Thruster layout design method and device for micro-nano satellite rendezvous and docking

Abstract

The present invention provides a thruster layout design method, device, and readable medium for micro-nano satellite rendezvous and docking, addressing the problem that existing layout design methods are unsuitable for micro-nano satellites and can generate significant interference forces on orbits during attitude control. The method includes constructing a satellite body coordinate system; determining the positions and thrust directions of four orbit control thrusters in the satellite body coordinate system based on a first constraint; selecting a first mounting surface and a second mounting surface from the satellite body coordinate system, wherein the first mounting surface and the second mounting surface are symmetrical about the satellite body; determining the positions and thrust directions of a first group of attitude control thrusters on the first mounting surface and a second group of attitude control thrusters on the second mounting surface based on a second constraint, wherein the second group of attitude control thrusters are symmetrically mounted with the first group of attitude control thrusters in a force couple configuration; wherein the first constraint is that the combined force of each of the four orbit control thrusters passes through the satellite's center of mass, and the second constraint is that the thrust generated by the attitude control thrusters does not pass through the satellite's center of mass.

Description

Technical Field

[0001] The present invention relates to the field of space rendezvous and docking, and in particular to a thruster layout design method, device and readable medium for micro-nano satellite rendezvous and docking. Background Art

[0002] Rendezvous and docking involves the simultaneous rendezvous and connection of two spacecraft at the same orbital position and speed, forming a single, integrated unit. This process places high demands on the spacecraft's six-degree-of-freedom maneuverability and control accuracy in both orbit and attitude, as well as the reliability of the attitude and orbit control system. As the primary actuator in the control system of such service-oriented spacecraft, the thruster's configuration and layout directly determine whether the spacecraft can meet these requirements and are a key factor influencing the success of the rendezvous and docking mission.

[0003] Currently, spacecraft that have successfully conducted rendezvous and docking missions in orbit typically have between 24 and 40 thrusters. For micro- and nanosatellites with small size and high functional density, directly adopting this thruster configuration and layout scheme presents significant engineering challenges. Furthermore, existing theoretical thruster layout methods based on feasible configuration matrices place all thrusters upright along the spacecraft's main axis. This creates significant interference forces on the orbit during attitude control, requiring counter-force injection from thrusters passing through the center of mass in opposite directions to offset the interference. This results in high propellant consumption and lacks resilience to thruster failures.

[0004] Based on this, the inventors of the present application have designed a thruster layout design method and device for micro-nano satellite rendezvous and docking, in order to solve the above technical problems. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to provide a thruster layout design method, device and readable medium for micro-nano satellite rendezvous and docking, so as to solve the problem that the existing layout design method is not suitable for micro-nano satellites and will generate a large interference force on the orbit during attitude control.

[0006] To solve the above technical problems, the present invention provides a thruster layout design method for micro-nano satellite rendezvous and docking, comprising: constructing a satellite body coordinate system, wherein the origin of the satellite body coordinate system is located at the satellite's center of mass and the three axes of the satellite body coordinate system are parallel and in the same direction as the axes of the orbital coordinate system when the satellite is in an earth-oriented attitude; determining the positions and thrust directions of four orbital control thrusters in the satellite body coordinate system according to a first constraint condition; selecting a first mounting surface and a second mounting surface from the satellite body coordinate system, wherein the first mounting surface and the second mounting surface are symmetrical about the satellite body; determining the positions and thrust directions of a first group of attitude control thrusters on the first mounting surface and the positions and thrust directions of a second group of attitude control thrusters on the second mounting surface according to a second constraint condition, wherein the second group of attitude control thrusters are symmetrically mounted with the first group of attitude control thrusters in the form of a force couple;

[0007] Among them, the first constraint condition is that the combined force of the four orbit control thrusters in pairs passes through the satellite's center of mass, and the second constraint condition is that the thrust generated by the attitude control thrusters does not exceed the satellite's center of mass.

[0008] Optionally, the first group of attitude control thrusters and the second group of attitude control thrusters each include four attitude control thrusters.

