An intelligent collaborative control method for autonomous orbiting and collision avoidance of spacecraft

Abstract

The present invention provides an intelligent collaborative control method for autonomous fly-by and collision threat avoidance of spacecraft, comprising the following steps: using a hybrid state machine to establish a hybrid model of the autonomous fly-by and collision threat avoidance control system, giving the control target corresponding to each state, and monitoring and managing the behavior state of the fly-by spacecraft based on the state machine model; designing a motion planner to give a reference motion trajectory of the fly-by spacecraft to ensure the flexibility of the conversion of each behavior state of the spacecraft; designing a reduced-order observer to estimate the unknown maneuvering parameters for the target spacecraft in view of the uncertainty of the maneuvering parameters for the controller design; designing a multi-modal motion planning and control strategy according to the control targets of different states to realize multi-modal adaptive tracking control. The present invention solves the problem that the spacecraft can effectively avoid collision threat events of space targets while realizing autonomous fly-by, thereby improving the mission execution capability and space operation safety of the spacecraft.

Description

Technical Field

[0001] The present invention belongs to the technical field of autonomous control of spacecraft, and in particular relates to an intelligent collaborative control method for autonomous orbiting and collision threat avoidance of a spacecraft. Background Art

[0002] Autonomous flyby plays an important role in supporting on-orbit services and emergency response, on-orbit control of anti-satellite weapons, space target tracking, monitoring, identification and reconnaissance, space rendezvous and docking, on-orbit refueling, and space debris removal. For on-orbit control of anti-satellite weapons, autonomous flyby is used to find a path and capture control point that is favorable for approaching the target. For space target tracking, monitoring, identification and reconnaissance, autonomous flyby and precise and stable attitude pointing control are used to achieve all-round and detailed observation and monitoring of the target spacecraft, fully obtain target information, and then accurately identify the target feature information and determine the target type and capability. For on-orbit services and emergency response, space rendezvous and docking, on-orbit refueling, and space debris removal, flyby is used to find a safe approach corridor, avoid collisions with non-cooperative targets, and ensure the safety of approach. At the same time, flyby is used to identify the target feature information and motion state, and find a suitable service and operation position. The autonomous flyby process is divided into multiple stages, including guided approach, relative position maintenance (fixed-point observation, photo imaging), and flyby.

[0003] However, with the increasing amount of space debris every year and the deployment of a large number of large constellations in orbit, the space orbit environment has become increasingly crowded and the risk of collision has increased. The autonomous flyby process may face collision threats from space debris or other spacecraft. Therefore, the spacecraft is required to have the ability to autonomously avoid collision threats during the flyby, and be able to autonomously adjust its orbit and effectively deal with possible collision threats with space debris or other spacecraft during the autonomous flyby.

[0004] The existing fly-by control method is only for the fly-by stage, the control mode is single, and the problem of avoiding space collision risk is not considered. Therefore, it is impossible to make autonomous decisions according to changes in space conditions, such as sudden changes in lighting conditions and increased collision risks, adjust behavior strategies, plan motion trajectories, and change controller modes, so that the spacecraft can adapt to changes in space conditions, autonomously complete corresponding control tasks and ensure the safety of the spacecraft. Therefore, it is necessary to increase the autonomous capability of the spacecraft control system. In order to enable the spacecraft to autonomously complete the entire fly-by process and to autonomously avoid collision risk events to ensure the execution of the mission and the safety of the spacecraft, the present invention designs an intelligent collaborative control method for autonomous fly-by and collision risk avoidance based on a hybrid state machine method. Summary of the invention

[0005] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and provide an intelligent collaborative control method for autonomous orbiting and collision threat avoidance of a spacecraft, which can achieve autonomous orbiting while avoiding collision threats, thereby ensuring mission accomplishment and the spacecraft's own safety.

[0006] The technical solution of the present invention is:

[0007] An intelligent collaborative control method for autonomous orbiting and collision threat avoidance of a spacecraft comprises the following steps:

[0008] (1) Setting the behavior states of the spacecraft in the process of autonomous flyby and collision threat avoidance, the control targets in each behavior state, setting the discrete events in the process of autonomous flyby and collision threat avoidance and the trigger conditions for the start and end of each discrete event, and setting the transition function between the behavior states according to the discrete events and their start and end trigger conditions. The value of the transition function is used to identify whether to perform the transition between the behavior states;

[0009] (2) After receiving the autonomous flyby start command, the flyby spacecraft enters the initial behavior state, determines the trigger state of each discrete event according to the relative motion parameters between the flyby spacecraft and the target spacecraft, the autonomous flyby mission requirements, and the collision threat situation, and calculates the value of the conversion function according to the trigger state of each discrete event. When the corresponding conversion function is satisfied, the behavior state is converted, otherwise the current behavior state is maintained;

[0010] (3) Under each behavior state, the flyby spacecraft performs motion planning based on the relative motion parameters with the target spacecraft and the control objectives of the current behavior state to obtain a reference trajectory for the current behavior state. The unknown maneuvering parameters of the target spacecraft are then estimated based on the current behavior state, and the control parameters are further determined by combining the reference trajectory and the estimated unknown maneuvering parameters to achieve the control objectives under the current behavior state.

