Spacecraft rendezvous and docking synchronization control method and system

Abstract

This application discloses a spacecraft rendezvous and docking synchronization control method and system, relating to the field of spacecraft control. The method includes: considering the physical characteristics of the spacecraft and space environment factors, constructing a pose model based on dual quaternions and formulating constraints; constructing a potential function reflecting the spacecraft's pose state based on the pose model and constraints; constructing a sliding mode control strategy based on dual quaternions based on the potential function to obtain the sliding mode surface; determining the control conditions for the synchronous convergence of each state component based on the theory of time synchronization stability; constructing a controller based on the sliding mode surface and control conditions; and generating control signals based on the real-time acquired spacecraft state to achieve rendezvous and docking control. This application can solve the problem that nonlinear terms in existing spacecraft motion and dynamics models lead to mutual coupling between various channels, and that linearized models cannot accurately represent the spacecraft's attitude and pose motion.

Description

Technical Field

[0001] This application relates to the field of spacecraft control, and in particular to a method and system for synchronous control of spacecraft rendezvous and docking. Background Technology

[0002] Spacecraft attitude control is crucial for various space missions, including astronomical observation, remote sensing, long-distance communication, and space science experiments. The high value, multiple constraints, and high precision requirements of spacecraft necessitate consideration of numerous factors in spacecraft control. Early research included methods such as the Riccati equation and variational integral methods using Lie groups; however, these methods, which seek optimization strategies for linear systems within nonlinearities, often exhibit poor performance. Artificial intelligence offers new methods for spacecraft control, but its effectiveness is often limited by onboard computing resources. Therefore, developing an efficient and accurate spacecraft control method to achieve low-energy, high-control-requirement attitude control in an on-orbit environment has significant theoretical value and application prospects.

[0003] For spacecraft rendezvous and docking, due to structural limitations, visual sensors typically have only a limited field of view (i.e., a limited line of sight angle). Therefore, it is necessary to appropriately control the attitude motion of the tracker to meet this field of view requirement (called field of view constraint). This control requires a high degree of precision and real-time performance; any tiny deviation can lead to rendezvous and docking failure.

[0004] However, most existing control methods only address a portion of a complex process or provide solutions to stability problems under specific conditions, failing to offer a general solution for spacecraft control under various constraints. For nonlinear systems, they are often linearized near the equilibrium point, and then controllers are designed using optimal control methods for linear systems. However, nonlinear terms in spacecraft motion and dynamics models lead to coupling between different channels, and linearized models cannot accurately represent the spacecraft's attitude and pose motion. Summary of the Invention

[0005] To address the aforementioned issues, this application provides a spacecraft rendezvous and docking synchronization control method and system.

[0006] To achieve the above objectives, this application provides the following solution: Firstly, this application provides a spacecraft rendezvous and docking synchronization control method, including: Considering the physical characteristics of the spacecraft and the space environment factors, a pose model for describing the six-degree-of-freedom motion of the spacecraft is constructed based on dual quaternions, and constraints are formulated according to the requirements of the space mission. Based on the pose model and the constraints, a potential function is constructed to reflect the pose state of the spacecraft. Based on the potential function, a sliding mode control strategy based on dual quaternions is constructed to obtain the sliding surface; Based on the theory of time-synchronous stability, the control conditions for the synchronous convergence of each state component are determined. A controller is constructed based on the sliding surface and the control conditions; The controller generates control signals based on the real-time acquired spacecraft status to achieve rendezvous and docking control of the spacecraft.

[0007] Secondly, this application provides a spacecraft rendezvous and docking synchronization control system, comprising: The pose model and constraint design module is used to consider the physical characteristics of the spacecraft and space environment factors, construct a pose model based on dual quaternions to describe the six degrees of freedom motion of the spacecraft, and formulate constraints according to the requirements of the space mission. The potential function design module based on dual quaternions is used to construct a potential function that reflects the attitude state of the spacecraft based on the pose model and the constraints. The artificial potential field function sliding surface design module is used to construct a sliding mode control strategy based on dual quaternions based on the potential function to obtain the sliding surface; The time synchronization stability control condition design module is used to determine the control conditions for the synchronous convergence of each state component based on the theory of time synchronization stability. The controller design module is used to construct a controller based on the sliding surface and the control conditions; the controller generates control signals based on the real-time acquired spacecraft status to realize the rendezvous and docking control of the spacecraft.