[0009] Optionally, determining the positions and thrust directions of the four orbital control thrusters in the satellite body coordinate system according to the first constraint condition includes: determining the plane in which the thrust axes of the four orbital control thrusters are located according to the two directions with the largest speed increments required for the rendezvous and docking mission; and constructing the position equation and thrust equation of each orbital control thruster on the plane according to the first constraint condition.

[0010] Optionally, for a rendezvous and docking mission within the same orbital plane, the two directions with the largest required velocity increments are set as the X-axis and Z-axis of the satellite body coordinate system. The position equation and thrust equation of each orbit control thruster are:

[0011] C9=[X g 0 Z g ], f9=[-sin(α) 0 -cos(α)]×f g

[0012] C 10 =[X g 0-Z g ], f 10 =[-sin(α) 0 cos(α)]×f g

[0013] C 11 =[-X g 0-Z g ], f 11 =[sin(α) 0 cos(α)]×fg

[0014] C 12 =[-X g 0 Z g ], f 12 =[sin(α) 0 -cos(α)]×f g

[0015] Among them, C9~C 12 are the positions of the four orbital control thrusters, f9~f 12 is the thrust of the four orbit control thrusters, f g is the thrust of a single orbital control thruster, and α is the angle between the thrust direction of the +Z-end orbital control thruster and the +Z axis.

[0016] Optionally, determining the position and thrust direction of the first group of attitude control thrusters on the first mounting surface based on the second constraint condition includes: constructing a position equation and a thrust equation for each attitude control thruster in the first group of attitude control thrusters on the first mounting surface based on the second constraint condition, so that the thrust component of each attitude control thruster in the k direction is not zero and can meet the k-direction orbital control requirement required for the rendezvous and docking mission, where the k direction is the direction with the smallest orbital control amount required for the rendezvous and docking mission.

[0017] Optionally, the position equation and thrust equation of each attitude control thruster in the first group of attitude control thrusters are:

[0018] C1=[X z Y z Z z ], f1=[-sin(γ)cos(β) -sin(γ)sin(β) -cos(γ)]×f z

[0019] C2=[X z -Y z Z z ], f2=[-sin(γ)cos(β) sin(γ)sin(β) -cos(γ)]×f z

[0020] C3=[X z -Y z -Z z ], f3=[-sin(γ)cos(β) sin(γ)sin(β) cos(γ)]×f z

[0021] C4=[X z Y z -Z z], f4=[-sin(γ)cos(β) -sin(γ)sin(β) cos(γ)]×f z

[0022] Among them, C1~C4 are the installation positions of the four attitude control thrusters, f1~f4 are the thrusts of the four attitude control thrusters, fz is the thrust size of a single attitude control thruster, β is the angle between the projection of the first group of attitude control thrusters on the XY plane and +X, and γ is the angle between the +Z end attitude control thruster vector and +Z.

[0023] Optionally, for a rendezvous and docking mission within the same orbital plane, the k direction is the Y-axis direction of the satellite body coordinate system.

[0024] Optionally, the first group of attitude control thrusters and the second group of attitude control thrusters form a dual-point backup with each other.

[0025] Optionally, the first group of attitude control thrusters and the second group of attitude control thrusters can serve as backup thrusters when any orbit control thruster fails.

[0026] To solve the above technical problems, the present invention provides a thruster layout design device for micro-nano satellite rendezvous and docking, comprising: a processor and a memory, wherein the memory stores programs or instructions that can be run on the processor, and the programs or instructions are executed by the processor to implement the method described above.

[0027] To solve the above technical problem, the present invention provides a computer-readable medium storing computer program code, which implements the above method when executed by a processor.