[0011] Preferably, the CW equation in the reference particle orbit coordinate system is used to describe the relative motion between the orbiting spacecraft and other space targets, wherein the origin of the reference particle orbit coordinate system is the reference particle, and Z o The Y axis points from the reference mass to the center of the earth. o The axis is perpendicular to the instantaneous orbital plane of the reference particle and points in the negative normal direction of the orbital plane. o The axis direction forms a right-hand coordinate system with the other two axes. The relative motion of the orbiting spacecraft and other space targets satisfies the following expression:

[0012]

[0013]

[0014] Among them, P represents the flying spacecraft, i represents other space targets, and ρPi =[x Pi ,y Pi ,z Pi ] T Represents the relative position vector of the orbiting spacecraft and other space targets; Represents the relative velocity vector between the orbiting spacecraft and other space targets; represents the relative acceleration vector between the orbiting spacecraft and other space targets, u P 、u i Represent the control input of P and i respectively, represents the reference particle orbital angular velocity.

[0015] Preferably, in step (1), the behavior states of the spacecraft during autonomous circling and collision threat avoidance and the control targets under each behavior state are set, specifically:

[0016] Guidance approach state q0: Through guidance, the orbiting spacecraft reaches the desired orbiting position of the target spacecraft, and the control target is ρ PE =ρ PE,ref ; where E represents the target spacecraft, ρ PE represents the relative position between the flyby spacecraft and the target spacecraft, ρ PE,ref represents the expected flyby position of the flyby spacecraft relative to the target spacecraft;

[0017] Relative position holding state q1: The relative position between the orbiting spacecraft and the target spacecraft remains stable, and the target spacecraft is observed. The control target is Among them, ρ Pi,k represents the observation location, represents the relative speed between the flyby spacecraft and the target spacecraft; represents the relative acceleration between the flyby spacecraft and the target spacecraft;

[0018] Flyby transfer state q2: The flyby spacecraft transfers from one observation position to another observation position. The control target is: Among them, ρ r,ref Indicates the reference position of the fly-around transfer. represents the fly-by transfer reference speed;

[0019] Collision threat avoidance state q3: The flyby spacecraft avoids the space target with collision threat, so that the distance between the flyby spacecraft and the spacecraft with collision threat meets the safe distance. The control target is: ||ρ Pi ||≥d s ; where ρ Pi is the distance between the orbiting spacecraft and the spacecraft with collision threat, d s Indicates the safe distance between spacecraft;

[0020] Return to evacuation state q4: After the orbiting mission is completed, the orbiter returns to the predetermined target orbit. The control objectives are: Among them, ρ P , denote the position, velocity and acceleration of the flying spacecraft in the reference particle orbit coordinate system, ρ c,ref , They respectively represent the predetermined target orbital position, orbital velocity and orbital acceleration in the reference particle orbital coordinate system.

[0021] Preferably, in step (1), discrete events in the spacecraft's autonomous flyby and collision threat avoidance and trigger conditions for the start and end of each discrete event are set, specifically including:

[0022] Fly-around task e1: e1=1 indicates the start of the fly-around task, and the trigger condition is task=1; e1=0 indicates the end of the fly-around task, and the trigger condition is n≥n ref ||t≥t t,ref ||task=0, where n represents the number of laps. ref represents the expected number of circling circles, t represents the total time spent on circling, and t t,ref Indicates the expected total duration of the circling mission. Task = 1 indicates that the autonomous circling start command has been received, and task = 0 indicates that the autonomous circling end command has been received.

[0023] Guided approach e2: e2 = 1 indicates the start of guided approach, and the trigger condition is ρ PE ≠ρ PE,ref ; e2 = 0 means that the guidance is close to the end, and the trigger condition is ρ PE =ρ PE,ref ,

[0024] Relative position hold e3: e3 = 1 indicates the start of relative position hold, the trigger condition is ρ PE =ρ PE,ref ; e3 = 0 means the relative position is held and the trigger condition is Among them, t o The observation time spent at a certain observation position, t o,ref The expected observation time at a certain observation location;

[0025] Target observation e4: e4 = 1 indicates the start of target observation. The trigger condition is e4=0 means the target observation is over, and the trigger condition is

[0026] Collision threat avoidance e5: e5 = 1 indicates the start of collision threat avoidance, and the triggering condition is ||ρ Pi ||-d c≤0&d||ρ Pi || / dt<0; e5=0 indicates the end of collision threat avoidance, and the trigger condition is ||ρ Pi ||≥d s , where d c Provides collision warning distances between spacecraft;

[0027] Fly-by transfer e6: e6 = 1 indicates the start of fly-by transfer, the triggering condition is e6=0 means the fly-around transfer is finished, and the trigger condition is and Where ψ = 5π / , which means ρ PE The unallowed backlight angle between the sun vector and the sun light vector, N represents the number of observation intervals divided in a flyby cycle, s represents the vector projection of the sun light vector in the reference orbit coordinate system under the geocentric inertial system, s g represents the projection of the sunlight vector in the geocentric inertial coordinate system, represents the transfer matrix from the geocentric inertial coordinate system to the reference orbit coordinate system, where φ represents the perigee argument, represents the orbital inclination, represents the right ascension of the ascending node; ρ PE (t k-1 ) represents the relative position of the flyby spacecraft and the target spacecraft at the last observed position;

[0028] Return evacuation e7: e7 = 1 means the return evacuation starts, and the trigger condition is e1 = 0; e7 = 0 means the evacuation ends, and the trigger condition is Where ρ c,ref , They respectively represent the predetermined target orbital position, orbital velocity and orbital acceleration in the reference orbital coordinate system.

