Posture and orbit coupling trajectory planning method and device suitable for asteroid landing exploration

Abstract

The application provides an attitude and orbit coupling trajectory planning method and device suitable for asteroid landing exploration, and belongs to the asteroid exploration field.The method comprises the following steps: establishing an asteroid landing dynamics model;establishing an asteroid landing orbit nominal trajectory model based on the dynamics model;optimizing the asteroid landing orbit nominal trajectory model to obtain an optimization result of the landing orbit nominal trajectory;calculating an asteroid landing attitude nominal trajectory based on the optimization result of the landing orbit nominal trajectory, so as to obtain a trajectory planning result of the asteroid landing exploration.The application can efficiently solve the feasible attitude nominal trajectory and the orbit nominal trajectory of the asteroid explorer in the process of developing the asteroid landing exploration, and provides effective technical support for the asteroid landing exploration.

Description

Technical Field

[0001] This invention belongs to the field of asteroid exploration, and specifically relates to an attitude-orbit coupling trajectory planning method and device suitable for asteroid landing exploration. Background Technology

[0002] The prevailing theory suggests that asteroids were formed synchronously with the solar system and did not undergo complex crustal evolution, likely preserving primordial materials. Therefore, landing and exploring asteroids is of great significance for studying the solar system. Furthermore, some asteroids contain abundant rare metals such as platinum and iridium, making asteroid exploration economically valuable. In addition, some near-Earth asteroids have orbits that overlap with the Earth-Sun orbit; over long periods of evolution, these asteroids may collide with Earth, and landing and exploring asteroids could also contribute to the development of an active asteroid defense system.

[0003] However, communication between the probe and Earth during asteroid exploration suffers from high latency and high energy consumption, necessitating the asteroid probe to plan its own landing trajectory. Currently, the mainstream methods for asteroid landing trajectory planning include the Gaussian pseudospectral method and the convex optimization method, but both suffer from low computational efficiency and cannot be performed in real-time on onboard computers. To better conduct asteroid landing exploration, it is necessary to establish an efficient asteroid landing trajectory planning method to provide effective guidance for asteroid exploration missions. Summary of the Invention

[0004] The purpose of this invention is to overcome the shortcomings of existing technologies and propose an attitude-orbit coupling trajectory planning method and apparatus suitable for asteroid landing exploration. This invention can efficiently calculate feasible nominal attitude trajectories and nominal orbital trajectories for asteroid probes during asteroid landing exploration, providing effective technical support for asteroid landing exploration.

[0005] A first aspect of this invention proposes an attitude-orbit coupling trajectory planning method suitable for asteroid landing exploration, comprising:

[0006] Establish a dynamic model for asteroid landing;

[0007] Based on the aforementioned dynamic model, a nominal trajectory model for asteroid landing orbit is established;

[0008] The nominal trajectory model of the asteroid landing orbit is optimized to obtain the optimized result of the nominal trajectory of the landing orbit;

[0009] Based on the optimization results of the nominal trajectory of the landing orbit, the nominal trajectory of the asteroid landing attitude is calculated to obtain the trajectory planning results of the asteroid landing exploration.

[0010] In a specific embodiment of the present invention, the asteroid landing dynamics model is constructed based on the asteroid body coordinate system; wherein, the asteroid body coordinate system has the asteroid's center of mass O as the origin, the x-axis as the principal axis of the asteroid's minimum moment of inertia, the y-axis as the principal axis of the asteroid's intermediate moment of inertia, and the z-axis as the principal axis of the asteroid's maximum moment of inertia, and the x-axis, y-axis and z-axis form a right-handed system.

[0011] In a specific embodiment of the present invention, the expression for the asteroid landing dynamics model is as follows:

[0012]

[0013] Where r is the position vector of the asteroid probe, and v is the velocity vector of the asteroid probe. and Let r and v represent the first derivatives, respectively, and T be the thrust vector of the asteroid probe. ω s Let g be the angular velocity of the asteroid's rotation. s (r) is the gravitational field vector of the asteroid at position vector r.

