Multi-constraint attached guidance method for space non-cooperative maneuvering target

Abstract

The application discloses a space non-cooperative maneuvering target multi-constraint attachment guidance method, and belongs to the technical field of spacecraft guidance and control. The application realizes the method as follows: a relative dynamics model under a line-of-sight coordinate system is established, the spacecraft movement is decomposed into lateral movement considering the attachment angle constraint and longitudinal movement needing to consider the terminal position and velocity constraint through simplifying the line-of-sight relative dynamics model, and a guidance law is respectively constructed; a guidance parameter self-adaptive adjustment method is introduced, the attachment position and velocity error caused by the target maneuvering is reduced while the attachment angle constraint is considered; when the directional attachment target is difficult in an extreme case, the attachment position and velocity precision is improved through an optimization angle constraint coefficient fuse strategy, and the space non-cooperative maneuvering target multi-constraint attachment guidance is realized. The analytical guidance law constructed by the application has high calculation efficiency, can simultaneously constrain the attachment position, velocity and angle, and improves the attachment position and velocity precision on the basis of considering the attachment angle.

Description

Technical Field

[0001] This invention relates to a spacecraft guidance method, and more particularly to a multi-constraint attachment guidance method for non-cooperative maneuvering targets in space, belonging to the field of spacecraft guidance and control technology. Background Technology

[0002] With the continuous development of low-Earth orbit satellite constellation technology, the number of non-cooperative targets in space is increasing, negatively impacting the space environment and security. To manage non-cooperative targets, corresponding disposal technologies are urgently needed. Contact disposal, as an important disposal method, requires spacecraft to attach to the target, and attachment guidance technology is one of the key technologies for achieving this. Considering that disposal missions typically require attaching to targets from a predetermined direction, the angle of the target to be attached should be constrained in the guidance method, i.e., directional attachment. When the target is non-cooperative and exhibits maneuvering behavior, the guidance method also needs to possess autonomy and efficiency. Therefore, it is necessary to study high-precision, highly autonomous guidance methods capable of achieving multi-constraint attachment. Currently, research on multi-constraint guidance methods for spacecraft typically employs two methods: trajectory optimization combined with optimal trajectory tracking, and designing guidance laws with bias terms. The former has low optimization problem-solving efficiency, cannot meet the real-time requirements of guidance, and is difficult to apply in practical engineering; the latter struggles to simultaneously consider position, velocity, and angle constraints, failing to meet the guidance requirements for non-cooperative target attachment. Researching multi-constraint autonomous guidance methods with terminal position, velocity, and angle constraints is of great significance for the successful implementation of contact-based disposal missions against non-cooperative targets. Summary of the Invention

[0003] The technical problem to be solved by the multi-constraint attachment guidance method for non-cooperative maneuvering targets in space disclosed in this invention is as follows: For the attachment guidance problem of non-cooperative maneuvering targets in space, a simplified relative dynamics model is constructed in the line-of-sight coordinate system. Based on this, a multi-constraint attachment guidance law is constructed in the line-of-sight and perpendicular line-of-sight directions. An adaptive adjustment method for guidance parameters and a melting strategy for angle constraint coefficients are introduced to achieve multi-constraint attachment guidance for non-cooperative maneuvering targets in space. The analytical guidance law constructed in this invention has high computational efficiency, can simultaneously constrain attachment position, velocity, and angle, and improves the accuracy of attachment position and velocity while taking into account the attachment angle.

[0004] The objective of this invention is achieved through the following technical solution.

[0005] This invention discloses a multi-constraint attachment guidance method for non-cooperative maneuvering targets in space. By simplifying the line-of-sight relative dynamics model, the spacecraft motion is decomposed into lateral motion considering attachment angle constraints and longitudinal motion considering terminal position and velocity constraints, and guidance laws are constructed for each. An adaptive adjustment method for guidance parameters is introduced to reduce attachment position and velocity errors caused by target maneuvering while considering attachment angle constraints. In extreme cases where directional attachment to the target is difficult, an optimized angle constraint coefficient melting strategy further improves the accuracy of attachment position and velocity, thereby achieving multi-constraint attachment guidance for non-cooperative maneuvering targets in space. The analytical guidance law constructed in this invention has high computational efficiency, can simultaneously constrain attachment position, velocity, and angle, and improves the accuracy of attachment position and velocity while taking into account the attachment angle.

