Variable configuration spacecraft system and method for non-cooperative target capture

Abstract

The application provides a variable configuration spacecraft system and method for non-cooperative target capture, and belongs to the field of deep space exploration. The method comprises the following steps: 1) a rigid combination flies to the vicinity of the target; 2) the combination is spun to be expanded into a ring belt; 3) the ring belt envelopes the target; and 4) the ring belt encloses the target. Large space debris and small celestial bodies both belong to non-cooperative targets, and cannot provide special interfaces for assisting capture. It is extremely difficult to develop general control mechanisms and methods to realize reliable capture and stable connection of various targets. In order to solve the problem, the application utilizes the inter-satellite connection to expand a plurality of controllable satellites into a ring-shaped multi-body tether satellite formation system, and carries out ring belt enveloping and then enclosing the tether to realize capture of a non-cooperative target with large scale, complex geometric shape, great uncertainty of motion state and inertia information according to a predetermined trajectory. The application has the advantages of high applicability, flexibility, high intelligence, reusability and small size.

Description

TECHNICAL FIELD

[0001] The application relates to a variable-configuration spacecraft system and method for capturing a non-cooperative target, and belongs to the technical field of deep space exploration. BACKGROUND

[0002] With the increasing strategic position of space in the fields of politics, economy, and military, space game competition is particularly prominent as the core ability of a major power, and is an important basis for seizing the right to control space and information in future wars, and capturing a non-cooperative target is one of the important means in space game competition. In addition, the seriousness and urgency of the space debris problem have been highly valued by various spacefaring countries and organizations, and since large space debris (abandoned spacecraft / rocket bodies) can induce a cascading collision effect of space debris, large debris in near-Earth orbit has become the primary object of debris removal. At the same time, with the in-depth and rapid development of space technology, the scope of human exploration has gradually expanded to the remote deep space, and asteroids in the solar system are the initial formation of the solar system, and in-depth research on asteroids is expected to reveal the origin, composition, and evolution mechanism of the solar system. Large space debris and small celestial bodies are both non-cooperative targets, and cannot provide a dedicated interface for auxiliary capture, so it is extremely difficult to develop a general manipulation mechanism and method to achieve reliable capture and stable connection of various targets. The difficulty lies in: first, there are many uncertain factors. The mass, rotation characteristics, topography, rocks, and weathering layer of asteroids cannot be directly obtained at present, and can only be calculated through model analysis, comparison with other small celestial bodies and ground meteorites, etc., which has great uncertainty, and the inertia and motion parameters of space debris are unknown. Second, the characteristics of the captured targets are quite different. Large space debris, asteroids, and a large number of small celestial bodies have great differences in shape and structure size, the rotation period of asteroids is from hundreds of seconds to tens of days, and asteroids have weak gravitational field, irregularity, and other characteristics.

[0003] Therefore, the skilled in the art is committed to developing a new method for capturing a non-cooperative target to solve the problem of insufficient existing expansion means. SUMMARY

[0004] The purpose of the present application is to solve the problems existing in the prior art, and to provide a variable-configuration spacecraft system and method for capturing a non-cooperative target.

[0005] The purpose of the present application is achieved by the following technical solutions:

[0006] The application provides a variable configuration spacecraft system for non-cooperative target capture, which is characterized by comprising a main node, a secondary node, a mechanical interface, an electric thruster, a rope winding and unwinding mechanism and a solar wing.

[0007] The rope-driven variable spacecraft operation scheme (enclosure process) is divided into four stages, namely, a combination body stage, an unfolding stage, an envelope stage and a closing stage.