[0008] According to the specific embodiments provided in this application, this application has the following technical effects: This application provides a spacecraft rendezvous and docking synchronization control method and system. By considering the physical characteristics of the spacecraft and space environment factors, a pose model for describing the six degrees of freedom motion of the spacecraft is constructed based on dual quaternions. In combination with constraints, a potential function for reflecting the pose state of the spacecraft is constructed. Based on the control of the original artificial potential field function, the anti-interference capability is improved and the energy consumption is reduced, thereby realizing the basic control of the six degrees of freedom of the spacecraft pose under line-of-sight constraints. By constructing a sliding mode control strategy based on dual quaternions using potential functions and combining it with control conditions derived from time-synchronous stability theory, a controller is built. This approach integrates the concept of time-synchronous stability with existing six-degree-of-freedom control methods, achieving consistent and orderly convergence of all state components. Furthermore, by precisely controlling the convergence speed of position and attitude, the controller ensures temporal synchronization, avoiding system instability caused by excessively fast or slow convergence of a particular state component. This not only improves control accuracy and reduces error accumulation but also effectively prevents excessive energy consumption due to over-adjustment. It achieves precise synchronous control for spacecraft rendezvous and docking, thereby addressing the problem in existing spacecraft motion and dynamics models where nonlinear terms lead to coupling between channels, and linearized models cannot accurately represent the spacecraft's attitude and pose. Attached Figure Description

[0009] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0010] Figure 1 A schematic flowchart illustrating a spacecraft rendezvous and docking synchronization control method provided in an embodiment of this application; Figure 2 A schematic diagram of a six-degree-of-freedom spacecraft control process provided in an embodiment of this application; Figure 3 A schematic diagram illustrating the time synchronization stabilization principle provided in an embodiment of this application; Figure 4 A schematic diagram of spacecraft control constraints based on dual quaternions provided for an embodiment of this application; Figure 5 A schematic diagram of sliding mode control based on an artificial potential field function provided in an embodiment of this application; Figure 6 A constrained controlled trajectory diagram of a spacecraft provided in an embodiment of this application; Figure 7 A spacecraft sight cone angle constraint diagram provided for an embodiment of this application; Figure 8 A time synchronization stabilization control effect diagram provided in an embodiment of this application; Figure 9 A schematic diagram comparing the energy consumption of a spacecraft rendezvous and docking synchronization control method provided in one embodiment of this application with that of traditional sliding mode control; Figure 10 A schematic diagram comparing the energy consumption of a spacecraft rendezvous and docking synchronization control method and a PD controller, as provided in an embodiment of this application; Figure 11 A functional module diagram of a spacecraft rendezvous and docking synchronization control system provided in one embodiment of this application; Figure 12 This is a schematic diagram of the structure of a computer device provided in an embodiment of this application. Detailed Implementation

[0011] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0012] To make the above-mentioned objectives, features and advantages of this application more apparent and understandable, the application will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0013] In an exemplary embodiment, this application provides a spacecraft rendezvous and docking synchronization control method, which, in specific applications, can mainly include two parts: the first part is designing sliding mode control based on an artificial potential field function, achieving basic six-degree-of-freedom control of the spacecraft's attitude under line-of-sight constraints by combining it with dual quaternions; the second part is considering time-synchronized stable control to achieve synchronous and orderly convergence of each attitude component of the spacecraft. Based on this, as... Figure 1 As shown, the spacecraft rendezvous and docking synchronization control method provided in this application includes: Step 100: Considering the physical characteristics of the spacecraft and the space environment factors, construct a pose model based on dual quaternions to describe the six-degree-of-freedom motion of the spacecraft, and formulate constraints according to the requirements of the space mission.

[0014] Step 101: Based on the pose model and constraints, construct a potential function to reflect the pose state of the spacecraft.

[0015] Step 102: Based on the potential function, construct a sliding mode control strategy based on dual quaternions to obtain the sliding surface.

[0016] Step 103: Based on the theory of time-synchronous stability, determine the control conditions for the synchronous convergence of each state component.

[0017] Step 104: Construct a controller based on the sliding surface and control conditions.

[0018] Step 105: The controller generates control signals based on the real-time acquired spacecraft status to achieve rendezvous and docking control of the spacecraft.

[0019] Based on the description of steps 100-102 above, this application provides a sliding mode control method for an artificial potential field function under line-of-sight constraints. To achieve the spacecraft's control objective, a potential function is designed that can spontaneously slide towards the target position regardless of the system's current state. To achieve this objective, a constraint condition is designed where the potential function's value is 0 only when the system is at the target point, and positive at other positions. Therefore, the potential function needs to always remain negative definite. Based on this, during the design of the potential function, the constraint condition is transformed into a sliding surface for sliding mode control. After the system state is drawn to the sliding surface, the condition that the derivative of the potential function remains negative definite is always maintained, thus achieving the control requirements. Through the principle of sliding mode control and utilizing the special properties of the sliding surface, the system state is precisely controlled, enabling the spacecraft to continuously adjust its motion state according to the changing trend of the potential function during its sliding towards the target position. This ensures that the direction of motion always points towards the target position, and the speed is reasonably controlled, avoiding excessive acceleration or deceleration, thereby achieving stable and efficient control.

[0020] Furthermore, by implementing steps 103-105 above, this application proposes a time-synchronized stable spacecraft control method. Time synchronization stability is a control method that allows the components of the state space to converge stably and synchronously to the target state. Time synchronization stability can be satisfied by the following lemmas: First, finite-time stability, which guarantees the system convergence within a finite time. Second, ratio persistence, which restricts the orderly and continuous convergence of each state component to the specified position. These are sufficient conditions for time synchronization stability and finite-time stability. Finite-time stability ensures that the spacecraft can respond quickly to control commands and accurately converge to the target state within a finite time window. Ratio persistence imposes strict constraints on the convergence process of each state component, ensuring that they converge orderly and consistently to the specified position.