[0028] The positive progress effect of the present invention is:

[0029] The thruster layout design method, device, and readable medium for micro-nano satellite rendezvous and docking of the present invention achieve ideal zero-force interference in satellite orbit control by symmetrically installing a first set of attitude control thrusters and a second set of attitude control thrusters in a couple configuration. That is, when the thrust magnitude and thrust direction deviation are not considered, the thruster output during attitude control is pure torque, and the output resultant force is zero. This avoids the problems encountered in existing technologies of generating large interference forces on the orbit during attitude control, requiring opposite-direction through-center thrusters to counteract and offset the interference, and resulting in high propellant consumption. The present invention reduces the number of thrusters required for a micro-nano satellite to complete rendezvous and docking missions to 12, effectively adapting to the structural installation constraints of micro-nano satellites with small size and high functional density. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] The above and other features, properties and advantages of the present invention will become more apparent through the following description in conjunction with the accompanying drawings and embodiments, in which:

[0031] Figure 1 A flow chart of a thruster layout design method for micro-nano satellite rendezvous and docking according to the present invention;

[0032] Figure 2 yes Figure 1 Flowchart of an embodiment of step S12;

[0033] Figure 3 (a) to (b) are schematic diagrams of the installation of four orbit control thrusters in the satellite body coordinate system;

[0034] Figure 4 (a) to (c) are schematic diagrams of the installation of the first set of attitude control thrusters in the satellite body coordinate system;

[0035] Figure 5 (a) to (c) are schematic diagrams of the installation of the second set of attitude control thrusters in the satellite body coordinate system;

[0036] Figure 6 (a) to (d) are schematic diagrams of thruster layouts suitable for rendezvous and docking of micro- and nano-satellites;

[0037] Figure 7 It is a system block diagram of a thruster layout design device according to an embodiment of the present application. DETAILED DESCRIPTION

[0038] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, specific embodiments of the present invention are described in detail below with reference to the accompanying drawings.

[0039] Embodiments of the present invention will now be described in detail with reference to the accompanying drawings. Reference will now be made in detail to preferred embodiments of the present invention, examples of which are illustrated in the accompanying drawings. Wherever possible, the same reference numerals will be used throughout the drawings to represent the same or similar parts. In addition, although the terms used in the present invention are selected from commonly known and commonly used terms, some of the terms mentioned in this specification may be selected by the applicant at his or her discretion, and their detailed meanings are explained in the relevant parts of the description herein. Furthermore, it is required that the present invention be understood not only by the actual terms used, but also by the meaning implied by each term.

[0040] Existing large-scale rendezvous and docking spacecraft, such as manned and cargo spacecraft, both internationally and domestically, typically have their orbital control engines and attitude control thrusters mounted upright along the spacecraft's main axis. This arrangement offers limited controllability and poor resilience to thruster failures. To address these shortcomings, most have adopted measures such as increasing the number of thrusters and implementing cold backup redundancy. Currently, spacecraft that have successfully conducted rendezvous and docking missions in orbit have between 24 and 40 thrusters. For micro- and nano-satellites with small size and high functional density, directly adopting this thruster configuration and layout scheme presents significant challenges in engineering practice. From a controllability perspective, some researchers have investigated theoretical thruster layout methods based on feasible configuration matrices. This layout method requires complex matrix construction and optimization processes. Furthermore, due to the vertical placement of all thrusters along the spacecraft's main axis, the resulting design generates significant disturbance forces on the orbit during attitude control, requiring counter-force injection from trans-center-of-mass thrusters in opposite directions to offset the disturbance forces. This results in high propellant consumption and lacks resilience to thruster failures. The purpose of this invention is to propose a thruster layout design method that adapts to the installation constraints of micro-nano satellites and meets the requirements of space missions for micro-nano satellite rendezvous and docking. The principle of this invention is to use two groups of eight thrusters installed in a couple configuration with thrust forces that do not exceed the center of mass, plus four additional thrusters with combined forces that exceed the center of mass. These thrusters form a multiple backup relationship, reliably providing six-degree-of-freedom control of the orbit and attitude of micro-nano satellite rendezvous and docking, and delivering pure torque output with ideal zero-force interference, thus facilitating the successful implementation of space rendezvous and docking missions requiring high-precision orbital control.

[0041] Figure 1 The figure is a flow chart of the thruster layout design method for micro-nano satellite rendezvous and docking according to the present invention. Figure 1 As shown, the thruster layout design method 100 includes the following steps:

[0042] Step S11: Constructing the satellite body coordinate system.

[0043] Construct satellite body coordinate system O b -X b Y b Z b , where the origin O b Located at the satellite's center of mass, the three axes of the satellite's body coordinate system are parallel and in the same direction as the axes of the orbital coordinate system when in the Earth-oriented attitude.