[0008] 1) The rigid combination body flies to the vicinity of the target: before reaching the target capture position, the ropes are wound in the rope winding and unwinding mechanism, the main node and the secondary node of the variable configuration spacecraft are fixed through the interface, and the combination body is a complete annular rigid combination body, and the main node and the secondary node spacecraft move to the target position by using the electric thruster;

[0009] 2) The combination body is spun to form a ring belt: after reaching the target capture position, the rope winding and unwinding mechanism releases the ropes, and the rigid combination body is unfolded into a ring belt connected by the nodes under the propulsion of the nodes;

[0010] 3) The ring belt envelopes the target: the non-cooperative target is observed to obtain its geometric shape and motion state, the Kalman filtering method is used to predict the motion, the dynamic game method and the artificial potential field method are used to plan the expected envelope configuration, the main node and the secondary node track the expected configuration under the action of the electric thruster, and the fuel consumption, safety and follow-up motion state maintenance factors consumed by orbit transfer are considered, the weight coefficients of the reconstruction fuel consumption function and the collision risk assessment function are considered, the configuration competition degree concept containing the above constraints is introduced as the target function, the local or global optimal result under multiple constraint conditions is obtained by calculating the extreme value of the target function, after preliminary screening, the adaptive coarse-grained parallel genetic algorithm, the pigeon algorithm and the neural network intelligent algorithm are used to search for the optimal solution in the solution space, and the optimal solution required is obtained, and the non-cooperative target is enveloped in the formation ring belt by using the continuous small thrust generated by the electric thruster of the main node and the secondary node, so that the ropes are kept in a relaxed state;

[0011] 4) The ring belt closes the target: when each node completes the envelope of the non-cooperative target, the rope winding and unwinding mechanism tightens the ropes, and the follow-up state and the final node realize the closing state by tightening the ropes.

[0012] The present invention provides a variable-configuration spacecraft system and method for capturing non-cooperative targets, which has the following advantages: (1) It is flexible and versatile, and the rope length can be adjusted to capture targets of different sizes; (2) It has no special requirements for the surface conditions of the target, because its working principle is that after the spacecraft completes the envelopment of the non-cooperative target, it tightens the rope so that the spacecraft tightly binds the target. It relies on the tightening of the rope to make the spacecraft firmly attached to the surface of the target, especially when exploring asteroids, it does not depend on the surface geological conditions; (3) It can use the rope tightening force to provide sufficient adhesion for drilling and sampling, and solve the adsorption problem under weak gravity conditions. Attached Figure Description

[0013] Figure 1 Design drawings for a variant spacecraft structure;

[0014] Figure 2 Flowchart for the operation of a tethered variant spacecraft;

[0015] Figure 3 This is a vector diagram of the spacecraft coordinate system and other data based on dual quaternions.

[0016] Figure 4 A schematic diagram of rope dynamics modeling;

[0017] Figure 5 This is a schematic diagram of the drag coordinate system and the body coordinate system;

[0018] Figure 6 This is a block diagram of the feedback control system for a spacecraft. Detailed Implementation

[0019] The present invention will be further described in detail below with reference to the accompanying drawings: This embodiment is implemented under the premise of the technical solution of the present invention, and detailed implementation methods are given, but the protection scope of the present invention is not limited to the following embodiments.

[0020] Example 1: This example relates to a variable-configuration spacecraft system and method for capturing non-cooperative targets, such as... Figure 1 As shown, this variable-configuration spacecraft consists of a main node, secondary nodes, a mechanical interface, electric propulsion, a cable retrieval mechanism, and solar panels. The main and secondary nodes are connected via mechanical interfaces. Each main and secondary node has electric propulsion and solar panels, and each main and secondary node contains a cable retrieval mechanism.

[0021] The tethered variant spacecraft operation plan (encirclement process) is divided into four stages, such as... Figure 2 As shown, the stages are the combination stage, the unfolding stage, the envelope stage, and the encirclement stage, respectively:

[0022] 1) The rigid combination flies to the vicinity of the target: before reaching the capture target position, the rope is collected in the rope winding and unwinding mechanism, the main and secondary nodes of the variable configuration spacecraft are fixed by the interface, and it is a complete ring-shaped rigid combination, which moves to the target position by using the electric propulsion on the main and secondary node spacecrafts;

[0023] 2) The combination is spun to develop into a ring belt: after reaching the capture target position, the rope winding and unwinding mechanism releases the rope, and the rigid combination develops into a rope-connected formation ring belt under the propulsion of the nodes;