[0021] In an exemplary embodiment of this application, to ensure the spacecraft operates safely and effectively, the pose model construction process in step 100 comprehensively considers the spacecraft's physical characteristics, including mass distribution and inertial tensor. It also comprehensively considers space environment factors, including the influence of gravitational field and atmospheric drag on the spacecraft's pose. Furthermore, strictly defined constraints include orbital altitude ranges and attitude angle limitations.

[0022] As another implementation method in this embodiment, dual quaternions can simultaneously represent rotation and translation information, making them suitable for describing the six degrees of freedom motion of a spacecraft. During the construction of the pose model, various motion characteristics of the spacecraft in space can be fully considered, such as translation, rotation, acceleration, and deceleration, as well as various external disturbances, such as field-of-view limitations and changes in the target spacecraft's position. Simultaneously, strict constraints are formulated based on mission requirements and the spacecraft's physical characteristics. These constraints cover aspects such as the spacecraft's position range, attitude angle limits, and velocity boundaries, ensuring that the spacecraft's motion occurs within a reasonable, safe, and effective range.

[0023] In an exemplary embodiment of this application, to achieve a unified representation and convergence guidance for six-degree-of-freedom pose, in this embodiment, a potential function that reflects the six-degree-of-freedom pose state of the spacecraft is designed based on the mathematical properties of dual quaternions. This potential function takes the dual quaternion deviation between the spacecraft's real-time pose and the target pose as its core input, takes a value of zero at the target pose, and a positive value at any other pose. Furthermore, its gradient direction can accurately guide the spacecraft to converge toward the target pose. Based on this, the implementation process of step 101 above can be described as follows: Step 11: Based on the pose model and constraints, determine the dual quaternion deviation between the spacecraft's real-time pose and the target pose.

[0024] Step 12: Using the dual quaternion deviation between the spacecraft's real-time pose and the target pose as input, construct a potential function to reflect the spacecraft's pose state. The potential function takes a value of zero at the target pose and a value of positive at other poses. Furthermore, the derivative of the potential function guides the spacecraft's motion toward the target pose. The designed potential function is expressed as: .

[0025] In the formula, Let be the potential function. The dual quaternion deviation between the real-time attitude of the spacecraft and the attitude of the target is given. Dual quaternion deviation between the spacecraft's real-time pose and the target pose The conjugate quaternion, This represents the quaternion inner product. For angular velocity deviation, For positional deviation, For linear velocity deviation, This is the matrix transpose. It is an adaptive weighting factor, and .

[0026] By assigning weights, we can achieve coordinated optimization of attitude, position, and velocity states.

[0027] In an exemplary embodiment of this application, the sliding mode control based on an artificial potential field function can be proven to achieve finite-time stability. When designing the sliding surface, while satisfying the sliding mode control requirements, the spacecraft's rotational inertia is combined with the designed sliding surface to achieve ratio consistency. This satisfies the characteristic of time-synchronous stability, allowing each state component to converge synchronously and stably. Based on this, the implementation process of step 102 provided above can include: Step 21: Based on the potential function, construct a sliding mode control strategy based on dual quaternions to obtain the initial sliding surface. The initial sliding surface is represented as: .

[0028] In the formula, For synovial membrane variables, This is the dual angular velocity. The sliding mode control parameter is a coefficient vector. Auxiliary state variables constructed for the sliding surface. , For vectorization operators, Let be the conjugate form of the relative dual quaternions between the target and the controlled spacecraft. Represents the Kronecker product. It is the gradient operator in the dual quaternion space. This represents the operational law for exchanging the real and imaginary parts of an even number. This represents the Hadamard product. This formula can generate appropriate control signals based on the deviation between the spacecraft's real-time state and the target state.

[0029] Step 22: Optimize the initial sliding surface using the spacecraft's inertial matrix to obtain the final sliding surface. In practical applications, during sliding mode control, this step can utilize the operational properties of dual quaternions, combined with the spacecraft's inertial matrix, to optimize the sliding surface, resulting in the following: .

[0030] In the formula, To account for the sliding membrane variable of the inertia matrix, For the inertial matrix of the spacecraft, This is the dual angular velocity.

[0031] By adjusting the sliding mode control parameters (The larger the parameter, the faster the convergence speed.) The appropriate convergence speed can be selected according to different operating conditions to enhance the anti-interference ability of the system and ensure the stable operation of the spacecraft in complex environments.