[0014] In a specific embodiment of the present invention, the nominal trajectory model expression for the asteroid landing orbit is as follows:

[0015] r d (t)=c0+c1t+c2t 2 +c3t 3 +c4t 4 +c5t 5 (2)

[0016] Where, r d (t) is the ideal position vector of the probe at time t, i.e. the nominal trajectory of the asteroid landing orbit; t represents time, and c0, c1, c2, c3, c4 and c5 are parameters.

[0017] In one specific embodiment of the present invention, the method further includes:

[0018] Based on the aforementioned asteroid landing dynamics model and the aforementioned asteroid landing orbit nominal trajectory model, the nominal thrust trajectory during the asteroid probe's landing process is established, as expressed below:

[0019]

[0020] Where m0 represents the initial mass of the asteroid probe. and r d The second and first derivatives of (t).

[0021] In one specific embodiment of the present invention, the constraints of the nominal trajectory model of the asteroid landing orbit include:

[0022] The initial state and target state constraints of the detector are expressed as follows:

[0023]

[0024] Among them, r0, r f These are the initial position vector and target position vector of the asteroid probe, respectively; v0, v f These are the initial velocity vector and target velocity vector of the asteroid probe, respectively, n0 and n... f These are the initial thruster direction vector and the target direction vector of the asteroid probe, T0 and T1, respectively. f These represent the initial thrust and the final thrust of the asteroid probe, t0 and t1, respectively. f For the start and end times of landing;

[0025] The collision avoidance constraint is expressed as follows:

[0026] d(r d (t))≥d min (5)

[0027] Where d(r) d (t) represents the position vector r. d (t) is the distance from the asteroid's surface, d min This is the minimum allowable distance for an asteroid probe to reach the asteroid surface.

[0028] The maximum thrust constraint is expressed as follows:

[0029] 0≤||T d (t)||≤T max (6)

[0030] Among them, T max The maximum thrust provided for the asteroid probe.

[0031] In a specific embodiment of the present invention, optimizing the nominal trajectory model of the asteroid landing orbit to obtain the optimized result of the nominal trajectory of the landing orbit includes:

[0032] 1) Determine the optimization index, expressed as follows:

[0033]

[0034] Where β is the weighting coefficient;

[0035] 2) Determine the range of values ​​for the parameter to be optimized, including:

[0036]

[0037] Where G is the gravitational constant, m s The mass of the asteroid;

[0038] 3) By finding T0 and T that satisfy the linear constraint (8) f and t f Under the condition of satisfying the nonlinear constraints (5) and (6), the optimization index (7) is minimized to obtain the optimized nominal trajectory r of the asteroid landing orbit. d (t).

[0039] In one specific embodiment of the present invention, the method further includes:

[0040] Based on the optimized asteroid landing orbit nominal trajectory r d (t), using equation (3) to obtain the nominal thrust trajectory T d (t), then the nominal trajectory of the asteroid landing attitude is:

[0041]

[0042] Where, n d (t) represents the ideal thruster orientation at time t, i.e., the nominal trajectory of the asteroid landing attitude;

[0043] r d (t) and n d (t) represents the trajectory planning result for asteroid landing exploration.

[0044] In one specific embodiment of the present invention, the optimized asteroid landing orbit nominal trajectory is obtained by using any one of the following algorithms: simulated annealing algorithm, genetic algorithm, and compass search algorithm.

[0045] A second aspect of the present invention provides an attitude-orbit coupling trajectory planning device suitable for asteroid landing exploration, comprising:

[0046] The dynamics model building module is used to establish asteroid landing dynamics models;

[0047] The nominal trajectory construction module is used to establish a nominal trajectory model for asteroid landing based on the dynamic model.

[0048] The nominal trajectory optimization module is used to optimize the nominal trajectory model of the asteroid landing trajectory to obtain the optimized result of the nominal trajectory of the landing trajectory.