[0029] Preferably, in step (1), the transition function between the behavior states is set according to the discrete event and the trigger conditions of its start and end, specifically including:

[0030] Transformation function T0:

[0031] δ(q0=1&e1=1&e7=0&e2=1&e5=0||q0=1&e1=0&e5=0&e7=1||q2=1&e1=1&e3=1&e5=0&e6= 0&e7=0||q2=1&e1=0&e5=0&e7=1||q3=1&e1=1&e5=0&e7=0||q3=1&e1=0&e5=0&e7=1)=1;

[0032] Transformation function T1:

[0033] δ(q0=1&e5=1||q1=1&e5=1||q2=1&e5=1||q4=1&e5=1)=1;

[0034] Transformation function T2:

[0035] δ(q1=1&e1=1&e2=0&e5=0&e6=1&e7=0)=1;

[0036] Transformation function T3:

[0037] δ(q1=1&e1=1&e2=1&e5=0)=1;

[0038] Transformation function T4:

[0039] δ(q1=1&e1=0&e5=0&e7=1)=1;

[0040] Among them, δ(·) represents the conforming function, and · represents all the variables in the function.

[0041] Preferably, in step (2), the behavior state conversion is performed according to the behavior state and the conversion function, specifically:

[0042] The initial behavior state is q0. If the conversion function T0 is satisfied, the behavior state is converted to q1. If the conversion function T1 is satisfied, the behavior state is converted to q3.

[0043] When the current behavior state is q1, if the conversion function T1 is satisfied, the behavior state is converted to q3; if the conversion function T2 is satisfied, the behavior state is converted to q2; if the conversion function T3 is satisfied, the behavior state is converted to q0; if the conversion function T4 is satisfied, the behavior state is converted to q4;

[0044] When the current behavior state is q2, if the conversion function T0 is satisfied, the behavior state is converted to q1; if the conversion function T1 is satisfied, the behavior state is converted to q3;

[0045] When the current behavior state is q3, if the transition function T0 is satisfied, the behavior state is converted to q1;

[0046] When the current behavior state is q4, if the transition function T1 is satisfied, the behavior state is converted to q3.

[0047] Preferably, in step (3), the flyby spacecraft performs motion planning according to the relative motion parameters with the target spacecraft and the control target of the current behavior state, specifically:

[0048] When the current behavior state is q0 or q4, multi-style planning is adopted, and the expression is as follows:

[0049] ρ PE=b0+b1t+b2t 2 +b3t 3 +b4t 4 +b5t 5

[0050]

[0051]

[0052]

[0053]

[0054] Among them, ρ PE (0) Respectively represent the previous behavior state q l The relative position, relative velocity, and relative acceleration between the flyby spacecraft and the target spacecraft at the termination time; ρ PE (t f ), They respectively represent the relative position, relative velocity, and relative acceleration between the orbiting spacecraft and the target spacecraft at the end of the current behavior state;

[0055] When the current behavior state is q3, potential function planning is adopted, and the expression is as follows:

[0056]

[0057]

[0058]

[0059]

[0060] Among them U s represents the total potential energy function, U ia and U ir denote the gravitational potential energy function and repulsive potential energy function between the orbiting spacecraft and the i-th space target, respectively; h denotes the total number of collision threat warning space targets; η1 and η2 denote the attraction factors; η3 denotes the repulsion factor; α and β are both constants greater than zero; ρ i1 represents the relative position vector of the i-th space target, Represents the relative velocity vector of the i-th space target.

[0061] Preferably, in step (3), the unknown maneuvering parameters of the target spacecraft are estimated according to the current behavior state, specifically:

[0062]

[0063]

[0064] in, Express The estimate, represents the estimation of f, f represents the unknown maneuvering parameters of the target, represents the estimation error, σ 1k , σ 2k is a constant greater than 0, determined according to actual conditions, and k represents the behavior state.

[0065] Preferably, in step (3), the control parameters are determined by combining the motion planning results and the estimated unknown maneuvering parameters, specifically:

[0066]

[0067] When k∈{0,1,2,3},

[0068] When k = 4, and

[0069] Among them, k∈{0,1,2,…,4} represents the control mode, corresponding to the behavior state of the spacecraft, and is the reference trajectory obtained by motion planning under each control mode, δ k >0 is a normal number.

[0070] The advantages of the present invention compared with the prior art are:

[0071] (1) The present invention establishes a hybrid model of the spacecraft autonomous fly-by and collision threat avoidance control system based on a hybrid state machine, which can effectively describe the behavior change rules of the autonomous fly-by and collision threat processes;

[0072] (2) The state transition function designed by the present invention realizes the monitoring and management of the spacecraft behavior state and ensures the coordination of the transition between the spacecraft behavior states;

[0073] (3) The present invention adapts to the changes in the behavior state of the spacecraft through a multi-modal reduced-order observer and effectively estimates the uncertain parameters in the control system;

[0074] (4) The present invention can effectively track the reference trajectory and achieve the control target of the corresponding behavior state through a multi-modal adaptive control strategy, and has strong adaptability and robustness; BRIEF DESCRIPTION OF THE DRAWINGS

[0075] Figure 1 A schematic flow chart of the intelligent collaborative control method for autonomous orbiting and collision threat avoidance of a spacecraft according to the present invention;

[0076] Figure 2 A schematic diagram of a hybrid model of the autonomous fly-around and collision threat avoidance control system of the present invention;

[0077] Figure 3 Schematic diagram of the reference particle orbit coordinate system of the present invention;

[0078] Figure 4 It is a schematic diagram of the control system of the flyby spacecraft of the present invention. DETAILED DESCRIPTION

[0079] The following detailed description of the present invention will make the features and advantages of the present invention more clear and explicit.