[0032] In this embodiment, the design of the sliding mode surface fully considers the dynamic characteristics and control requirements of the spacecraft, enabling it to accurately guide the system state towards the target state. By continuously adjusting the parameters of the sliding mode control, the control effect can be optimized, ensuring that the spacecraft maintains stable gliding under various complex conditions. For example, when facing external disturbances, the sliding mode control can react quickly, adjusting the spacecraft's motion to prevent it from being affected by the disturbances or to quickly return to a normal state. Simultaneously, the potential function is closely integrated with the sliding mode surface, using changes in the potential function to drive the sliding mode control action, achieving coordinated operation between the two. This allows the spacecraft to accurately track the target trajectory in complex space environments and complete various mission requirements.

[0033] In an exemplary embodiment of this application, to ensure that the spacecraft's position, attitude, and other state components converge collaboratively in a predetermined manner, the implementation process of step 104 can be described as follows: Constructing a synthetic Lyapunov function, based on the theory of time-synchronous stability, and through derivation and simulation verification, determining the control conditions for the synchronous convergence of each state component, so as to ensure that the control system remains stable throughout the entire control process and avoids oscillations or divergences. The control conditions may include setting parameters such as convergence rate and convergence ratio.

[0034] In practical applications, constructing synthetic Lyapunov functions We can obtain: .

[0035] .

[0036] .

[0037] .

[0038] In the formula, It is a Lyapunov function. For vector dot product, This is the first derivative of the Lyapunov function with respect to time. Let be the first derivative of the sliding surface function with respect to time. The first derivative of the dual angular velocity with respect to time, Let be the time derivative of the Lyapunov function corresponding to the sliding surface. Let be the α power of the Lyapunov function. It is a positive constant gain.

[0039] Proof obtained At that time, the system satisfies the global asymptotic stability condition throughout the entire control process.

[0040] Furthermore, to achieve time-synchronized stable control, the control conditions can be further refined and improved. Based on the lemmas of finite-time stability and ratio persistence, and combined with the specific characteristics of the spacecraft, conditions that ensure the synchronous and stable convergence of each state component can be determined. This includes the rational selection of control parameters, in-depth analysis of the system's dynamic characteristics, and a trade-off between convergence speed and convergence accuracy. For example, by adjusting parameters related to the moment of inertia and optimizing the performance of the sliding surface, each state component can converge to the target state according to a predetermined ratio and time rhythm. Simultaneously, a comprehensive evaluation of factors that may affect time-synchronized stability, such as system delay and measurement errors, should be conducted, and corresponding compensation measures should be formulated to ensure that the system can reliably achieve time-synchronized stable control in actual operation.

[0041] In an exemplary embodiment of this application, the controller in step 104 above can be an integrated controller that can be directly engineered for use in the spacecraft rendezvous and docking control system. This controller is the core execution unit of the entire spacecraft rendezvous and docking control system, and its specific design and implementation process is as follows: Step 41: Controller Input Acquisition. The six-degree-of-freedom operating status of the controlled spacecraft is acquired in real time through the spacecraft's onboard sensor array. This includes core state variables such as the dual quaternion pose deviation relative to the target spacecraft in the body coordinate system, dual angular velocity, linear velocity and angular velocity deviation, and real-time values ​​of sensor field of view and approximation cone angle. After filtering and preprocessing the raw data, it is input to the control calculation unit.

[0042] Step 42: Core Design of the Control Law. Based on the dual quaternion spacecraft dynamics and kinematic equations, combined with the potential function obtained above, the constructed sliding surface, and the control conditions for the synchronous convergence of each state component, a controller that simultaneously satisfies the constraint characteristics and time synchronization stability requirements is designed, expressed as: .

[0043] In the formula, For spacecraft control forces and torque vectors, For dual cross product matrices, The angular velocity relative to the inertial frame in the controlled spacecraft's own coordinate system. Let ω be the angular velocity of the controlled spacecraft in the target spacecraft's body coordinate system. To exchange the potential functions of the real and imaginary terms, It is the power of the symbolic function. For exponential parameters, . As an adaptive weighting factor, The formula uses the spacecraft's inertial matrix. It can be a controlled spacecraft in a coordinate system The constant known inertia matrix under the given conditions.

[0044] This controller ensures the efficient and stable operation of the entire spacecraft control system by monitoring the status in real time, rapidly calculating and accurately outputting control signals, and promptly handling and fault-tolerantly controlling various fault conditions. The process of control command calculation and allocation can be described as follows: substituting the preprocessed real-time state quantities into the controller's expression formula, first calculating the gradient of the potential function and the real-time value of the sliding mode surface corresponding to the attitude deviation, then completing the real-time calculation of the control law through dual quaternion operations to obtain the total control quantity, and finally allocating the control quantity according to the configuration matrix of the spacecraft's thrusters and attitude actuators to generate the corresponding drive commands for the actuators.

[0045] Based on the above description, the full-process execution logic of the spacecraft rendezvous and docking synchronization control method under line-of-sight constraints provided in this application can be as follows: Figure 2 As shown. Figure 2 The paper sequentially demonstrates the closed-loop control process, from pose model and constraint construction to dual quaternion potential function design, artificial potential field sliding surface construction, time-synchronous stable control condition design, and controller calculation and execution. It clarifies the input-output relationships, data transmission links, and core design objectives of each stage, and reconstructs the complete technical path of this application from spacecraft motion modeling and constraint definition to control algorithm design and controller engineering implementation. This provides intuitive visualization support for understanding the overall control architecture and collaborative working logic of each mode in this application.