[0049] The attitude nominal trajectory optimization module is used to calculate the asteroid landing attitude nominal trajectory based on the optimization results of the landing orbit nominal trajectory, so as to obtain the trajectory planning results of the asteroid landing exploration.

[0050] A third aspect of the present invention provides an electronic device comprising:

[0051] At least one processor; and a memory communicatively connected to said at least one processor;

[0052] The memory stores instructions that can be executed by the at least one processor, and the instructions are configured to execute the above-described attitude-orbit coupling trajectory planning method suitable for asteroid landing exploration.

[0053] A fourth aspect of the present invention provides a computer-readable storage medium storing computer instructions for causing the computer to execute the above-described attitude-orbit coupling trajectory planning method applicable to asteroid landing exploration.

[0054] The features and beneficial effects of this invention are as follows:

[0055] During asteroid landing exploration, the asteroid probe needs to autonomously plan a feasible landing trajectory. This invention first proposes a polynomial-form nominal trajectory for asteroid landing exploration. Second, it proposes an efficient combinatorial optimization method to achieve efficient trajectory optimization. Finally, based on the optimized nominal trajectory, this invention calculates an ideal attitude nominal trajectory, enabling efficient solution of attitude-orbit coupled trajectories suitable for asteroid landing exploration. This invention can be used in future deep space exploration missions such as asteroid landing exploration and asteroid impact defense, providing effective technical support for asteroid landing exploration and possessing high application value. Attached Figure Description

[0056] Figure 1 This is an overall flowchart of an attitude-orbit coupling trajectory planning method applicable to asteroid landing exploration, as shown in a specific embodiment of the present invention.

[0057] Figure 2 This is a schematic diagram of the asteroid body coordinate system in a specific embodiment of the present invention.

[0058] Figure 3 This is a schematic diagram of the landing trajectory of an asteroid probe in a specific embodiment of the present invention. Detailed Implementation

[0059] This invention proposes an attitude-orbit coupling trajectory planning method and device suitable for asteroid landing exploration. The invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0060] This invention proposes an attitude-orbit coupling trajectory planning method suitable for asteroid landing exploration, comprising:

[0061] Establish a dynamic model for asteroid landing;

[0062] Based on the aforementioned dynamic model, a nominal trajectory model for asteroid landing orbit is established;

[0063] The nominal trajectory model of the asteroid landing orbit is optimized to obtain the optimized result of the nominal trajectory of the landing orbit;

[0064] Based on the optimization results of the nominal trajectory of the landing orbit, the nominal trajectory of the asteroid landing attitude is calculated to obtain the trajectory planning results of the asteroid landing exploration.

[0065] In a specific embodiment of the present invention, the overall process of the attitude-orbit coupling trajectory planning method suitable for asteroid landing exploration is as follows: Figure 1 As shown, it includes the following steps:

[0066] 1) Establish a coordinate system.

[0067] The coordinate system used in this embodiment of the invention is , and the following descriptions are all based on this coordinate system. Figure 2 This is a schematic diagram of the asteroid body coordinate system in a specific embodiment of the present invention. Figure 2 This shows the asteroid Bennu and its corresponding asteroid body coordinate system, such as... Figure 2 As shown, the asteroid's body coordinate system has the asteroid's center of mass O as the origin of the coordinate system. The x-axis is the principal axis of the asteroid's minimum moment of inertia, the y-axis is the principal axis of the asteroid's intermediate moment of inertia, and the z-axis is the principal axis of the asteroid's maximum moment of inertia. The three axes form a right-handed system.

[0068] 2) Based on the coordinate system established in step 1), establish an asteroid landing dynamics model.

