deep space exploration rendezvous guidance device may include more or fewer components than shown in the figures, or combine certain components, or split certain components, or have different component arrangements. The illustrated components can be implemented in hardware, software, or a combination of software and hardware.

[0108] Regarding the information interaction, execution process, etc. between the various modules within the above device, since it is based on the same concept as the method embodiments of the present invention, the specific content can be referred to the description in the method embodiments of the present invention and will not be elaborated here.

[0109] The embodiments of the present invention further provide a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, a deep space exploration rendezvous guidance method in any embodiment of the present invention is implemented.

[0110] The embodiments of the present invention further provide a computer-readable storage medium. A computer program is stored on the computer-readable storage medium. When the computer program is executed by a processor, the processor is enabled to execute a deep space exploration rendezvous guidance method in any embodiment of the present invention.

[0111] Specifically, a system or device equipped with a storage medium can be provided. A software program code for implementing the functions in any one of the above embodiments is stored on the storage medium, and the computer (or CPU or MPU) of the system or device is enabled to read and execute the program code stored in the storage medium.

[0112] In this case, the program code read from the storage medium itself can implement the functions of any one of the above embodiments. Therefore, the program code and the storage medium storing the program code constitute a part of the present invention.

[0113] Examples of storage media for providing program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Optionally, the program code can be downloaded from a server computer via a communication network.

[0114] In addition, it should be clear that not only can the functions of any of the above embodiments be achieved by executing the program code read by a computer, but also by causing an operating system or the like operating on the computer based on the instructions of the program code to complete part or all of the actual operations.

[0115] Furthermore, it can be understood that the program code read from the storage medium is written into the memory provided in the expansion board inserted into the computer or into the memory provided in the expansion module connected to the computer, and then based on the instructions of the program code, the CPU or the like installed on the expansion board or the expansion module is caused to execute part or all of the actual operations, thereby achieving the functions of any of the above embodiments.

[0116] It should be noted that in this article, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the term "comprising", "including" or any other variant thereof is intended to cover non-exclusive inclusion, so that a process, method, article or device comprising a series of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article or device.

[0117] Those of ordinary skill in the art can understand that all or part of the steps for implementing the above method embodiments can be completed by hardware related to program instructions. The foregoing program can be stored in a computer-readable storage medium. When the program is executed, it performs the steps including the above method embodiments; and the foregoing storage medium includes various media such as ROM, RAM, magnetic disks, or optical disks that can store program code.

[0118] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them; 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 for 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 invention.

Claims

1. A rendezvous guidance method for deep space exploration, characterized in that, it is applied to the on-board guidance system of a tracking spacecraft, and includes: When the tracking spacecraft reaches the orbital position from the Earth to space, it is guided according to the guidance instructions of the ground control system, so that the tracking spacecraft reaches the remote anchor point from the orbital position; wherein, the guidance instructions are determined according to the pre-determined orbital position and the orbital information of the remote anchor point; Based on the pre-acquired orbital information of the remote anchor point and the nominal value of the near anchor point, autonomous remote guidance is performed to make the tracking spacecraft reach the near anchor point from the remote anchor point; Based on the nominal value of the near anchor point and the orbital information of the target spacecraft, autonomous near guidance is performed to enable the tracking spacecraft to perform autonomous rendezvous and docking with the target spacecraft; wherein, the target spacecraft orbits around the target celestial body; The orbital information of the remote anchor point is determined by the following method: Based on the mission constraints and the capture orbit, a first constraint model is determined; Based on the time constraint from the remote anchor point to the near anchor point, a second constraint model is determined; Based on the effective measurement field of view of the on-board remote sensor, a third constraint model is determined; Based on the accuracy handshake condition between ground guidance and on-board autonomous guidance, a fourth constraint model is determined; Based on the first constraint model, the second constraint model, the third constraint model and the fourth constraint model, the orbital information of the remote anchor point is determined; The second constraint model is: N 1 = ∈ [2, N max ​ Where N 1 and N 2 are the number of phase adjustment pulses and the number of fine adjustment pulses required from the remote anchor point to the near - range anchor point respectively, t 1 and t 2 are the operating periods of the orbits where the phase adjustment pulses and the fine adjustment pulses are located respectively, N max is the maximum value of the number of phase adjustment pulses, and t max is the time constraint.