[0024] 3) Envelope stage: observe the non-cooperative target to obtain its geometric shape and motion state, predict its motion by using Kalman filtering method, plan the expected envelope configuration by using dynamic game method and artificial potential field method, and track the expected configuration by the main node (1) and the secondary node (2) under the action of the electric propulsion (4), considering the fuel consumption, safety and the maintenance consumption factors of the follow-up motion state, comprehensively considering the key constraint factors in different stages, respectively assigning weight coefficients to the reconstruction fuel consumption function and the collision risk assessment function, introducing the configuration competitiveness concept containing the above constraints as the objective function, calculating the extreme value of the objective function to obtain the local or global optimal result under multiple constraint conditions, after preliminary screening, using adaptive coarse-grained parallel genetic algorithm, pigeon algorithm and neural network intelligent algorithm to search for the optimal solution in the solution space, obtaining the required optimal solution, and using the continuous small thrust generated by the electric propulsion (4) on the main node (1) and the secondary node (2) to envelope the non-cooperative target in the formation ring belt, so that the rope remains in a relaxed state;

[0025] 4) Ring belt surrounding stage: a parking area is planned in advance on the non-cooperative target, the nodes track the expected surrounding configuration, the rope winding and unwinding mechanism tightens the rope, and the final node realizes the surrounding state by tightening the rope.

[0026] The technical route for the variable body spacecraft variable and formation process dynamics modeling is as follows: firstly, the relative attitude and orbit integrated dynamics modeling of distributed nodes is carried out, in the dual algebra framework, the body coordinate system of the inertia coordinate system, the reference coordinate system with the non-cooperative target mass center as the reference, the node mass center system and the expected coordinate system in the expected configuration are established, the coordinate system and other vectors are shown as follows: Figure 3 The main coordinate systems involved are as follows:

[0027] 1) Inertial coordinate system: the small planet in this paper is a small planet in the solar system, so the heliocentric ecliptic coordinate system is taken as the inertial system, denoted as Ψ I .

[0028] 2) Relative motion reference coordinate system: the coordinate origin O is located at the mass center of the asteroid, the x-axis points from the sun to the asteroid, the z-axis is along the angular momentum direction of the asteroid, and the y-axis forms a right-handed orthogonal coordinate system with the x-axis and the z-axis.

[0029] 3) Body coordinate system: The body coordinate system of the i-th composite node is denoted as Ψ. i The origin is located at the centroid of the i-th node, and the three coordinate axes coincide with the principal inertial axes of the spacecraft. The asteroid's body coordinate system is denoted as Ψc, and the three coordinate axes coincide with the principal inertial axes of the asteroid.

[0030] 4) Desired coordinate system: The coordinate system of the i-th combined node is denoted as Ψ. Ti This coordinate system is determined according to task requirements and is a virtual coordinate system that describes the coordinate system Ψ of the composite object's nodes. i The final expected relative state.

[0031] The unit dual quaternion can be used to describe the six degrees of freedom motion of a coordinate system, including both rotational and translational motion. For coordinate system Ψ... c To coordinate system Ψ i The transformation can be achieved using the quaternion q. ic Rotation followed by translational motion p ic Alternatively, it can be translated first and then rotated, which can be described using dual quaternions:

[0032]

[0033] In the formula, Coordinate system Ψ c To coordinate system Ψ i The dual quaternion of the transformation, q ic Let the coordinate system of the i-th composite node be relative to the coordinate system Ψ of the asteroid. c The posture quaternion, and The position vectors of the center of mass of the i-th composite node relative to the center of mass of the asteroid in coordinate system Ψ are respectively. c and coordinate system Ψ i The following amount, For the conjugate of the dual quaternions of the reference coordinate system relative to the inertial coordinate system, Let ε be the dual quaternion of the volume coordinate system of node i relative to the inertial coordinate system, and let ε be the dual element.

[0034] In engineering, the kinematic equations of rigid body rotation are generally expressed in quaternion form as follows:

[0035]

[0036] In the formula, ω i ic ω c ic Let ω be the angular velocity vector of the center of mass of the i-th composite node relative to the center of mass of the asteroid in coordinate system Ψ. iand the coordinate system Ψ c .