[0080] An intelligent collaborative control method for autonomous orbiting and collision avoidance of spacecraft, such as Figure 1 As shown, the specific implementation steps are as follows:

[0081] 1) According to the analysis of the spacecraft autonomous flyby and collision threat avoidance mission, a hybrid model of the autonomous flyby and collision threat avoidance control system is established using a hybrid state machine, such as Figure 2 As shown, it is mainly divided into discrete event dynamic system model, relative orbit dynamics model, and the mutual conversion relationship between discrete events and relative orbit positions:

[0082] (1) Discrete Event Dynamic System Model

[0083] According to the circumvention process and threat avoidance process, the event set E = {e i}, i = 1, 2, ..., 7, where e1: fly-by mission, e2: guided approach, e3: relative position maintenance, e4: target observation, e5: collision threat avoidance, e6: fly-by transfer, e7: return and evacuation. i =1 or e i = 0 (i = 1, 2, ..., 7) indicates the occurrence or end of an event. All events have their trigger conditions, and the specific definitions are shown in the relevant design of the conversion function (3).

[0084] Further, according to the event set and process state, the autonomous circumvention process and threat avoidance process are divided into five behavioral states: guided approach, relative position maintenance, circumvention transfer, collision threat avoidance and return evacuation, forming a complete state machine, and establishing a finite state machine model. The state machine conversion relationship is as follows: Figure 2 As shown, Q = {q i}, i = 0, 1, 2, ..., 4 represents a finite state set, where q0 represents guided approach, q1 represents relative position maintenance, q2 represents fly-around transfer, q3 represents collision threat avoidance, and q4 represents return evacuation. T = {T i}, i = 0, 1, 2, 3, 4 represents the transition function between states, and the specific definition is shown in (3), where T0 represents the transition function from q0 to q1, q2 to q1, and q3 to q1; T1 represents the transition function from q0 to q3, q1 to q3, q2 to q3, and q4 to q3; T2 represents the transition function from q1 to q2; T3 represents the transition function from q1 to q0; T4 represents the transition function from q1 to q4.

[0085] (2) Relative orbital dynamics model

[0086] In order to conveniently describe the relative orbital motion behavior of the spacecraft, the relevant coordinate system is first constructed, such as Figure 3 As shown, F I Represents the geocentric inertial coordinate system, which is defined as: the origin is the center of the earth, the X axis points to the vernal equinox, the Z axis points to the celestial north pole, and the Y axis and the other two axes form a right-handed coordinate system. o represents the reference particle orbit coordinate system, the origin is the reference particle, Z o The Y axis points from the reference mass to the center of the Earth. o The axis is perpendicular to the instantaneous orbital plane and points to the negative normal direction of the orbital plane. The reference particle is generally selected on a circular orbit near the spacecraft, or on a circular orbit near the target spacecraft. o The direction of the axis forms a right-hand coordinate system with the other two axes.

[0087] In the reference particle orbit coordinate system, the CW equation is used to describe the relative motion between spacecraft P and spacecraft E relative to the reference particle, as well as the relative motion between other space targets and the reference particle:

[0088]

[0089]

[0090] Where i = {P, E, C}, P represents spacecraft P, E represents spacecraft E, and C represents other spacecraft. i1 =[x i ,y i ,z i ] T represents the relative position vector, ρ i2 =[v xi ,v yi ,v zi ] T represents the relative velocity vector, u i =[u xi ,u yi ,u zi ] T represents the control input of spacecraft i. Where A1 and A2 are:

[0091]

[0092] In the formula represents the reference particle orbital angular velocity.

[0093] Define ρ PE =ρ P1 -ρ E1 and represents the relative position and velocity between spacecraft P and E, and defines ρ PC =ρ P1 -ρ C1 and The relative position and velocity between the spacecraft P and other space targets are

[0094]

[0095]

[0096] Where i = {E, C}, ρ Pi =[x Pi ,y Pi ,z Pi ] T ,

[0097] While completing the autonomous flyby of spacecraft E, spacecraft P must be able to avoid collision threats between spacecraft E and other space target spacecraft C. Considering that spacecraft P has multiple behavior states during the autonomous flyby, it is necessary to establish corresponding behavior state control targets, and then design control strategies to ensure that the control targets are achieved, so that spacecraft P enters the corresponding behavior state.

[0098] Analyze the relative motion relationship under each behavior state and establish the corresponding behavior state control target as follows:

[0099] ①Guided approach state

[0100] The guided approach state refers to the guidance of the spacecraft P to reach the expected circling position of the spacecraft E, including the initial expected circling position and the expected circling position after collision threat avoidance, that is,

[0101] ρ PE =ρ PE,ref

[0102] In the formula, ρ PE,ref is the expected flyby position of spacecraft E.

[0103] ②Relative position holding state

[0104] In the relative position holding state, the spacecraft P is mainly used for state adjustment or imaging observation of the spacecraft E at a favorable position. In this state, the relative position between the spacecraft P and the spacecraft E or other space targets is required to remain stable, that is,

[0105]

[0106] Where i = {E, C}, ρ Pi,k Indicates a fixed relative position.