[0006] The present invention discloses a multi-constraint attachment guidance method for non-cooperative maneuvering targets in space, comprising the following steps:

[0007] Step 1: Establish a relative dynamic model in the line-of-sight coordinate system, and simplify the relative dynamic model by decomposing the spacecraft motion into motion perpendicular to the line-of-sight direction and motion along the line-of-sight direction, referred to as lateral motion and longitudinal motion, respectively. Construct attachment guidance laws in two directions: longitudinally satisfying terminal position and velocity constraints, and laterally satisfying attachment angle constraints.

[0008] The specific implementation method of step one is as follows:

[0009] Establish line-of-sight coordinate system O P X L Y L Z L The origin is located at the spacecraft's center of mass O. P O P X L The axis coincides with the line of sight from the spacecraft to the target; the spacecraft pointing towards the target is positive; O P Z L Vertical tracking of the axis and the target are in the same plane, and with O P X L The axis is vertical, pointing upwards is positive, O P Y L The axis is given by the right-hand rule.

[0010] At the end of the attachment mission, the spacecraft and target are in the same orbital plane, and a relative dynamic model is established in the line-of-sight coordinate system within this plane. The relative position between the target and the spacecraft is ρ, also known as the line of sight, which is the origin of the line-of-sight coordinate system. L X L The axis, the line of sight ρ, and the O coordinate system of the orbital coordinate system T Y T The angle between the axes is the line-of-sight angle q. The orbital coordinate system refers to a system where the origin is located at the target's centroid, O... T XT The axis coincides with the target position vector, and its direction is from the Earth's center to the target; O T Y T The axis is within the target orbital plane; O T Z T A non-inertial coordinate system in which the axis is perpendicular to the target orbital plane, points in the direction of the target orbital angular velocity, and satisfies the right-hand rule;

[0011] ρ0 is O L X L The unit vector of the axis, q0 is O L Y L The unit vector of the axis, k is O L Z L The unit vectors of the axes, ρ0, q0, and k together constitute the three unit vectors of the orthogonal coordinate system. The derivative of ρ with respect to time is expressed as:

[0012]

[0013] In the formula, ω represents the orbital angular velocity of the target, and ρ represents the relative distance between the target and the spacecraft.

[0014] From geometric relations we know

[0015]

[0016] In the formula, r T r P Represent the geocentric distance between the target and the spacecraft in the inertial coordinate system; μ represents the geocentric gravitational constant; F P Indicates the control quantities applied to a spacecraft; m P Indicates the mass of the spacecraft; Δf d Let r be an unknown but bounded relative perturbation acceleration. When the distance between the two spacecraft is much smaller than the distance between their centers on Earth, there exists an approximation r. T ≈r P Substituting into equation (2) yields

[0017]

[0018] Combining equation (3) and equation (1), we obtain the lateral motion equation and the longitudinal motion equation in the line-of-sight coordinate system.

[0019]

[0020]

[0021] In the formula, a q a ρ These represent lateral acceleration and longitudinal acceleration, respectively; f dq f dρLet represent the lateral perturbation acceleration and the longitudinal perturbation acceleration, respectively; Δζ and Δυ are the system's uncertain disturbance terms, expressed as follows:

[0022]

[0023] Equations (4) and (5) describe the changes in the line-of-sight angle and the relative distance, respectively. Based on equations (4) and (5), guidance laws are designed in both the lateral and longitudinal directions. The lateral guidance law is constructed based on sliding mode control theory.

[0024]

[0025] In the formula, q d Target attachment angle; K q =K q0 +K q1 K is the lateral sliding mode coefficient. q0 K is the angle constraint coefficient. q1 K represents the damping coefficient introduced to reduce system chattering. q1 >0;η q η is the lateral switching coefficient. q >0;t go Remaining flight time for the spacecraft

[0026]

[0027] In the formula, ε1 and ε2 are small positive quantities introduced to avoid the numerical result of the sliding surface being always equal to zero.