[0044] Step S12: determining the positions and thrust directions of the four orbit control thrusters in the satellite body coordinate system according to a first constraint condition, wherein the first constraint condition is that the combined force of each of the four orbit control thrusters passes through the satellite's center of mass.

[0045] Figure 2 yes Figure 1 Flowchart of step S12 in the embodiment. Figure 2 As shown, step S12 includes:

[0046] Step S121: determining the planes where the thrust axes of the four orbital control thrusters are located based on the two directions with the largest speed increments required for the rendezvous and docking mission;

[0047] Step S122: constructing the position equation and thrust equation of each orbit control thruster on the plane according to the first constraint condition, and obtaining the position and thrust direction of the four orbit control thrusters in the satellite body coordinate system.

[0048] The following steps S121 to S122 are described in detail:

[0049] The satellite body coordinate system O b -X b Y b Z b As the reference system, the thrust axis planes of the four orbital control thrusters are determined according to the two directions (i, j) with the largest velocity increment required for the rendezvous and docking mission, ensuring that the resultant forces in the four directions i, j, -i, and -j can be generated and the first constraint condition is satisfied, that is, the resultant forces in each direction pass through the center of mass. Among them, the combination (i, j) belongs to (X b ,Y b )、(X b ,Z b ) and (Y b ,Z b ), that is, (i, j)∈{(X b ,Y b ), (X b ,Z b ), (Y b ,Z b )}.

[0050] Figure 3 (a) to (b) are schematic diagrams of the installation of four orbit control thrusters in the satellite body coordinate system. Figure 3 As shown in (a), the four orbit control thrusters are numbered F9, F10, F11 and F12. Figure 3 As shown in (b), for rendezvous and docking missions within the same orbital plane, the two directions (i, j) with the largest required velocity increments are set to the X-axis and Z-axis of the satellite's coordinate system. In other words, the two directions with the largest required velocity increments are (i, j) = (Xb, Zb). The triangles next to the numbers represent orbital control thrusters. The arrows in the middle of the triangles represent the thrust f generated by the orbital control thrusters. m (m=9,10,11,12) thrust direction. Among them, F9 orbit control thruster and F12 orbit control thruster are at +Z b Axis direction, called +Z bThe F10 and F11 orbital control thrusters are at -Z. b Axis direction, called -Z b End orbit control thruster. For +Z b For the end orbit control thruster, the thrust direction of F9 orbit control thruster is the same as +Z b The angle between the axes is α, and the thrust direction of the F12 orbit control thruster is +Z b The angle between the axes is also α. b End orbit control thruster, the thrust direction of F10 orbit control thruster is the same as -Z b The angle between the axes is α, and the thrust direction of the F11 orbit control thruster is -Z b The angle between the axes is also α. The value of α ranges from (0° to 90°), and the value of α can be set according to the orbital control requirements of the rendezvous and docking mission.

[0051] According to the first constraint condition, the position equation and thrust equation of each orbital control thruster are constructed in the Xb-Zb plane as follows:

[0052] C9=[X g 0 Z g ], f9=[-sin(α) 0 -cos(α)]×f g

[0053] C 10 =[X g 0-Z g ], f 10 =]-sin(α) 0 cos(α)]×f g

[0054] C 11 =[-X g 0-Z g ], f 11 =[sin(α) 0 cos(α)]×f g

[0055] C 12 =[-X g 0 Z g ], f 12 =[sin(α) 0 -cos(α)]×f g

[0056] Among them, C9~C 12 are the positions of the four orbital control thrusters, f9~f 12 is the thrust of the four orbit control thrusters in the satellite body coordinate system, f g is the thrust of a single orbital thruster, α is +Z b The thrust direction of the end orbit control thruster is related to +Zb The angle between the axes.

[0057] Step S13: Selecting a first mounting surface and a second mounting surface from the satellite body coordinate system, wherein the first mounting surface and the second mounting surface are symmetrical about the satellite body.