[0069] In this embodiment of the invention, a landing dynamics model for a probe in an asteroid environment is established, and the expression is as follows:

[0070]

[0071] Where r is the position vector of the asteroid probe, and v is the velocity vector of the asteroid probe. and Let r and v represent the first derivatives, respectively, and T be the thrust vector of the detector. ω s Let g be the angular velocity of the asteroid's rotation. s (r) represents the gravitational field vector of the asteroid at position vector r. The asteroid's gravitational field vector can be calculated using the asteroid polyhedral model combined with the polyhedral method. (The polyhedral method is a commonly used method for calculating gravitational fields in the field of asteroid exploration).

[0072] 3) Based on the asteroid landing dynamics model established in step 2), establish the nominal trajectory model for the asteroid landing orbit. The specific steps are as follows:

[0073] 3-1) Establish a polynomial-form nominal trajectory model for asteroid landing orbits.

[0074] In this embodiment of the invention, a fifth-order polynomial is used to establish the nominal trajectory model of the landing orbit. Figure 3 This is a schematic diagram of the landing trajectory of an asteroid probe in a specific embodiment of the present invention. Figure 3 This paper demonstrates a feasible fifth-order polynomial nominal orbital trajectory from the initial position of an asteroid probe to the target position, according to a specific embodiment of the present invention. The expression for the asteroid landing orbit nominal trajectory model is as follows:

[0075] r d (t)=c0+c1t+c2t 2 +c3t 3 +c4t 4 +c5t 5 (2)

[0076] Where, r d (t) is the ideal position vector of the probe at time t, i.e. the nominal trajectory of the asteroid landing orbit; t represents time, and c0, c1, c2, c3, c4 and c5 are parameters to be determined.

[0077] 3-2) Establish the nominal thrust trajectory during the landing process of the asteroid probe.

[0078] In this embodiment, by combining equations (1) and (2), the nominal thrust trajectory T of the asteroid probe during its landing process under the nominal trajectory can be obtained. d (t), the expression is as follows:

[0079]

[0080] Where m0 represents the initial mass of the asteroid probe. and r d The second and first derivatives of (t).

[0081] 3-3) Determine the constraints of the nominal trajectory model.

[0082] In this embodiment, the nominal track r d (t) is constrained by the initial state of the detector and the target state, and its expression is as follows:

[0083]

[0084] Among them, r0, r fThese are the initial position vector and the target position vector of the asteroid probe, respectively. v0, v f Let n0 and n' be the initial velocity vector of the detector and the velocity vector of the target, respectively. f These are the initial thruster direction vector and the target direction vector of the probe, respectively. T0, T f These represent the initial thrust and the final thrust of the probe, t0 and t1, respectively. f Let t0 be the start and end times of landing. Without loss of generality, we can assume t0 = 0. Among the variables mentioned above, r0 and r... f v0, v f n0, n f The landing and exploration mission was determined in advance, T0, T f and t f Let r be an optimizable variable. Combining equations (2) and (3), we can obtain r. d Only determined by T0, T f and t f Three variables to be optimized.

[0085] To avoid collision with the asteroid's surface, the nominal orbital trajectory r d (t) also needs to satisfy anti-collision constraints, as shown in the following expression:

[0086] d(r d (t))≥d min (5)

[0087] Where d(r) d (t) represents the position vector r. d The distance (t) to the asteroid surface can be calculated using an asteroid polyhedral model combined with spherical harmonic series methods. min This is the minimum allowable distance for an asteroid probe to reach the surface of an asteroid.

[0088] During the landing process, the asteroid probe's thrust must also meet the maximum thrust constraint of the thruster, as expressed below:

[0089] 0≤||T d (t)||≤T max (6)

[0090] Among them, T max The maximum thrust provided for the asteroid probe.

[0091] 4) Optimize the nominal trajectory model of the asteroid landing orbit established in step 3) to obtain the optimized result of the nominal trajectory of the landing orbit.

[0092] In this embodiment of the invention, a common combinatorial optimization method is used to optimize the nominal trajectory of the asteroid landing orbit. The specific steps are as follows:

[0093] 4-1) Determine the optimization indicators.