[0046] In one exemplary embodiment of this application, corresponding to the time synchronization stability control condition design step in step 103 of this application, Figure 3 The comparison demonstrates the difference in control input between the traditional symbolic function and the unit vector symbolic function used in this application, clearly presenting the core principle and mechanism of achieving synchronous convergence of various state components of the spacecraft, reducing invalid control input, and lowering control energy consumption by passing the ratio persistence constraint for time synchronization stability. This provides a visual theoretical support for the design of the time synchronization convergence control conditions in step 103. Figure 3 middle, x 1 is a vector on the first dimension of the system's state space. x 2 is a vector on the second coordinate axis of the system state space. x 3 is a vector on the third-dimensional coordinate axis of the system state space. x This is the system state vector.

[0047] In an exemplary embodiment of this application, corresponding to the pose model and constraint design step 100 of the present invention, Figure 4 The visualization presents the line-of-sight constraints and the cone-shaped safety zone of the approach corridor constraints during the spacecraft rendezvous and docking process. It clearly defines the line-of-sight axis of the controlled spacecraft, the docking axis of the target spacecraft, and the geometric boundary of the constraint cone half-angle. It fully restores the spacecraft safety operation constraints established in step 100, and provides a geometric model basis for subsequent potential function constraint fusion and sliding mode surface constraint design. Figure 4 middle,( X L , Y L , Z L () represents the coordinates of the target spacecraft's body coordinate system. 0 L Let O be the origin of the target spacecraft's coordinate system. X B , Y B , Z B () represents the coordinates of the controlled spacecraft's body coordinate system. 0 B It is the origin of the coordinate system of the controlled spacecraft.

[0048] In an exemplary embodiment of this application, corresponding to the sliding surface design step 102 of the present invention, Figure 5 The paper demonstrates the high potential energy forbidden zone of the artificial potential field, the low potential energy region of the target docking pose, and the complete process of the sliding surface guiding system state converging towards the target point along a constraint-compliant trajectory. It intuitively presents the core control logic of the combination of artificial potential field function and sliding control in step 102, and reflects the working mechanism of how the sliding surface avoids local minima of the potential function and ensures that the derivative of the potential function is negative throughout the process.

[0049] In one exemplary embodiment of this application, the actual control effect of the controller corresponding to step 104 of the present invention is verified. Figure 6 The paper demonstrates the spatial position and attitude of the controlled spacecraft at key time points such as 0s, 10s, 20s, 35s, and 80s during the rendezvous and docking process, as well as the complete approach trajectory relative to the target spacecraft. It intuitively presents the entire controlled motion process of the spacecraft from its initial pose to the target docking pose under the control of the controller in this application, and verifies the trajectory tracking capability and constraint execution capability of the method provided in this application.

[0050] In an exemplary embodiment of this application, corresponding to the constraint design in step 100 and the potential function constraint fusion step in step 101 of the present invention, Figure 7The image shows the line-of-sight angle α(t) (corresponding to) throughout the controlled process in polar coordinates. Figure 7 (a) part) and the cone angle β(t) (corresponding to Figure 7 The trajectory of the change in part (b) is marked with the preset constraint boundary and control starting point. At the same time, the constraint execution effect of the method provided in this application is compared with that of the traditional method. It clearly shows that this application can strictly limit the line of sight angle and the approach cone angle within the preset safety boundary, and verifies the effective execution of the constraint conditions and the constraint guidance effect of the potential function.

[0051] In one exemplary embodiment of this application, the time synchronization stabilization control effect verification corresponding to step 103 of the present invention is described. Figure 8 The text uses time as the horizontal axis to show the methods used (…). Figure 8 (a) and not using ( Figure 8 The sliding surface convergence process and pose error convergence curve of the time synchronization stabilization strategy in part (b) of this application intuitively demonstrate that the method provided in this application can achieve synchronous convergence of the spacecraft's position, attitude, and velocity components, eliminating the oscillations and overshoots caused by asynchronous convergence of components in traditional methods, and verifying the improvement effect of the time synchronization stabilization strategy on control accuracy and convergence stability.

[0052] In one exemplary embodiment of this application, the core advantages of the method provided in this application are verified. Figure 9 and Figure 10 The study statistically compared the total energy consumption of the method provided in this application (corresponding to TSC in the figure), traditional sliding mode control (corresponding to SMC in the figure), and traditional PD control (corresponding to PD in the figure) at different time periods throughout the entire rendezvous and docking process. It was found that the method provided in this application can achieve lower energy consumption in all control stages, including 0-10s, 10-20s, 20-100s, and 100-300s. The total energy consumption is only about 21% of that of traditional PD control and about 85% of that of traditional sliding mode control, which directly demonstrates the energy consumption optimization advantages and engineering application value of the method provided in this application.