[0037] Taking the derivative of equation (1) and substituting equation (2), we have

[0038]

[0039] Definition:

[0040]

[0041] We have

[0042]

[0043]

[0044] To study the dynamic equation of the coordinate system Ψ c relative to the inertial system Ψ I , we first give the following dual vector:

[0045]

[0046] In the formula, is the dual angular momentum of the ith combined body node, is the dual angular velocity of the ith combined body node relative to the inertial system, which satisfies the definition, is the dual inertia matrix of the ith combined body node, which is specifically defined as:

[0047]

[0048] In the formula, m i , J i are the mass and moment of inertia of the ith combined body node, I 3×3 is a three-order unit matrix.

[0049] Taking the derivative of equation (8), we have

[0050]

[0051] In the formula, is the external force acting on each node.

[0052] From the definition and difference characteristics of , we have It can also be written as:

[0053]

[0054] In the formula, is the body coordinate system of the ith combined body node relative to the asteroid body coordinate system Ψ cthe conjugate of the attitude quaternion of the i-th spacecraft, the dual angular velocity of the asteroid body with respect to the asteroid body coordinates.

[0055] Differentiate equation 11 and substitute equation 10 and 7 into it, we have

[0056]

[0057] When the relative velocity is easy to measure, equation 11 can be rewritten as

[0058]

[0059]

[0060] where, is the dual external force of node i, is the dual control force, is the dual gravity, including the sun gravity and the asteroid irregular weak gravity, is the bounded unpredictable dual disturbance force, is the dual cable force of the i-th combined node received from other nodes.

[0061] On this basis, the dynamics modeling of the tether system between nodes is carried out, as shown in equation 12. Figure 4 The motion state of the spacecraft at both ends of the cable is taken as the boundary condition, the model is simplified into a variable length multi-rigid body series link, the mass is concentrated in the spherical hinge, and there is no friction, no axial expansion, no torsion, and no material damping between adjacent links. The force of the cable is considered as gravity and weak gravity, the kinematics equation of the link and the dynamics equation of the spherical hinge are established, and the tension iteration matrix is iterated to solve. The solved cable drag force is converted into the dual cable force of the nodes at both ends combined with the Laplacian matrix of the cable connection topology, and the detailed process is as follows:

[0062] In order to better model the dynamics of the cable, the body coordinate system and the drag coordinate system are defined, as shown in equation 13. Figure 5

[0063] Body coordinate system: this coordinate system is consistent with the node body coordinate system, the body coordinate system of the i-th combined node is denoted as i , the origin is located at the mass center of the i-th node, and the directions of the three coordinate axes are respectively coincident with the principal axes of inertia of the spacecraft.

[0064] Drag coordinate system: since the cable connection point does not necessarily pass through the mass center, and in order to facilitate subsequent theoretical derivation, this coordinate system is specially built. The three coordinate axes of the drag coordinate system are parallel to the body coordinate system, only the origin of the coordinate system is placed at the drag point connected with node j, denoted as

[0065] ​In order to distinguish the cable force in different coordinate systems, the cable force is defined as the pair cable force and the drag force, as follows Figure 6 The drag force is the force of the cable between two nodes acting on the drag point, and the pair cable force is the resultant force of all the cable drag forces connected to the node in the body coordinate system. The specific definitions are as follows:

[0066] The cable drag force acting on the drag point of node i and the drag point of node j is defined as:

[0067]

[0068] In the formula, are the x, y, and z components of the cable drag force acting on the drag point of node i and the drag point of node j.

[0069] The cable drag force acting on the drag point of node i and the drag point of node j is converted to the center of mass of the combination node, which can be expressed as:

[0070]

[0071] In the formula, is the pair cable force of the i-th node in the body coordinate system, are the cable force and cable moment in the body coordinate system of the node, respectively, and L b is the displacement vector of the center of mass of the i-th node pointing to the cable inlet and outlet.

[0072] The pair cable force of node i is obtained by adding all the drag force vectors connected to node i, which can be expressed as:

[0073]

[0074] The pair cable force to be solved has been converted to the drag force between the two drag points to be solved.

[0075] Assuming that there is a cable connection between two nodes i and j, the linear velocity and acceleration of the drag point of node i are defined as and The values can be calculated by the motion state of the center of mass of the node and the displacement vector of the center of mass of the node pointing to the cable inlet and outlet. Node j is defined in the same way, and will not be repeated here.