[0058] Optionally, the first group of attitude control thrusters and the second group of attitude control thrusters each include four attitude control thrusters. In other words, the eight attitude control thrusters are divided into two groups, each with four. Two symmetry planes m and -m of the star are selected for the installation of the first group of attitude control thrusters and the second group of attitude control thrusters, where m∈{ Xb For the convenience of description, it is assumed that the first set of thrusters is installed on the m-plane, with thrusters numbered F1 to F4, and the second set of thrusters is installed on the -m-plane, with thrusters numbered F5 to F8.

[0059] Step S14: Determine the position and thrust direction of the first set of attitude control thrusters on the first mounting surface and the position and thrust direction of the second set of attitude control thrusters on the second mounting surface based on the second constraint. The second set of attitude control thrusters is symmetrically mounted with the first set of attitude control thrusters in a couple configuration. The second constraint is that the thrust generated by the attitude control thrusters does not exceed the satellite's center of mass.

[0060] Optionally, step S14 includes constructing position equations and thrust equations for each attitude control thruster within the first group of attitude control thrusters on the first mounting surface based on the second constraint, such that the thrust component of each attitude control thruster in the k-direction is non-zero and satisfies the k-direction orbital control requirements for the rendezvous and docking mission, where the k-direction is the direction of minimal orbital control required for the rendezvous and docking mission. Other considerations regarding thrust direction and thruster mounting positions are consistent with those for conventional four-slant-mounted attitude control thrusters, namely, meeting attitude control torque requirements, satisfying constraints such as the satellite structural installation size, and eliminating plume effects.

[0061] For example, when designing the thrust direction of one group of thrusters (such as the first group of thrusters), it is ensured that the component force of each thruster in the k direction is not zero and can meet the k-direction orbit control requirements required for the rendezvous and docking mission, where k = (X b ,Y b ,Z b )-(i, j) is the direction that minimizes the amount of orbital control required for the rendezvous and docking mission. Other considerations regarding thrust direction and thruster installation location are consistent with those for traditional four-slant-mounted attitude control thrusters, ensuring they meet attitude control torque requirements, meet satellite structural installation dimensions, and avoid plume impacts.

[0062] Figure 4 (a) to (c) are schematic diagrams of the installation of the first set of attitude control thrusters in the satellite body coordinate system. Taking the first set of attitude control thrusters installed on the +X plane of the satellite as an example, the first set of 4 attitude control thrusters F1 to F4 are installed as follows: Figure 4(a) As shown. For the rendezvous and docking mission within the same orbital plane, the k direction is the Y-axis direction of the satellite's coordinate system. In other words, for the rendezvous and docking mission within the same orbital plane, the direction with the minimum orbital control amount is k = (X b ,Y b ,Z b )-(X b ,Z b )=Y b .like Figure 4 (b) and Figure 4 As shown in (c), at this time, according to the second constraint, that is, the thrust generated by the attitude control thruster does not exceed the satellite's center of mass, at X b Y b The position equation and thrust equation of each attitude control thruster in the first group of attitude control thrusters constructed on the surface are:

[0063] C1=[X z Y z Z z ], f1=[-sin(γ)cos(β) -sin(γ)sin(β) -cos(γ)]×f z

[0064] C2=[X z -Y z Z z ], f2=[-sin(γ)cos(β) sin(γ)sin(β) -cos(γ)]×f z

[0065] C3=[X z -Y z -Z z ], f3=[-sin(γ)cos(β) sin(γ)sin(β) cos(γ)]×f z

[0066] C4=[X z Y z -Z z ], f4=[-sin(γ)cos(β) -sin(γ)sin(β) cos(γ)]×f z

[0067] Among them, C1~C4 are the installation positions of F1~F4 attitude control thrusters, f1~f4 are the thrusts of F1~F4 attitude control thrusters, and f zis the thrust of a single attitude control thruster, β is the angle between the projection of the first group of attitude control thruster axes on the XY plane and +X, and γ is the angle between the +Z end attitude control thruster vector and +Z. β and γ can be set by comprehensively considering the requirements of attitude control torque, the installation size of the satellite structure, and the absence of plume influence.