[0053] In summary, the spacecraft rendezvous and docking synchronization control method provided in this application, by constructing a kinematic and dynamic model of the spacecraft using dual quaternions, integrates position and attitude information into a unified framework, effectively reducing model complexity and improving computational efficiency. Based on this model, an advanced control algorithm is designed, fully considering time synchronization factors during the control process, enabling precise coordination of control commands for each degree of freedom. This overcomes the coordination deficiencies and time lag problems of traditional control methods in multi-degree-of-freedom control, significantly improving the stability and accuracy of spacecraft control. It can be widely applied to various spacecraft orbit adjustment and attitude stabilization control tasks, and is of great significance for enhancing spacecraft operational reliability and expanding the application scenarios of space missions, powerfully promoting the further development of aerospace technology.

[0054] Based on the same inventive concept, this application also provides a spacecraft rendezvous and docking synchronization control system for implementing the aforementioned spacecraft rendezvous and docking synchronization control method. The solution provided by this system is similar to the implementation scheme described in the above method; therefore, the specific limitations of one or more spacecraft rendezvous and docking synchronization control system embodiments provided below can be found in the limitations of the spacecraft rendezvous and docking synchronization control method described above, and will not be repeated here.

[0055] In one exemplary embodiment, such as Figure 11 As shown, a spacecraft rendezvous and docking synchronization control system is provided, including: a pose model and constraint design module, a potential function design module based on dual quaternions, an artificial potential field function sliding surface design module, a time synchronization stable control condition design module, and a controller design module.

[0056] The pose model and constraint design module considers the physical characteristics of the spacecraft and space environment factors. Based on dual quaternions, it constructs a pose model to describe the six degrees of freedom motion of the spacecraft and formulates constraints according to the requirements of the space mission. For example, the six-degree-of-freedom motion pose model of the spacecraft, constructed based on dual quaternions, comprehensively considers the spacecraft's physical characteristics such as mass distribution and inertial tensor, as well as space environment factors such as gravitational field and atmospheric drag. According to the requirements of the space mission, it clarifies key conditions such as the orbital altitude range, attitude angle limits, and visual sensor field-of-view constraints, ensuring that the spacecraft's motion always meets safety and mission execution requirements, providing accurate model support for subsequent control modules. In particular, formulating constraints such as orbital altitude, attitude angle, and field-of-view angle in conjunction with mission requirements ensures that the spacecraft operates within a safe and effective range.

[0057] The potential function design module based on dual quaternions is used to construct potential functions reflecting the spacecraft's attitude state based on the pose model and constraints. For example, leveraging the unique mathematical expression and properties of dual quaternions, a potential function reflecting the deviation between the spacecraft's real-time pose and the target pose is constructed. This potential function satisfies the core condition that it is 0 only at the target pose, and positive at all other poses, with its derivative always remaining negative definite to guide the spacecraft towards the target pose. Simultaneously, the potential function structure is optimized to balance anti-interference performance and energy consumption control, reducing propellant consumption and extending the spacecraft's lifespan. By guiding the spacecraft towards the target pose through potential energy changes, the potential function is optimized to improve anti-interference capabilities and reduce energy consumption.

[0058] The artificial potential field function sliding surface design module is used to construct a sliding mode control strategy based on dual quaternions based on the potential function, and obtain the sliding surface. For example, by combining the above potential function to construct a dual quaternion sliding mode control strategy, the sliding surface expression is determined. The operational characteristics of dual quaternions are used to simplify calculations, improve control real-time performance and accuracy, adjust sliding mode parameters to enhance the system's anti-interference capability, and achieve precise trajectory tracking of spacecraft.

[0059] Furthermore, by combining sliding mode control theory, the expression for the sliding surface is derived, transforming constraints such as line-of-sight constraints into sliding surface constraints. The computational properties of dual quaternions are utilized to simplify the control algorithm's calculation process and improve real-time control performance. Adjusting the sliding mode control parameters enhances the system's anti-interference capability, ensuring that the spacecraft can stably track the target trajectory and maintain the sliding surface convergence and maintenance conditions even when facing external disturbances in complex space environments.

[0060] The Time Synchronization Stability Control Condition Design Module is used to determine the control conditions for the synchronous convergence of each state component based on the theory of time synchronization stability. For example, based on finite-time stability and the ratio persistence lemma, parameters such as convergence rate and convergence ratio are set, and the spacecraft's moment of inertia is combined with the sliding mode surface to ensure the synchronous and orderly convergence of state components such as attitude and velocity. The module also analyzes system stability and formulates compensation measures for delays and measurement errors.

[0061] Furthermore, based on the finite-time stability and ratio persistence lemma, control parameters for the synchronous convergence of all state components of the spacecraft are determined, and the convergence rate and convergence ratio requirements are clarified. The spacecraft's moment of inertia is combined with the sliding surface design to ensure that the sliding surface possesses ratio consistency characteristics, guaranteeing that state components such as position, attitude, and velocity converge in an orderly manner according to predetermined ratios. Influencing factors are assessed, and targeted compensation measures are formulated. Theoretical derivation and simulation verification ensure the stability of the control system throughout the entire process, avoiding oscillations or divergences.