[0076] In the drag coordinate system , the following definitions are made:

[0077] The cable is divided into equal length segments with variable length, and the mass is concentrated in the spherical hinge. From the geometric relationship, we have:

[0078] r k = r k-1 + δ k (16)

[0079] In the formula, r k ,r k-1 Spherical hinges k and k-1 are respectively located in the coordinate system. The position vector in δ k Let be the spatial distance vector from spherical hinge k to spherical hinge k-1.

[0080] δ k In terms of unit vector and length, the specific form is:

[0081] δ k =l k n = l k [sinθ k1 cosθ k2 cosθ k1 cosθ k2 sinθ k2 ] T (17)

[0082] In the formula, l k The length of link k varies with time, where n is δ. k A unit vector with independent variable θ. k1 ,θ k2 These correspond to link k relative to plane Z, respectively. w O w Y w and X w O w Y w The deflection angle.

[0083] Taking the first and second derivatives of Equation 17 respectively, we get:

[0084]

[0085] And due to dragging the coordinate system With respect to the inertial frame of reference, attitude changes occur. Differentiating Equation 12 yields:

[0086]

[0087] The derivation is easy to obtain:

[0088]

[0089] From the above relationships, we can see that:

[0090]

[0091] Multiply equation 20 by... and Substituting 15 and 16, we get:

[0092]

[0093]

[0094] The motion equation of the link k is obtained, where:

[0095]

[0096] where There are 12 independent variables, and the motion of all links can be solved by iterative calculation.

[0097] Then the dynamics of the link is analyzed.

[0098] According to Newton's second law, the acceleration a k of the spherical hinge k is:

[0099]

[0100] where Q k is the external force on the spherical hinge k, mainly the sun's gravity and the asteroid's gravity, t k and t k+1 are the internal tension of the link k and the link k+1. m k = l k ρ is the mass of the link k (concentrated at the corresponding spherical hinge), and ρ is the density of the rope.

[0101] Since t k is the internal force of the rope, it cannot be directly measured, and external constraints need to be introduced for real-time calculation.

[0102] From equation 18, the following constraints are satisfied:

[0103]

[0104] Taking the second-order derivative of equation 26, we get:

[0105]

[0106] And there is the following relationship:

[0107]

[0108] n k = [sinθ k1 cosθ k2 ,cosθ k1 cosθ k2 ,sinθ k2 ] T (29)

[0109]

[0110]

[0111] t k = -t k n k (32)

[0112] From the adjacent link acceleration relationship, we have:

[0113]

[0114] Substitute equation 26 into 23 to obtain the algebraic linear equations of adjacent link tension:

[0115]

[0116] In particular, for the starting boundary point, we have the relationship:

[0117]

[0118] Starting boundary algebraic equation:

[0119]

[0120] Convert to tension form:

[0121]

[0122] For the end boundary, we have the end boundary link acceleration relationship:

[0123]

[0124] End boundary link acceleration algebraic equation:

[0125]

[0126] End boundary link tension algebraic equation:

[0127]

[0128] Write the obtained tension algebraic equations in matrix form, we have:

[0129]

[0130] By solving, we can get the internal stress between the links, and by iterative solution, we can get the rope state. Thus, the rope modeling between nodes is completed.

[0131] The surrounding capture configuration planning and intelligent control strategy idea is as follows: according to four constraint conditions of non-cooperative target motion state, geometric shape, tracking process fuel loss and collision safety, the expected parking configuration, envelope configuration and surrounding configuration are autonomously planned and reconfigured and optimized. Under the single constraint condition, such as safety constraint, the safety obstacle avoidance distance is set, and the artificial potential field method is used to realize the collision avoidance between nodes and the collision between nodes and obstacles by using the potential function of the relative distance between nodes as the variable. Under the multi-constraint condition, the fuel consumption, safety and the maintenance consumption of the follow-up motion state consumed by the variable orbit need to be considered, the weight coefficients are assigned to the reconfiguration fuel consumption function and the collision risk assessment function according to the focus of different stages, the configuration competition degree concept containing the above constraints is introduced as the target function, the local or global optimal result under the multi-constraint condition is obtained by calculating the extreme value of the target function, and the configuration optimization is actually a search problem of multiple solutions. After preliminary screening, intelligent algorithms such as adaptive coarse-grained parallel genetic algorithm, pigeon algorithm and neural network can be used to search for the optimal solution in the solution space.