[0094] The optimization index used in this embodiment of the invention takes into account both fuel consumption and landing time during the landing and exploration process, and its expression is as follows:

[0095]

[0096] Wherein, β is a weighting coefficient, which can be adjusted according to the specific landing mission. Generally, the value of β is in the range of [0, 0.01). In one embodiment of the present invention, β is 0.001.

[0097] 4-2) Determine the range of values ​​for the parameter to be optimized.

[0098] In this embodiment, combining equations (2) and (3), it can be seen that the parameters to be optimized in the nominal trajectory of the asteroid landing orbit are T0 and T1. f and t f The range of values ​​for these three parameters is as follows:

[0099]

[0100] Where G is the gravitational constant, m s The mass of the asteroid.

[0101] 4-3) Calculate the optimized result of the nominal trajectory of the landing track.

[0102] The optimization problem to be solved in this embodiment of the invention can be described as: finding T0 and T that satisfy the linear constraint equation (8). f and t f Under the condition of satisfying the nonlinear constraints (5) and (6), the optimization index (7) is minimized. This optimization problem belongs to the combinatorial optimization problem, and commonly used high-efficiency combinatorial optimization methods can be used, such as simulated annealing algorithm, genetic algorithm, compass search algorithm, etc. In this embodiment of the invention, the compass search algorithm is used for optimization calculation to obtain the optimized asteroid landing orbit nominal trajectory r. d (t).

[0103] 5) Based on the results of step 4), calculate the nominal trajectory of the asteroid's landing attitude.

[0104] In this embodiment of the invention, the thruster orientation of the lander during landing, i.e., the attitude nominal trajectory, is determined based on the nominal orbital trajectory. Through steps 2) to 4), the optimized asteroid landing orbit nominal trajectory r can be obtained. d (t), and then combined with equation (3), we obtain the nominal thrust trajectory T. d (t), then the nominal trajectory of the asteroid landing attitude is:

[0105]

[0106] Where, n d (t) represents the ideal thruster orientation at time t, i.e., the nominal trajectory of the asteroid landing attitude.

[0107] r d (t) and n d (t) represents the trajectory planning result for asteroid landing exploration.

[0108] Steps 4) and 5) yield the optimized nominal trajectory and attitude trajectory for asteroid landing, respectively. During the asteroid landing mission, by further integrating the thrusters and attitude controllers to track the nominal trajectory and attitude trajectory, asteroid landing and exploration can be achieved.

[0109] To achieve the above embodiments, a second aspect of the present invention proposes an attitude-orbit coupling trajectory planning device suitable for asteroid landing exploration, comprising:

[0110] The dynamics model building module is used to establish asteroid landing dynamics models;

[0111] The nominal trajectory construction module is used to establish a nominal trajectory model for asteroid landing based on the dynamic model.

[0112] The nominal trajectory optimization module is used to optimize the nominal trajectory model of the asteroid landing trajectory to obtain the optimized result of the nominal trajectory of the landing trajectory.

[0113] The attitude nominal trajectory optimization module is used to calculate the asteroid landing attitude nominal trajectory based on the optimization results of the landing orbit nominal trajectory, so as to obtain the trajectory planning results of the asteroid landing exploration.

[0114] It should be noted that the foregoing explanation of an embodiment of an attitude-orbit coupling trajectory planning method suitable for asteroid landing exploration also applies to an attitude-orbit coupling trajectory planning device suitable for asteroid landing exploration in this embodiment, and will not be repeated here. According to an embodiment of the present invention, an attitude-orbit coupling trajectory planning device suitable for asteroid landing exploration establishes an asteroid landing dynamics model; based on the dynamics model, establishes an asteroid landing orbit nominal trajectory model; optimizes the asteroid landing orbit nominal trajectory model to obtain the optimized landing orbit nominal trajectory; and based on the optimized landing orbit nominal trajectory, calculates the asteroid landing attitude nominal trajectory to obtain the trajectory planning result for asteroid landing exploration. This enables efficient calculation of feasible attitude and orbit nominal trajectories for asteroid probes during asteroid landing exploration, providing effective technical support for asteroid landing exploration.