[0068] Figure 4 (b) and Figure 4 (c) f 1T is the thrust f1 at X b Y b Projection on the surface, f 2T is the thrust f2 at X b Y b Projection on the surface, f 3T is the thrust f3 at X b Y b Projection on the surface, f 4T The thrust f4 is in X b Y b Projection on the surface.

[0069] Figure 5 Figures (a) to (c) are schematic diagrams of the second set of attitude control thrusters installed in the satellite's coordinate system. Since the second set of attitude control thrusters F5 to F8 are installed symmetrically with the first set of attitude control thrusters F1 to F4 in a couple arrangement, the position equations and thrust equations for each attitude control thruster in the second set are:

[0070] C5=[-X z Y z Z z ], f5=[sin(γ)cos(β) -sin(γ)sin(β) -cos(γ)]×f z

[0071] C6=[-X z -Y z Z z ], f6=[sin(γ)cos(β) sin(γ)sin(β) -cos(γ)]×f z

[0072] C7=[-X z -Y z -Z z ], f7=[sin(γ)cos(β) sin(γ)sin(β) cos(γ)]×f z

[0073] C8=[-X z Y z -Z z], f8=[sin(γ)cos(β) -sin(γ)sin(β) cos(γ)]×f z

[0074] Among them, C5~C8 are the installation positions of F5~F8 attitude control thrusters, f5~f8 are the thrusts of F5~F8 attitude control thrusters, and f z is the thrust of a single attitude control thruster, β is the angle between the projection of the first group of attitude control thrusters on the XY plane and +X, and γ is the angle between the +Z end attitude control thruster vector and +Z.

[0075] At this point, the thruster layout suitable for micro-nano satellite rendezvous and docking is completed. Figure 6 (a) to (d) are schematic diagrams of thruster layouts suitable for rendezvous and docking of micro-nano satellites. Figure 6 As shown in (a), the present invention reduces the number of thrusters required for a micro-nano satellite to complete the rendezvous and docking mission to 12, which can better adapt to the structural installation constraints of micro-nano satellites with small size and high functional density. Figure 6 As shown in (b) to (c), the attitude control thrusters F1 to F4 and the attitude control thrusters F5 to F8 are symmetrically installed in the form of a couple, which can achieve ideal zero-force interference in satellite orbit control. That is, when the thrust size and thrust direction deviation are not considered, the thruster output is pure torque during attitude control, and the output resultant force is 0. This avoids the problem in the existing technology that a large interference force is generated on the orbit during attitude control, and the opposite direction of the center of mass thruster is needed to spray and offset the interference of the force, resulting in a large consumption of propellant.

[0076] The F1~F4 attitude control thrusters and the F5~F8 attitude control thrusters form a dual-point backup with each other, and can also serve as backup thrusters when any orbit control thruster fails, greatly improving system reliability.

[0077] When the F1 to F8 attitude control thrusters are in normal working condition, the attitude control torque is not affected by the thrust center of mass deviation, and has a strong ability to cope with the rendezvous and docking process, as well as the center of mass changes caused by different working stages of the assembly after the docking is completed.

[0078] like Figure 6 As shown in (b) to (d), the attitude control thruster can be used as an attitude and orbit reuse thruster for orbit control in the direction with the smallest orbit control amount and the lowest requirement for the rendezvous and docking mission. While providing the torque required for attitude control, it also provides the resultant force required for orbit control.

[0079] The present application also includes a thruster layout design device, comprising a memory and a processor, wherein the memory is used to store instructions executable by the processor; and the processor is used to execute the instructions to implement the thruster layout design method described above.

[0080] Figure 7 This is a system block diagram of a thruster layout design device according to an embodiment of the present application. Figure 7 As shown, the thruster layout design device 700 may include an internal communication bus 701, a processor 702, a read-only memory (ROM) 703, a random access memory (RAM) 704, and a communication port 705. When used on a personal computer, the thruster layout design device 700 may also include a hard disk 706. The internal communication bus 701 can enable data communication between the components of the thruster layout design device 700. The processor 702 can make judgments and issue prompts. In some embodiments, the processor 702 can be composed of one or more processors. The communication port 705 can enable data communication between the thruster layout design device 700 and the outside world. In some embodiments, the thruster layout design device 700 can send and receive information and data from a network via the communication port 705. The thruster layout design device 700 may also include various forms of program storage units and data storage units, such as a hard disk 706, a read-only memory (ROM) 703, and a random access memory (RAM) 704, capable of storing various data files used for computer processing and / or communication, as well as possible program instructions executed by the processor 702. The processor executes these instructions to implement the main part of the method. The results of the processor processing are transmitted to the user device via a communication port and displayed on the user interface.