[0028] Similarly, constructing a longitudinal guidance law

[0029]

[0030] In the formula, K ρ K is the longitudinal sliding mode coefficient. ρ >1;η ρ η is the vertical switching coefficient. ρ >0.

[0031] Step 2: Reduce the guidance terminal error caused by target maneuvering through parameter adaptive adjustment method, keep the longitudinal guidance parameters constant, and design the lateral adhesion angle constraint parameter K. q0 The curve showing the change in relative distance allows the spacecraft to appropriately relax the attachment angle constraint and tilt its maneuverability towards longitudinal maneuvering in order to improve the accuracy of terminal position and velocity.

[0032] The specific implementation method for step two is as follows:

[0033] Parameter K q0 The size should vary with the relative distance between the spacecraft and the target, therefore we have

[0034] K q0 =Ψ(ρ,Δρ) (10)

[0035] In the formula, Ψ represents the parameter variation function with relative distance; ρ represents the current relative distance; Future distance represents the trend of relative distance increasing or decreasing over a period of time in the future.

[0036] In the lateral guidance law (7), K q =K q0 +K q1 Parameter K q The magnitude of K affects the chattering of the control curve. q When the value decreases, the switching coefficient η q The increased proportion of the sign function and its corresponding sign function in the calculation of maneuvering acceleration leads to increased oscillations. When K q If the value is too small, the system will oscillate significantly on both sides of the sliding mode, making it difficult to apply in actual attachment tasks. Therefore, the design parameter K is necessary. q When K is a constant, the parameter K is... q1 It is subject to change, therefore

[0037] K q1 =K q -K q0 =K q -Ψ(ρ,Δρ) (11)

[0038] Based on this, adjustment coefficients χ1 and χ2 are introduced, and a function is designed.

[0039] K q0 =Ψ(ρ,Δρ)=χ1g k (ρ)+χ2h k (Δρ) (12)

[0040] In the formula, χ1+χ2=K q And χ1<χ2; χ1<χ2 makes the function more susceptible to future trends; g k with h k These represent the curves for the ρ term and the Δρ term, respectively.

[0041] To improve sensitivity to initial changes in the relative distance between the spacecraft and the target, and to ensure the guidance law reacts as soon as unsuitable parameters emerge, the curve is designed as an upward-opening quadratic function, with the following shape expressions:

[0042]

[0043]

[0044] In the formula, ρ s , Δρs ρ represents the ideal relative distance and the ideal future distance, respectively. max , Δρ max These represent the maximum permissible relative distance and the maximum permissible future distance, respectively. Δρ s =0, ρ s ρ max , Δρ max All are positive values.

[0045] Combining the above-mentioned parameter adaptive adjustment method and step one, a guidance law is constructed to generate spacecraft guidance commands and update the spacecraft and target states. When the relative distance between the target and the spacecraft exceeds ρ... max If the distance between the spacecraft and the target continues to increase rapidly, proceed to step three. When the relative distance between the spacecraft and the target enters the preset acceptable terminal error range, guidance commands cease to be generated, and the multi-constraint attachment guidance mission for non-cooperative maneuvering targets in space is completed.

[0046] Step 3: When the relative distance between the target and the spacecraft exceeds ρ max If the error continues to increase rapidly, the situation is considered an extreme case. A circuit breaker strategy is proposed for this extreme case: when the extreme case occurs, the angle constraint coefficient is reduced to zero, and the spacecraft's maneuverability is quickly concentrated to meet the terminal position and velocity constraints. When the reset condition is met, the angle constraint coefficient is adjusted using the adjustment method in step two. By designing an adaptive adjustment curve for parameters and a circuit breaker strategy for the angle constraint coefficient, the autonomous and rational allocation of the spacecraft's maneuverability is achieved. This effectively reduces the terminal position and velocity errors caused by target maneuvering while taking into account the attachment angle constraints, thereby improving the multi-constraint attachment guidance accuracy for non-cooperative maneuvering targets in space.

[0047] The specific implementation method for step three is as follows:

[0048] When ρ and Δρ need to satisfy the condition of equation (15), K q0 =0.

[0049]

[0050] At this point, the relative distance between the spacecraft and the target is large and still tends to increase rapidly. Therefore, this situation is considered unsuitable for considering attachment angle constraints and is called an extreme case.