[0115] To implement the above embodiments, a third aspect of the present invention provides an electronic device, comprising:

[0116] At least one processor; and a memory communicatively connected to said at least one processor;

[0117] The memory stores instructions that can be executed by the at least one processor, and the instructions are configured to execute the above-described attitude-orbit coupling trajectory planning method suitable for asteroid landing exploration.

[0118] To implement the above embodiments, a fourth aspect of the present invention provides a computer-readable storage medium storing computer instructions for causing the computer to execute the above-described attitude-orbit coupling trajectory planning method applicable to asteroid landing exploration.

[0119] It should be noted that the computer-readable medium described in this disclosure can be a computer-readable signal medium or a computer-readable storage medium, or any combination thereof. A computer-readable storage medium can be, for example,—but not limited to—an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples of a computer-readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination thereof. In this disclosure, a computer-readable storage medium can be any tangible medium containing or storing a program that can be used by or in connection with an instruction execution system, apparatus, or device. In this disclosure, a computer-readable signal medium can include a data signal propagated in baseband or as part of a carrier wave, carrying computer-readable program code. Such propagated data signals can take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. A computer-readable signal medium can be any computer-readable medium other than a computer-readable storage medium, which can send, propagate, or transmit a program for use by or in connection with an instruction execution system, apparatus, or device. The program code contained on the computer-readable medium can be transmitted using any suitable medium, including but not limited to: wires, optical fibers, RF (radio frequency), etc., or any suitable combination thereof.

[0120] The aforementioned computer-readable medium may be included in the aforementioned electronic device; or it may exist independently and not assembled into the electronic device. The aforementioned computer-readable medium carries one or more programs, which, when executed by the electronic device, cause the electronic device to perform an attitude-orbit coupling trajectory planning method applicable to asteroid landing exploration according to the above embodiments.

[0121] Computer program code for performing the operations of this disclosure can be written in one or more programming languages ​​or a combination thereof, including object-oriented programming languages ​​such as Java, Smalltalk, and C++, and conventional procedural programming languages ​​such as the "C" language or similar programming languages. The program code can be executed entirely on the user's computer, partially on the user's computer, as a standalone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In cases involving remote computers, the remote computer can be connected to the user's computer via any type of network—including a local area network (LAN) or a wide area network (WAN)—or can be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0122] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., refer to specific features, structures, materials, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of this application. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0123] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this application, "multiple" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0124] Any process or method described in the flowchart or otherwise herein can be understood as representing a module, segment, or portion of code comprising one or more executable instructions for implementing a particular logical function or process, and the scope of the preferred embodiments of this application includes additional implementations in which functions may be performed not in the order shown or discussed, including substantially simultaneously or in reverse order depending on the function involved, as will be understood by those skilled in the art to which embodiments of this application pertain.

[0125] The logic and / or steps represented in the flowchart or otherwise described herein, for example, can be considered as a sequenced list of executable instructions for implementing logical functions, and can be embodied in any computer-readable medium for use by, or in conjunction with, an instruction execution system, apparatus, or device (such as a computer-based system, a processor-included system, or other system that can fetch and execute instructions from, an instruction execution system, apparatus, or device). For the purposes of this specification, "computer-readable medium" can be any means that can contain, store, communicate, propagate, or transmit programs for use by, or in conjunction with, an instruction execution system, apparatus, or device. More specific examples (a non-exhaustive list) of computer-readable media include: an electrical connection having one or more wires (electronic device), a portable computer disk drive (magnetic device), random access memory (RAM), read-only memory (ROM), erasable and editable read-only memory (EPROM or flash memory), fiber optic devices, and portable optical disc read-only memory (CDROM). Furthermore, computer-readable media can even be paper or other suitable media on which programs can be printed, because programs can be obtained electronically, for example, by optically scanning the paper or other media, followed by editing, interpreting, or otherwise processing as necessary, and then stored in computer memory.