[0107] ③Fly-by transfer state

[0108] The purpose of flyby transfer is to achieve multi-directional observation of the target and better understand the target. It is generally a coplanar flyby, that is, spacecraft P and spacecraft E are always in the same plane. In order to ensure that spacecraft P and spacecraft E are in the same plane, Figure 3 As shown, assuming that the radius of the flyby is r r,ref , the fly-around plane is the plane where z=0, then the following coplanar fly-around trajectory can be designed.

[0109]

[0110] Solving the above formula, we can get the trajectory curve parameter ρ r,ref and The fly-around transfer process is actually the process of tracking the reference trajectory curve, that is:

[0111]

[0112] ④Collision avoidance state

[0113] In the collision avoidance state, the main task is to avoid collision threats. Suppose the safe distance between spacecraft is d s , the collision warning distance between spacecraft is d c , when the distance between them is ||ρ Pi ||≤d c , and d||ρ Pi When || / dt≤0, there is a collision threat. At this time, the spacecraft avoids the collision threat by changing its behavior state so that the relative distance satisfies

[0114] ||ρ Pi ||≥d s

[0115] ⑤Return to evacuation status

[0116] When the flyby mission is completed, the spacecraft needs to return to the predetermined target orbit, that is:

[0117]

[0118] Where ρc,ref , They respectively represent the predetermined target orbital position, orbital velocity and orbital acceleration in the reference orbital coordinate system.

[0119] (3) Establishing the conversion relationship between discrete events and relative orbital positions

[0120] Combine the event set, relative motion parameters and control objectives in (1) to design event e i ,i=1,2,…,7 trigger conditions to ensure that the spacecraft can respond to events autonomously.

[0121] e1=1 indicates that the circling mission has started, and the trigger condition is task=1, that is, the circling mission instruction has been received; e1=0 indicates that the circling mission has ended, and the trigger condition is n≥n ref ||t≥t t,ref ||task=0, target observation is completed, n represents the number of circling circles, n ref represents the expected number of circling circles, t represents the total time spent on circling, and t t,ref Indicates the total expected duration of the flyby mission.

[0122] e2=1 indicates the start of guided approach, and the trigger condition is ρ PE ≠ρ PE,ref ; e2 = 0 means the short-range approach is over, and the trigger condition is ρ PE =ρ PE,ref .

[0123] e3=1 indicates the start of relative position holding, and the trigger condition is ρ PE =ρ PE,ref ; e3 = 0 means the relative position is held and the trigger condition is t o Indicates the observation time spent on a certain observation position, t o,ref Indicates the expected observation time at a certain observation position.

[0124] e4=1 indicates the start of target observation, and the trigger condition is e4=0 means the target observation is over, and the trigger condition is

[0125] e5=1 indicates the start of collision threat avoidance, and the triggering condition is ||ρ Pi ||-d c ≤0&d||ρ Pi || / dt<0, i={E,O}; O represents the spacecraft that poses a collision threat to spacecraft P; e5=0 indicates that the collision threat avoidance is over and the relative distance to the target is safe. The trigger condition is ||ρ Pi ||≥d s .

[0126] e6=1 indicates the start of fly-by transfer, and the triggering condition is e6=0 means the fly-around transfer is finished, and the trigger condition is

[0127] and

[0128] Where ψ=5π / 6, which means ρ PE The unallowed backlight angle between the sunlight vector and the sun's light vector. N is a positive integer, designed according to the actual situation. s represents the vector projection of the sun's light vector in the reference orbit coordinate system under the geocentric inertial system. s g represents the projection of the sunlight vector in the geocentric inertial coordinate system, represents the transfer matrix from the geocentric inertial coordinate system to the orbital coordinate system; ρ PE (t k-1 ) indicates the relative position of the target last observed.

[0129] e7=1 indicates the start of return evacuation, and the trigger condition is e1=0; e7=0 indicates the end of evacuation, and the trigger condition is

[0130] According to the designed event trigger conditions, in order to coordinate the system operation, ensure the completion of the flyby mission and the safety of the spacecraft itself, the behavior state q is designed. i to q j The conversion function between i≠j, i,j=0,1,2,…,5 is:

[0131] 11) The conversion function T0 from q0 to q1, q2 to q1, and q3 to q1 is designed as:

[0132] δ(q0=1&e1=1&e7=0&e2=1&e5=0||q0=1&e1=0&e5=0&e7=1||q2=1&e1=1&e3=1&e5=0&e6= 0&e7=0||q2=1&e1=0&e5=0&e7=1||q3=1&e1=1&e5=0&e7=0||q3=1&e1=0&e5=0&e7=1)=1

[0133] 12) The conversion function T1 from q0 to q3, q1 to q3, q2 to q3, and q4 to q3 is:

[0134] δ(q0=1&e5=1||q1=1&e5=1||q2=1&e5=1||q4=1&e5=1)=1

[0135] 13) The conversion function T2 from q1 to q2 is:

[0136] δ(q1=1&e1=1&e2=0&e5=0&e6=1&e7=0)=1;

[0137] 14) The conversion function T3 from q1 to q0 is:

[0138] δ(q1=1&e1=1&e2=1&e5=0)=1;

[0139] 15) The conversion function T4 from q1 to q4 is:

[0140] δ(q1=1&e1=0&e5=0&e7=1)=1.