2. The method according to claim 1, characterized in that, The third constraint model is: where S 1 is the spatial orientation of the remote anchor relative to the target spacecraft, a T is the semi-major axis of the target spacecraft's orbit, hp and ha are the distances from the perigee of the orbit where the remote anchor is located to the center of the target celestial body and from the apogee to the center of the target celestial body, respectively, θ is the phase angle difference between the remote anchor and the target spacecraft, δ a is the effective measurement field of view of the on-board remote sensor, a c is the semi-major axis of the orbit where the remote anchor is located, N 1 and N 2 are the number of phase adjustment pulses and the number of precise adjustment pulses required from the remote anchor to the near anchor, respectively.

3. The method according to any one of claims 1-2, characterized in that, The nominal value of the near anchor point is determined by the following method: Based on the effective measurement field of view of the on-board near sensor, the relative distance in the track direction between the near anchor point and the target spacecraft is determined; Based on the volume of the target celestial body, it is judged whether to ignore the position-velocity coupling term; Based on the judgment result and the relative distance in the track direction, the relative radial position, the track velocity and the radial velocity of the near anchor point are determined.

4. The method according to claim 3, characterized in that, When the judgment result is not to ignore, the relative radial position, the track velocity and the radial velocity are calculated by the following formula: Wherein, x is the relative distance in the trace direction between the short-range anchor point and the target spacecraft, z is the relative radial position of the short-range anchor point, and r T is the distance between the target spacecraft and the center of the celestial body, and r C is the distance between the orbit where the short-range anchor point is located and the center of the celestial body, μ is the gravitational coefficient of the celestial body, VX is the trace direction velocity of the short-range anchor point, and VZ is the radial velocity of the short-range anchor point.

5. The method according to claim 3, characterized in that, When the judgment result is to ignore, the relative radial position, the track velocity and the radial velocity are calculated by the following formula: vz = 0 where x is the relative distance in the track direction between the short-range anchor point and the target spacecraft, z is the relative radial position of the short-range anchor point, and r T is the distance between the target spacecraft and the center of the celestial body, μ is the celestial gravitational coefficient, vx is the velocity in the track direction of the short-range anchor point, and vz is the radial velocity of the short-range anchor point.

6. A rendezvous guidance device for deep space exploration, used to implement the method according to any one of claims 1-5, characterized in that, It is arranged in the on-board guidance system of the tracking spacecraft and includes: The first guidance unit is used to guide according to the guidance instructions of the ground control system when the tracking spacecraft reaches the orbital position from the Earth to space, so that the tracking spacecraft reaches the remote anchor point from the orbital position; wherein, the guidance instructions are determined according to the pre-determined orbital position and the orbital information of the remote anchor point; A second guidance unit, configured to perform autonomous remote guidance based on the pre-acquired orbital information of the remote anchor point and the nominal value of the short-range anchor point, so that the tracking spacecraft reaches the short-range anchor point from the remote anchor point; A third guidance unit, configured to perform autonomous short-range guidance based on the nominal value of the short-range anchor point and the orbital information of the target spacecraft, so that the tracking spacecraft performs autonomous rendezvous and docking with the target spacecraft; wherein, the target spacecraft orbits around a target celestial body.

7. A computing device, comprising a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, the method according to any one of claims 1-5 is implemented.

8. A computer-readable storage medium, on which a computer program is stored, and when the computer program is executed on a computer, the computer is made to execute the method according to any one of claims 1-5.

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