[0126] It should be understood that various parts of this application can be implemented using hardware, software, firmware, or a combination thereof. In the above embodiments, multiple steps or methods can be implemented using software or firmware stored in memory and executed by a suitable instruction execution system. For example, if implemented in hardware, as in another embodiment, it can be implemented using any one or a combination of the following techniques known in the art: discrete logic circuits having logic gates for implementing logical functions on data signals, application-specific integrated circuits (ASICs) having suitable combinational logic gates, programmable gate arrays (PGAs), field-programmable gate arrays (FPGAs), etc.

[0127] Those skilled in the art will understand that all or part of the steps of the methods in the above embodiments can be implemented by a program instructing related hardware. The program can be stored in a computer-readable storage medium, and when executed, the program includes one or a combination of the steps of the method embodiments.

[0128] Furthermore, the functional units in the various embodiments of this application can be integrated into a processing module, or each unit can exist physically separately, or two or more units can be integrated into a module. The integrated module can be implemented in hardware or as a software functional module. If the integrated module is implemented as a software functional module and sold or used as an independent product, it can also be stored in a computer-readable storage medium.

[0129] The storage medium mentioned above can be a read-only memory, a disk, or an optical disk, etc. Although embodiments of this application have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting this application. Those skilled in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of this application.

Claims

1. A method for attitude-orbit coupling trajectory planning suitable for asteroid landing exploration, characterized in that, include: Establish a dynamic model for asteroid landing; Based on the aforementioned dynamic model, a nominal trajectory model for asteroid landing orbit is established; The nominal trajectory model of the asteroid landing orbit is optimized to obtain the optimized result of the nominal trajectory of the landing orbit; Based on the optimization results of the nominal trajectory of the landing orbit, the nominal trajectory of the asteroid landing attitude is calculated to obtain the trajectory planning results of the asteroid landing exploration. The asteroid landing dynamics model is constructed based on the asteroid body coordinate system. The asteroid body coordinate system has the asteroid's center of mass O as the origin, the x-axis as the principal axis of the asteroid's minimum moment of inertia, the y-axis as the principal axis of the asteroid's intermediate moment of inertia, and the z-axis as the principal axis of the asteroid's maximum moment of inertia. The x-axis, y-axis, and z-axis form a right-handed system. The expression for the asteroid landing dynamics model is as follows: (1) in, This represents the position vector of the asteroid probe. Let be the velocity vector of the asteroid probe. and They represent and The first derivative, This represents the thrust vector of the asteroid probe. Let be the asteroid's angular velocity of rotation. Position vector The gravitational field vector of the asteroid at that location; The nominal trajectory model expression for the asteroid landing orbit is as follows: (2) in, Let t be the ideal position vector of the probe at time t, i.e., the nominal trajectory of the asteroid landing orbit; t represents time. , , , , and For parameters; Based on the aforementioned asteroid landing dynamics model and the aforementioned asteroid landing orbit nominal trajectory model, the nominal thrust trajectory during the asteroid probe's landing process is established, as expressed below: (3) in, This indicates the initial mass of the asteroid probe. and They are respectively The second and first derivatives; The constraints of the nominal trajectory model for the asteroid landing orbit include: The initial state and target state constraints of the detector are expressed as follows: (4) in, , These are the initial position vector and the target position vector of the asteroid probe, respectively; , These are the initial velocity vector and the target velocity vector of the asteroid probe, respectively. , These are the initial thruster direction vector and the target direction vector of the asteroid probe, respectively. , These represent the initial thrust and the final thrust of the asteroid probe, respectively. , For the start and end times of landing; The collision avoidance constraint is expressed as follows: (5) in, Position vector Distance to the surface of the asteroid, This is the minimum allowable distance for an asteroid probe to reach the asteroid surface. The maximum thrust constraint is expressed as follows: (6) in, The maximum thrust provided for the asteroid probe.