[0141] Through logical analysis, it can be seen that the state transition functions are mutually exclusive, that is, the state transition functions cannot be true at the same time, there is no conflict, the logic is self-consistent, and the coordination of the system can be guaranteed.

[0142] According to the descriptions of (1), (2) and (3), the hybrid model of the autonomous fly-around control system can be described by the following hybrid state machine model:

[0143] H=(Q,X,E,U,T,Y,f,Init,Inv)

[0144] Where Q = {q i}, i=0,1,2,…,4 represents a finite state set, represents a continuous variable, E={e i}, i=1,2,…,7 represents a discrete event set, U={u pi (t)} represents the segmented control strategy of the continuous dynamic system, T = {T i}, i=0,1,2,…,4 state transition function, Y,f,Init,Inv are the output variable set of the continuous dynamic system, the differential equation of the continuous dynamic system, the initial value of the system and the constraint boundary of the continuous variable respectively.

[0145] 2) According to the hybrid model established in 1), the control objectives of the spacecraft P in different behavior states are different, so the control strategies adopted are also different. In order to reduce the impact of the control system dynamic performance during the control strategy switching process, motion planning is required to make the motion curve smooth. Here, motion planning uses a combination of polynomial planning and potential function planning. Polynomial planning is used in states q0 and q4, and potential function planning is used in state q3.

[0146] ① Motion planning under the guided approach state q0 and return evacuation state q4

[0147] ρ Pi =b0+b1t+b2t 2 +b3t3 +b4t 4 +b5t 5

[0148]

[0149]

[0150] Where b0, b1, b2, b3, b4 and b5 are parameter vectors. The initial constraint is the previous state q l The relative position, velocity and acceleration information at the end time of

[0151]

[0152] The termination time constraint condition is the relative position, velocity and acceleration information corresponding to the control target of the current behavior state.

[0153] The terminal constraint condition of the guidance approaching state q0 is:

[0154] if

[0155] but

[0156] if

[0157] but

[0158] The terminal constraint for returning to the evacuation state q4 is:

[0159]

[0160] According to the initial conditions and terminal constraints, given the time length t

[0161]

[0162]

[0163] in

[0164]

[0165]

[0166] ②Motion planning for collision avoidance state q3

[0167] During the autonomous orbiting process, the spacecraft P must ensure both the completion of the mission and its own safety. Therefore, when a collision threat avoidance warning event occurs, the spacecraft needs to enter the threat avoidance state and perform replanning to ensure the avoidance of the collision threat and the safe operation of the spacecraft. In the threat avoidance state, the replanning adopts a model-based potential function planning method. The model adopts the CW equation, and the potential function is constructed as follows:

[0168]

[0169]

[0170]

[0171]

[0172]

[0173] Among them U s represents the total potential energy function, U ia and U ir They represent the gravitational potential energy function and repulsive potential energy function between the i-th target, respectively, h represents the total number of collision threat warning space targets, η1 and η2 represent attraction factors, and η3 represents repulsion factor.

[0174] When planning motion, the planning strategy to be adopted needs to be determined according to the behavior state of the spacecraft P. The fly-by transfer process q2 and the relative position holding state q1, the motion behavior are already given, so no further planning is required. After planning the reference motion trajectory, the spacecraft P can achieve the corresponding control target by tracking the established motion trajectory.

[0175] 3) Uncertain parameter estimation

[0176] Before designing the controller, considering that the maneuvering parameters of spacecraft E and other space target spacecraft O are unknown, it is necessary to estimate them and then feed them back to the controller. The relative distance and relative speed between the spacecraft E and the target can be obtained through the navigation system, and then on this basis, a reduced-order observer is designed to estimate the uncertain parameters. The form of the reduced-order observer is as follows:

[0177]

[0178]

[0179] In the formula Express The estimate, represents the estimation of f, f represents the unknown maneuvering parameters of the target, represents the estimation error, σ2k represents the parameters to be designed, σ 1k Represents the parameters to be designed, and k represents the behavior state.

[0180] 4) Adaptive tracking controller design

[0181] According to the motion planning results given by the planner and the estimation of uncertainty parameters given by the observer, an adaptive tracking control strategy is designed to achieve the control target under different behavior states. Considering that the control target and the controlled object change under different behavior states, the spacecraft controller is designed to be multimodal to adapt to the tracking of the reference trajectory under different behavior states. The entire continuous variable closed-loop control system framework is as follows: Figure 4 Assume that state q k , the reference trajectory given by the planner under k∈{0,1,…,4} is and The multi-mode adaptive tracking controller designed by the present invention is in the following form:

[0182]

[0183] Where δ k >0 is a positive constant, k∈{0,1,2,…,4} represents the controller mode, corresponding to the behavior state of the spacecraft. In the case of k∈{0,1,2,3}, and When k=4 and