[0062] The controller design module is used to construct a controller based on the sliding mode surface and control conditions. The controller generates control signals based on real-time acquired spacecraft status to achieve rendezvous and docking control. For example, based on dual quaternion dynamics and kinematic equations, combined with a sliding mode surface containing a rotational inertia matrix, integrating potential function properties and time synchronization stability requirements, and comprehensively considering spacecraft position, attitude, velocity, and other state data, it calculates and outputs control commands to the actuators to achieve fault-tolerant control and fault handling.

[0063] Furthermore, based on the dual quaternion dynamics equations and kinematic equations, the design results of sliding mode surfaces containing rotational inertia matrices are integrated, incorporating potential function properties and time synchronization stability requirements to construct a core control algorithm. Real-time acquisition of spacecraft position, attitude, velocity, and other state data is performed, and control commands are rapidly calculated after comprehensive analysis to ensure efficient and stable operation during the spacecraft rendezvous and docking process.

[0064] Based on the above description, compared with the prior art, this application has at least the following advantages: (1) This application proposes a sliding mode control method for spacecraft based on dual quaternions, which has higher anti-interference capability and lower energy consumption on the basis of the original artificial potential field function control. It can simultaneously represent the translational and rotational motions of the spacecraft in a concise and unified mathematical form, which greatly reduces the computational complexity and improves the execution efficiency of the control algorithm compared with the traditional description method. By optimizing the control strategy, unnecessary control actions are reduced, enabling the spacecraft to utilize energy more efficiently, avoiding excessive propellant consumption, thereby extending the service life of the spacecraft and reducing mission costs.

[0065] (2) This application combines the concept of time synchronization stability with existing six-degree-of-freedom control methods to achieve consistent and orderly convergence of all state components, resulting in higher control accuracy and lower energy consumption. Based on the concept of time synchronization stability, it ensures that the spacecraft's position, attitude, and velocity, among other state components, converge to the target state consistently and orderly. During the control process, the principle of ratio continuity is strictly followed, ensuring that each state component converges according to a predetermined ratio. By precisely controlling the convergence speed of position and attitude, they are synchronized in time, avoiding system instability caused by excessively fast or slow convergence of a certain state component. This precise control method not only improves control accuracy and reduces error accumulation but also effectively avoids excessive energy consumption due to over-adjustment.

[0066] In one exemplary embodiment, a computer device is provided, which may be a server or a terminal, and its internal structure diagram may be as follows. Figure 12As shown, this computer device includes a processor, memory, input / output (I / O) interfaces, and a communication interface. The processor, memory, and I / O interfaces are connected via a system bus, and the communication interface is also connected to the system bus via the I / O interfaces. The processor provides computational and control capabilities. The memory includes non-volatile storage media and internal memory. The non-volatile storage media stores the operating system, computer programs, and a database. The internal memory provides the environment for the operating system and computer programs stored in the non-volatile storage media. The database stores spacecraft rendezvous and docking synchronization control data. The I / O interfaces are used for information exchange between the processor and external devices. The communication interface is used for communication with external terminals via a network connection. When the computer program is executed by the processor, it implements a spacecraft rendezvous and docking synchronization control method.

[0067] Those skilled in the art will understand that Figure 12 The structure shown is merely a block diagram of a portion of the structure related to the present application and does not constitute a limitation on the computer equipment to which the present application is applied. Specific computer equipment may include, for example, [the following is a list of possible additional structures]. Figure 12 The diagram shows more or fewer components, or combinations of certain components, or different component arrangements.

[0068] In one exemplary embodiment, a computer device is provided, including a memory and a processor, wherein the memory stores a computer program, and the processor executes the computer program to implement the steps in the above-described method embodiments.

[0069] In one exemplary embodiment, a computer-readable storage medium is provided storing a computer program that, when executed by a processor, implements the steps in the above-described method embodiments.

[0070] In one exemplary embodiment, a computer program product is provided, including a computer program that, when executed by a processor, implements the steps in the above-described method embodiments.

[0071] It should be noted that the user information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data used for analysis, data stored, data displayed, etc.) involved in this application are all information and data authorized by the user or fully authorized by all parties, and the collection, use and processing of the relevant data must comply with relevant regulations.

[0072] Those skilled in the art will understand that all or part of the processes in the above embodiments can be implemented by a computer program instructing related hardware. The computer program can be stored in a non-volatile computer-readable storage medium, and when executed, it can include the processes of the embodiments described above. Any references to memory, databases, or other media used in the embodiments provided in this application can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (ROM), magnetic tape, floppy disk, flash memory, optical memory, high-density embedded non-volatile memory, resistive random access memory (RRAM), magnetic random access memory (MRAM), ferroelectric random access memory (FRAM), phase change memory (PCM), graphene memory, etc. Volatile memory can include random access memory (RAM) or external cache memory, etc. By way of illustration and not limitation, RAM can take many forms, such as Static Random Access Memory (SRAM) or Dynamic Random Access Memory (DRAM).