2. The method according to claim 1, characterized in that, The optimization of the nominal trajectory model of the asteroid landing orbit to obtain the optimized nominal trajectory of the landing orbit includes: 1) Determine the optimization metric, expressed as follows: (7) in, These are the weighting coefficients; 2) Determine the range of values ​​for the parameter to be optimized, including: (8) in, The gravitational constant is... The mass of the asteroid; 3) By finding the linear constraint (8) , and Under the condition of satisfying the nonlinear constraints (5) and (6), the optimization index (7) is minimized to obtain the optimized nominal trajectory of the asteroid landing orbit. .

3. The method according to claim 2, characterized in that, The method further includes: Based on the optimized asteroid landing orbit nominal trajectory The nominal thrust trajectory is obtained using equation (3). The nominal trajectory for asteroid landing attitude is: (8) in, This indicates the ideal thruster orientation at time t, i.e., the nominal trajectory of the asteroid landing attitude; and This refers to the trajectory planning results for asteroid landing exploration.

4. The method according to claim 2, characterized in that, The optimized asteroid landing orbit nominal trajectory is obtained using any one of the following algorithms: simulated annealing algorithm, genetic algorithm, or compass search algorithm.

5. An attitude-orbit coupling trajectory planning device suitable for asteroid landing exploration, characterized in that, include: The dynamics model building module is used to establish asteroid landing dynamics models; The nominal trajectory construction module is used to establish a nominal trajectory model for asteroid landing based on the dynamic model. The nominal trajectory optimization module is used to optimize the nominal trajectory model of the asteroid landing trajectory to obtain the optimized result of the nominal trajectory of the landing trajectory. The attitude nominal trajectory optimization module is used to calculate the asteroid landing attitude nominal trajectory based on the optimization results of the landing orbit nominal trajectory, so as to obtain the trajectory planning results of the asteroid landing exploration; The asteroid landing dynamics model is constructed based on the asteroid body coordinate system. The asteroid body coordinate system has the asteroid's center of mass O as the origin, the x-axis as the principal axis of the asteroid's minimum moment of inertia, the y-axis as the principal axis of the asteroid's intermediate moment of inertia, and the z-axis as the principal axis of the asteroid's maximum moment of inertia. The x-axis, y-axis, and z-axis form a right-handed system. The expression for the asteroid landing dynamics model is as follows: (1) in, This represents the position vector of the asteroid probe. Let be the velocity vector of the asteroid probe. and They represent and The first derivative, This represents the thrust vector of the asteroid probe. Let be the asteroid's angular velocity of rotation. Position vector The gravitational field vector of the asteroid at that location; The nominal trajectory model expression for the asteroid landing orbit is as follows: (2) in, Let t be the ideal position vector of the probe at time t, i.e., the nominal trajectory of the asteroid landing orbit; t represents time. , , , , and For parameters; Based on the aforementioned asteroid landing dynamics model and the aforementioned asteroid landing orbit nominal trajectory model, the nominal thrust trajectory during the asteroid probe's landing process is established, as expressed below: (3) in, This indicates the initial mass of the asteroid probe. and They are respectively The second and first derivatives; The constraints of the nominal trajectory model for the asteroid landing orbit include: The initial state and target state constraints of the detector are expressed as follows: (4) in, , These are the initial position vector and the target position vector of the asteroid probe, respectively; , These are the initial velocity vector and the target velocity vector of the asteroid probe, respectively. , These are the initial thruster direction vector and the target direction vector of the asteroid probe, respectively. , These represent the initial thrust and the final thrust of the asteroid probe, respectively. , For the start and end times of landing; The collision avoidance constraint is expressed as follows: (5) in, Position vector Distance to the surface of the asteroid, This is the minimum allowable distance for an asteroid probe to reach the asteroid surface. The maximum thrust constraint is expressed as follows: (6) in, The maximum thrust provided for the asteroid probe.

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