[0184] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. An intelligent collaborative control method for autonomous orbiting and collision threat avoidance of a spacecraft, characterized in that: The following steps are involved: (1) Setting the behavior states of the spacecraft in the process of autonomous flyby and collision threat avoidance, the control targets in each behavior state, setting the discrete events in the process of autonomous flyby and collision threat avoidance and the trigger conditions for the start and end of each discrete event, and setting the transition function between the behavior states according to the discrete events and their start and end trigger conditions. The value of the transition function is used to identify whether to perform the transition between the behavior states; (2) After receiving the autonomous flyby start command, the flyby spacecraft enters the initial behavior state, determines the trigger state of each discrete event according to the relative motion parameters between the flyby spacecraft and the target spacecraft, the autonomous flyby mission requirements, and the collision threat situation, and calculates the value of the conversion function according to the trigger state of each discrete event. When the corresponding conversion function is satisfied, the behavior state is converted, otherwise the current behavior state is maintained; (3) In each behavior state, the flyby spacecraft performs motion planning based on the relative motion parameters with the target spacecraft and the control target of the current behavior state to obtain the reference trajectory of the current behavior state; Then, the unknown maneuvering parameters of the target spacecraft are estimated according to the current behavior state, and the control parameters are further determined by combining the reference trajectory and the estimated unknown maneuvering parameters to achieve the control target under the current behavior state; The relative motion between the flying spacecraft and other space targets is described by the CW equation in the reference particle orbit coordinate system, where the origin of the reference particle orbit coordinate system is the reference particle, Z o The Y axis points from the reference mass to the center of the Earth. o The axis is perpendicular to the instantaneous orbital plane of the reference particle and points in the negative normal direction of the orbital plane. o The axis direction forms a right-hand coordinate system with the other two axes. The relative motion of the orbiting spacecraft and other space targets satisfies the following expression: Among them, P represents the flying spacecraft, i represents other space targets, and ρ Pi =[x Pi ,y Pi ,z Pi ] T Represents the relative position vector of the orbiting spacecraft and other space targets; Represents the relative velocity vector between the orbiting spacecraft and other space targets; represents the relative acceleration vector between the orbiting spacecraft and other space targets, u P 、u i Represent the control input of P and i respectively, represents the reference particle orbital angular velocity; In step (1), the behavior states of the spacecraft during autonomous circling and collision threat avoidance and the control targets under each behavior state are set, specifically: Guidance approach state q0: Through guidance, the orbiting spacecraft reaches the desired orbiting position of the target spacecraft, and the control target is ρ PE =ρ PE,ref ; where E represents the target spacecraft, ρ PE represents the relative position between the flyby spacecraft and the target spacecraft, ρ PE,ref represents the expected flyby position of the flyby spacecraft relative to the target spacecraft; Relative position holding state q1: The relative position between the orbiting spacecraft and the target spacecraft remains stable, the target spacecraft is observed, and the control target is ρ PE =ρ PE,k , Among them, ρ Pi,k represents the observation location, represents the relative speed between the flyby spacecraft and the target spacecraft; represents the relative acceleration between the flyby spacecraft and the target spacecraft; Flyby transfer state q2: The flyby spacecraft transfers from one observation position to another observation position, and the control target is: ρ PE =ρ r,ref , Among them, ρ r,ref Indicates the reference position of the fly-around transfer. represents the fly-by transfer reference speed; Collision threat avoidance state q3: The flyby spacecraft avoids the space target with collision threat, so that the distance between the flyby spacecraft and the spacecraft with collision threat meets the safe distance. The control target is: ||ρ Pi ||≥d s ; where ρ Pi is the distance between the orbiting spacecraft and the spacecraft with collision threat, d s Indicates the safe distance between spacecraft; Return to evacuation state q4: After the orbiting mission is completed, the orbiter returns to the predetermined target orbit, and the control target is: ρ P =ρ c,ref , Among them, ρ P , denote the position, velocity and acceleration of the flying spacecraft in the reference particle orbit coordinate system, ρ c,ref , They represent the predetermined target orbital position, orbital velocity and orbital acceleration in the reference particle orbital coordinate system respectively; In the step (1), discrete events in the autonomous flyby and collision threat avoidance of the spacecraft and trigger conditions for the start and end of each discrete event are set, specifically including: Fly-around task e1: e1=1 indicates the start of the fly-around task, and the trigger condition is task=1; e1=0 indicates the end of the fly-around task, and the trigger condition is n≥n ref ||t≥t t,ref ||task=0, where n represents the number of laps. ref represents the expected number of circling circles, t represents the total time spent on circling, and t t,ref Indicates the expected total duration of the circling mission. Task = 1 indicates that the autonomous circling start command has been received, and task = 0 indicates that the autonomous circling end command has been received. Guided approach e2: e2 = 1 indicates the start of guided approach, and the trigger condition is ρ PE ≠ρ PE,ref ; e2 = 0 means that the guidance is close to the end, and the trigger condition is ρ PE =ρ PE,ref , Relative position hold e3: e3 = 1 indicates the start of relative position hold, the trigger condition is ρ PE =ρ PE,ref ; e3 = 0 means the relative position is held and the trigger condition is Among them, t o The observation time spent at a certain observation position, t