[0081] The above-mentioned operating method may be implemented as a computer program, stored in the hard disk 706 , and loaded into the processor 702 for execution to implement the thruster layout design method of the present application.

[0082] The present application also includes a computer-readable medium storing computer program code, which, when executed by a processor, implements the thruster layout design method described above.

[0083] When the thruster layout design method is implemented as a computer program, it can also be stored in a computer-readable storage medium as an article of manufacture. For example, a computer-readable storage medium may include, but is not limited to, magnetic storage devices (e.g., hard disks, floppy disks, magnetic strips), optical disks (e.g., compact disks (CDs), digital versatile disks (DVDs)), smart cards, and flash memory devices (e.g., electrically erasable programmable read-only memories (EPROMs), cards, sticks, key drives). In addition, the various storage media described herein can represent one or more devices and / or other machine-readable media for storing information. The term "machine-readable medium" may include, but is not limited to, wireless channels and various other media (and / or storage media) that can store, contain, and / or carry code and / or instructions and / or data.

[0084] It should be understood that the embodiments described above are merely illustrative. The embodiments described herein may be implemented in hardware, software, firmware, middleware, microcode, or any combination thereof. For hardware implementation, the processor may be implemented within one or more application specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DSPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors, and / or other electronic units designed to perform the functions described herein, or a combination thereof.

[0085] Some aspects of the present application can be performed entirely by hardware, entirely by software (including firmware, resident software, microcode, etc.), or by a combination of hardware and software. The above hardware or software can be referred to as "data blocks", "modules", "engines", "units", "components" or "systems". The processor can be one or more application-specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DAPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors or combinations thereof. In addition, various aspects of the present application may be expressed as computer products located in one or more computer-readable media, which include computer-readable program code. For example, computer-readable media may include, but are not limited to, magnetic storage devices (e.g., hard disks, floppy disks, tapes...), optical disks (e.g., compact disks CDs, digital versatile disks DVDs...), smart cards, and flash memory devices (e.g., cards, sticks, key drives...).

[0086] A computer-readable medium may include a propagated data signal embodying computer program code, for example, in baseband or as part of a carrier wave. The propagated signal may be in a variety of forms, including electromagnetic, optical, etc., or a suitable combination thereof. A computer-readable medium may be any computer-readable medium other than a computer-readable storage medium that can be connected to an instruction execution system, apparatus, or device to communicate, propagate, or transmit the program for use. The program code on the computer-readable medium may be transmitted via any suitable medium, including radio, cable, fiber optic cable, radio frequency signal, or similar medium, or any combination of the above.

[0087] The basic concepts have been described above. It will be apparent to those skilled in the art that the above disclosures are merely examples and do not limit the present application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and revisions to the present application. Such modifications, improvements, and revisions are suggested in the present application and remain within the spirit and scope of the exemplary embodiments of the present application.

Claims

1. A thruster layout design method for micro-nano satellite rendezvous and docking, characterized in that: include: Constructing a satellite body coordinate system, wherein the origin of the satellite body coordinate system is located at the satellite's center of mass and the three axes of the satellite body coordinate system are parallel and in the same direction as the axes of the orbital coordinate system in the Earth orientation attitude; Determining the positions and thrust directions of the four orbit control thrusters in the satellite body coordinate system according to the first constraint condition; Selecting a first mounting surface and a second mounting surface from the satellite body coordinate system, wherein the first mounting surface and the second mounting surface are symmetrical about the satellite body; determining, based on a second constraint condition, the position and thrust direction of a first set of attitude control thrusters on the first mounting surface, and the position and thrust direction of a second set of attitude control thrusters on the second mounting surface, wherein the second set of attitude control thrusters and the first set of attitude control thrusters are symmetrically mounted in a force couple arrangement; Among them, the first constraint condition is that the combined force of the four orbit control thrusters in pairs passes through the satellite's center of mass, and the second constraint condition is that the thrust generated by the attitude control thrusters does not exceed the satellite's center of mass.