[0073] The databases involved in the embodiments provided in this application may include at least one type of relational database and non-relational database. Non-relational databases may include, but are not limited to, blockchain-based distributed databases. The processors involved in the embodiments provided in this application may be general-purpose processors, central processing units, graphics processing units, digital signal processors, programmable logic devices, quantum computing-based data processing logic devices, etc., and are not limited to these.

[0074] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0075] This document uses specific examples to illustrate the principles and implementation methods of this application. The descriptions of the above embodiments are only for the purpose of helping to understand the methods and core ideas of this application. Furthermore, those skilled in the art will recognize that, based on the ideas of this application, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of this application.

Claims

1. A method for synchronous control of spacecraft rendezvous and docking, characterized in that, include: Considering the physical characteristics of the spacecraft and the space environment factors, a pose model for describing the six-degree-of-freedom motion of the spacecraft is constructed based on dual quaternions, and constraints are formulated according to the requirements of the space mission. Based on the pose model and the constraints, a potential function is constructed to reflect the pose state of the spacecraft. Based on the potential function, a sliding mode control strategy based on dual quaternions is constructed to obtain the sliding surface; Based on the theory of time-synchronous stability, the control conditions for the synchronous convergence of each state component are determined. A controller is constructed based on the sliding surface and the control conditions; The controller generates control signals based on the real-time acquired spacecraft status to achieve rendezvous and docking control of the spacecraft.

2. The spacecraft rendezvous and docking synchronization control method according to claim 1, characterized in that, Based on the pose model and the constraints, a potential function reflecting the spacecraft's pose state is constructed, including: Based on the pose model and the constraints, the dual quaternion deviation between the spacecraft's real-time pose and the target pose is determined. The dual quaternion deviation between the spacecraft's real-time pose and the target pose is used as input to construct a potential function that reflects the spacecraft's pose state. The potential function takes a value of zero at the target pose and a value of positive at other poses. The derivative of the potential function is used to guide the spacecraft to move toward the target pose.

3. The spacecraft rendezvous and docking synchronization control method according to claim 1, characterized in that, Based on the potential function, a sliding mode control strategy based on dual quaternions is constructed to obtain the sliding surface, including: Based on the potential function, a sliding mode control strategy based on dual quaternions is constructed to obtain the initial sliding surface; The initial glial surface is optimized by combining the inertial matrix of the spacecraft to obtain the glial surface.

4. The spacecraft rendezvous and docking synchronization control method according to claim 3, characterized in that, The synovial surface is represented as follows: ； In the formula, For sliding mode variables, For the inertial matrix of the spacecraft, For the dual angular velocity of the spacecraft, These are the sliding mode control parameters. Auxiliary state variables constructed for the sliding surface. This represents the operational law for exchanging the real and imaginary parts of an even number. This represents the Hadamard product.

5. The spacecraft rendezvous and docking synchronization control method according to claim 1, characterized in that, Based on the theory of time-synchronous stability, the control conditions for the synchronous convergence of each state component are determined, including: A synthetic Lyapunov function is constructed, and based on the theory of time-synchronous stability, the control conditions for the synchronous convergence of each state component are determined through derivation and simulation verification.

6. The spacecraft rendezvous and docking synchronization control method according to claim 1, characterized in that, The controller is represented as: ； In the formula, A controller for representing force and torque, For dual cross product matrices, For the inertial matrix of the spacecraft, Let be the dual angular velocity vector of the controlled spacecraft in its body coordinate system. for, For dual angular velocity, Let be the dual angular velocity vector of the controlled spacecraft in the target point coordinate system. These are the sliding mode control parameters. Represents the Hadamard product. For vectorization operators, Let be the conjugate form of the relative dual quaternions between the target and the controlled spacecraft. Represents the Kronecker product. For the gradient operator in the dual quaternion space, Let be the potential function. For sliding mode variables, Let be the sign function of the power-to-reach law. This represents the operational law for exchanging the real and imaginary parts of an even number; For the exponent parameter of the power-law approach, ; As an adaptive weighting factor, .

7. A spacecraft rendezvous and docking synchronization control system, characterized in that, include: The pose model and constraint design module is used to consider the physical characteristics of the spacecraft and space environment factors, construct a pose model based on dual quaternions to describe the six degrees of freedom motion of the spacecraft, and formulate constraints according to the requirements of the space mission. The potential function design module based on dual quaternions is used to construct a potential function that reflects the attitude state of the spacecraft based on the pose model and the constraints. The artificial potential field function sliding surface design module is used to construct a sliding mode control strategy based on dual quaternions based on the potential function to obtain the sliding surface; The time synchronization stability control condition design module is used to determine the control conditions for the synchronous convergence of each state component based on the theory of time synchronization stability. The controller design module is used to construct a controller based on the sliding surface and the control conditions; the controller generates control signals based on the real-time acquired spacecraft status to realize the rendezvous and docking control of the spacecraft.

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