o,ref The expected observation time at a certain observation location; Target observation e4: e4 = 1 indicates the start of target observation. The trigger condition is e4=0 means the target observation is over, and the trigger condition is Collision threat avoidance e5: e5 = 1 indicates the start of collision threat avoidance, and the triggering condition is ||ρ Pi ||-d c ≤0&d||ρ Pi || / dt<0; e5=0 indicates the end of collision threat avoidance, and the trigger condition is ||ρ Pi ||≥d s , where d c Provides collision warning distances between spacecraft; Fly-by transfer e6: e6 = 1 indicates the start of fly-by transfer, the triggering condition is e6=0 means the fly-around transfer is finished, and the trigger condition is and Where ψ = 5π / , which means ρ PE The unallowed backlight angle between the sun vector and the sun light vector, N represents the number of observation intervals divided in a flyby cycle, s represents the vector projection of the sun light vector in the reference orbit coordinate system under the geocentric inertial system, s g represents the projection of the sunlight vector in the geocentric inertial coordinate system, represents the transfer matrix from the geocentric inertial coordinate system to the reference orbit coordinate system, where φ represents the perigee argument, represents the orbital inclination, represents the right ascension of the ascending node; ρ PE (t k-1 ) represents the relative position of the flyby spacecraft and the target spacecraft at the last observed position; Return evacuation e7: e7 = 1 means the return evacuation starts, and the trigger condition is e1 = 0; e7 = 0 means the evacuation ends, and the trigger condition is Where ρ c,ref , They represent the predetermined target orbital position, orbital velocity and orbital acceleration in the reference orbital coordinate system respectively; In the step (1), the transition function between the behavior states is set according to the discrete event and the trigger conditions of its start and end, specifically including: Transformation function T0: δ(q0=1&e1=1&e7=0&e2=1&e5=0||q0=1&e1=0&e5=0&e7=1|| q2=1&e1=1&e3=1&e5=0&e6=0&e7=0||q2=1&e1=0&e5=0&e7=1|| q3=1&e1=1&e5=0&e7=0||q3=1&e1=0&e5=0&e7=1)=1; Transformation function T1: δ(q0=1&e5=1||q1=1&e5=1||q2=1&e5=1||q4=1&e5=1)=1; Transformation function T2: δ(q1=1&e1=1&e2=0&e5=0&e6=1&e7=0)=1; Transformation function T3: δ(q1=1&e1=1&e2=1&e5=0)=1; Transformation function T4: δ(q1=1&e1=0&e5=0&e7=1)=1; Among them, δ(·) represents the conforming function, · represents all the variables in the function; In the step (2), the behavior state conversion is performed according to the behavior state and the conversion function, specifically: The initial behavior state is q0. If the conversion function T0 is satisfied, the behavior state is converted to q1. If the conversion function T1 is satisfied, the behavior state is converted to q3. When the current behavior state is q1, if the conversion function T1 is satisfied, the behavior state is converted to q3; if the conversion function T2 is satisfied, the behavior state is converted to q2; if the conversion function T3 is satisfied, the behavior state is converted to q0; if the conversion function T4 is satisfied, the behavior state is converted to q4; When the current behavior state is q2, if the conversion function T0 is satisfied, the behavior state is converted to q1; if the conversion function T1 is satisfied, the behavior state is converted to q3; When the current behavior state is q3, if the transition function T0 is satisfied, the behavior state is converted to q1; When the current behavior state is q4, if the transition function T1 is satisfied, the behavior state is converted to q3; In step (3), the flyby spacecraft performs motion planning according to the relative motion parameters with the target spacecraft and the control target of the current behavior state, specifically: When the current behavior state is q0 or q4, multi-style planning is adopted, and the expression is as follows: <h2 style=";text-align:left;direction:ltr">ρ<h2 style=";text-align:left;direction:ltr"> PE <h2 style=";text-align:left;direction:ltr"> =b0+b1t+b2t<h2 style=";text-align:left;direction:ltr"> 2 <h2 style=";text-align:left;direction:ltr"> +b3t<h2 style=";text-align:left;direction:ltr"> 3 <h2 style=";text-align:left;direction:ltr"> +b4t<h2 style=";text-align:left;direction:ltr"> 4 <h2 style=";text-align:left;direction:ltr"> +b5t<h2 style=";text-align:left;direction:ltr"> 5 b0=ρ PE (0), Among them, ρ PE (0) Respectively represent the previous behavior state q l The relative position, relative velocity, and relative acceleration between the flyby spacecraft and the target spacecraft at the termination time; ρ PE (t f ), They respectively represent the relative position, relative velocity, and relative acceleration between the orbiting spacecraft and the target spacecraft at the end of the current behavior state; When the current behavior state is q3, potential function planning is adopted, and the expression is as follows: Among them U s represents the total potential energy function, U ia and U ir denote the gravitational potential energy function and repulsive potential energy function between the orbiting spacecraft and the i-th space target, respectively; h denotes the total number of collision threat warning space targets; η1 and η2 denote the attraction factors; η3 denotes the repulsion factor; α and β are both constants greater than zero; ρ i1 represents the relative position vector of the i-th space target, represents the relative velocity vector of the i-th space target; In step (3), the unknown maneuvering parameters of the target spacecraft are estimated according to the current behavior state, specifically: in, Represents the PE The estimate, represents the estimation of f, f represents the unknown maneuvering parameters of the target, represents the estimation error, σ 1k , σ 2k is a constant greater than 0, determined according to actual conditions, and k represents the behavior state.

2. According to claim 1, the intelligent collaborative control method for autonomous orbiting and collision threat avoidance of a spacecraft is characterized in that: In step (3), the control parameters are determined by combining the motion planning results and the estimated unknown maneuvering parameters, specifically: When k∈{0,1,2,3}, When k = 4, and Among them, k∈{0,1,2,…,4} represents the control mode, corresponding to the behavior state of the spacecraft, and is the reference trajectory obtained by motion planning under each control mode, δ k >0 is a normal number.

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