2. The method according to claim 1, wherein The first group of attitude control thrusters and the second group of attitude control thrusters each include four attitude control thrusters.

3. The method according to claim 1, wherein Determining the positions and thrust directions of the four orbit control thrusters in the satellite body coordinate system according to the first constraint condition includes: The planes where the thrust axes of the four orbital control thrusters are located are determined based on the two directions with the largest speed increments required for the rendezvous and docking mission; The position equation and thrust equation of each orbit control thruster are constructed on the plane according to the first constraint condition to obtain the position and thrust direction of the four orbit control thrusters in the satellite body coordinate system.

4. The method according to claim 3, wherein For the rendezvous and docking mission within the same orbital plane, the two directions with the largest required velocity increments are set as the X-axis and Z-axis of the satellite body coordinate system. The position equation and thrust equation of each orbit control thruster are: C9=[X g 0 Z g ],f9=[-sin(α) 0 -cos(α)]×f g C 10 =[X g 0 -Z g ],f 10 =[-sin(α) 0 cos(α)]×f g C 11 =[-X g 0 -Z g ],f 11 =[sin(α) 0 cos(α)]×f g C 12 =[-X g 0 Z g ],f 12 =[sin(α) 0 -cos(α)]×f g Among them, C9~C 12 are the positions of the four orbital control thrusters, f9~f 12 is the thrust of the four orbit control thrusters, f g is the thrust of a single orbital control thruster, and α is the angle between the thrust direction of the +Z-end orbital control thruster and the +Z axis.

5. The method according to claim 2, wherein Determining the position and thrust direction of the first group of attitude control thrusters on the first mounting surface according to the second constraint condition includes: Based on the second constraint, position equations and thrust equations for each attitude control thruster in the first group of attitude control thrusters are constructed on the first mounting surface, such that the thrust component of each attitude control thruster in the k direction is non-zero and can meet the k-direction orbital control requirements for the rendezvous and docking mission, where the k direction is the direction in which the orbital control amount required for the rendezvous and docking mission is minimized.

6. The method according to claim 5, wherein The position equation and thrust equation of each attitude control thruster in the first group of attitude control thrusters are: C1=[X z Y z Z z ],f1=[-sin(γ)cos(β) -sin(γ)sin(β) -cos(γ)]×f z C2=[X z -Y z Z z ],f2=[-sin(γ)cos(β) sin(γ)sin(β) -cos(γ)]×f z C3=[X z -Y z -Z z ],f3=[-sin(γ)cos(β) sin(γ)sin(β)cos(γ)]×f z C4=[X z Y z -Z z ],f4=[-sin(γ)cos(β) -sin(γ)sin(β)cos(γ)]×f z Among them, C1~C4 are the installation positions of the four attitude control thrusters, f1~f4 are the thrusts of the four attitude control thrusters, and f z is the thrust of a single attitude control thruster, β is the angle between the projection of the first group of attitude control thrusters on the XY plane and +X, and γ is the angle between the +Z end attitude control thruster vector and +Z.

7. The method according to claim 5, wherein For the rendezvous and docking mission within the same orbital plane, the k direction is the Y-axis direction of the satellite body coordinate system.

8. The method according to claim 5, wherein The first group of attitude control thrusters and the second group of attitude control thrusters form a dual-point backup with each other.

9. The method according to claim 5, wherein The first group of attitude control thrusters and the second group of attitude control thrusters can serve as backup thrusters when any orbit control thruster fails.

10. A thruster layout design device for micro-nano satellite rendezvous and docking, comprising: A processor and a memory, wherein the memory stores a program or instruction that can be run on the processor, and the program or instruction is executed by the processor to implement the method according to any one of claims 1 to 9.

11. A computer-readable medium storing computer program code, wherein the computer program code implements the method according to any one of claims 1 to 9 when executed by a processor.

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