[0132] The control block diagram of the node is as shown in Figure 6 The expected position, attitude and velocity of each node in the expected configuration are taken as expected inputs (the expected position of the node relative to the 6-DOF transformation dual quaternion of the non-cooperative target and the dual screw), the node motion state feedback (the actual position of the node relative to the 6-DOF transformation dual quaternion of the non-cooperative target and the dual screw) is obtained, the relative error is obtained, the distributed adaptive trajectory tracking controller is used, the electric thrust is used as the actuator, and the control force and control moment are input into the dynamics model of the node. The rope dynamics and the node dynamics are strongly coupled, that is, the node dynamics model contains the dual rope force, and the rope dynamics model depends on the motion state of the nodes at both ends of the rope, so the dual control force of the node is executed first, and then the rope winding and unwinding mechanism is iterated.

[0133] The above is only the preferred specific embodiment of the present application, these specific embodiments are different implementation manners based on the overall concept of the present application, and the protection scope of the present application is not limited thereto, any person skilled in the art can easily think of changes or replacements within the technical range disclosed by the present application, which should be covered within the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A variable configuration spacecraft system for non-cooperative target capture, characterized by, The variable configuration spacecraft system for capturing non-cooperative target comprises a main node (1), a secondary node (2), a mechanical interface (3), an electric thruster (4), a rope winding and unwinding mechanism (5) and a solar wing (6), the electric thruster (4), the solar wing (6) and the mechanical interface (3) are arranged on each main node (1) and secondary node (2), a set of rope winding and unwinding mechanisms (5) are arranged in each main node (1) and secondary node (2), adjacent nodes are connected through ropes and the mechanical interface (3), and the excess ropes are retracted in the rope winding and unwinding mechanisms (5), the solar wing (6) absorbs solar energy and charges each node respectively; The variable configuration spacecraft system is a heterogeneous combination, which comprises one main node and n secondary nodes, wherein n is greater than or equal to 2; The implementation method of the variable configuration spacecraft system comprises the following steps: 1) flying the combination to the vicinity of the non-cooperative target: using the electric thrusters on the main node (1) and the secondary node (2) to perform orbit maneuvering, so that the combination approaches the non-cooperative target; 2) spinning to expand into a ring belt: after reaching the position of the non-cooperative target, the mechanical interfaces (3) between the nodes are disconnected, the rope winding and unwinding mechanisms (5) release the ropes, and the combination is expanded into a ring belt connected by ropes under the propulsion of the electric thrusters (4) on the main node (1) and the secondary node (2); 3) enveloping the target with the ring belt: observing the non-cooperative target to obtain its geometric shape and motion state, predicting its motion using Kalman filtering method, planning an expected envelope configuration using dynamic game method and artificial potential field method, and tracking the expected envelope configuration under the action of the electric thrusters (4) on the main node (1) and the secondary node (2), considering the fuel consumption, safety and maintenance consumption factors of the orbit transfer, comprehensively considering the key constraints in different stages, respectively assigning weight coefficients to the reconstruction fuel consumption function and the collision risk assessment function, introducing the configuration competitiveness concept containing the above constraints as the objective function, calculating the extreme value of the objective function to obtain the local or global optimal result under multiple constraint conditions, and after preliminary screening, using the adaptive coarse-grained parallel genetic algorithm, pigeon algorithm and neural network intelligent algorithm to search for the optimal solution in the solution space, obtaining the required optimal solution, and using the continuous small thrust generated by the electric thrusters (4) on the main node (1) and the secondary node (2) to envelop the non-cooperative target in the formation ring belt, so that the ropes remain in a relaxed state; 4) ring belt surrounding target: after each node completes the enveloping of the non-cooperative target, the rope winding and unwinding mechanisms (5) tighten the ropes, realizing the follow-up state and the final node using the tightening of the ropes to realize the surrounding state.

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