[0051] At this time, the lateral sliding mode system s q for

[0052]

[0053] Lateral guidance of spacecraft does not consider angle constraints, but only ensures that the spacecraft's flight direction points to the target, thereby converging the relative position and velocity in longitudinal motion. In other words, all of the spacecraft's maneuverability is used to satisfy the terminal position and velocity constraints.

[0054] When K q0 When K = 0, q0 The magnitude of the fuse no longer changes with relative distance and velocity; this is called a fuse. q0 Keep it at 0 until the reset condition is met, then recalculate K according to formula (12) in step two. q0 .

[0055] Define σ k To accept terminal error, the reset condition is:

[0056]

[0057] The circuit breaker strategy is constructed as shown in formula (18).

[0058]

[0059] In the formula, i represents the guidance cycle number.

[0060] By designing an adaptive adjustment curve for parameters and a fusion strategy as shown in formula (18), the autonomous and reasonable allocation of spacecraft maneuverability is achieved. While taking into account the attachment angle constraint, the terminal position and velocity error caused by target maneuvering is effectively reduced, and the multi-constraint attachment guidance accuracy of space non-cooperative maneuvering targets is improved.

[0061] Beneficial effects:

[0062] 1. Addressing the problem of attachment guidance for non-cooperative maneuvering targets, this invention discloses a multi-constraint attachment guidance method for space non-cooperative maneuvering targets. By simplifying the line-of-sight dynamics model, the multi-constraint guidance problem is decomposed into two problems: the design of a longitudinal guidance law considering position and velocity constraints, and the design of a lateral guidance law considering attachment angle constraints. Based on this, an analytical guidance law is designed, which can achieve the multi-constraint guidance mission objectives of position, velocity, and angle while meeting real-time requirements.

[0063] 2. The space non-cooperative maneuvering target multi-constraint attachment guidance method disclosed in this invention achieves autonomous and reasonable allocation of spacecraft maneuverability by designing an adaptive adjustment curve of parameters and a melting strategy of angle constraint coefficient. While taking into account the attachment angle constraint, it effectively reduces the terminal position and velocity errors caused by target maneuvering and improves the multi-constraint attachment guidance accuracy of space non-cooperative maneuvering targets. Attached Figure Description

[0064] Figure 1This is a flowchart of the multi-constraint attachment guidance method for non-cooperative maneuvering targets in space disclosed in this invention;

[0065] Figure 2 This is a schematic diagram of the guidance parameter variation curve;

[0066] Figure 3 The flight trajectory is attached from multiple angles in a relative coordinate system;

[0067] Figure 4 This is a magnified view of a multi-angle attached flight trajectory in a relative coordinate system.

[0068] Figure 5 This is the relative distance curve between the spacecraft and the target;

[0069] Figure 6 The relative velocity curve between the spacecraft and the target;

[0070] Figure 7 This is the line-of-sight curve of the spacecraft relative to the target. Detailed Implementation

[0071] To better illustrate the purpose and advantages of the present invention, the following description, in conjunction with an embodiment and corresponding drawings, further explains the invention.

[0072] To verify the feasibility of the method, a simulation of the spacecraft attaching to the target is performed, taking a spacecraft and a target in the equatorial plane as an example. The initial target state is r. T0 =[7378,100,0] T (Unit: km), v T0 = [-0.9962×10 -4 ,7.350,0] T (Unit: km / s), the target's acceleration is 0.5 m / s². 2 The maneuvering direction is along the positive direction of the velocity normal; the spacecraft's state is r. P0 =[7378,0,0] T (Unit: km), v P0 =[0,7.350,0] T (Unit: km / s), m P0 =100kg, the upper limit of spacecraft thrust is 100N, and the engine specific impulse is 350s.

[0073] like Figure 1 As shown in the figure, the specific implementation steps of the space non-cooperative maneuvering target multi-constraint attachment guidance method disclosed in this embodiment are as follows:

[0074] Step 1: Establish a relative dynamic model in the line-of-sight coordinate system and simplify it, decomposing the spacecraft motion into motion perpendicular to the line-of-sight direction and motion along the line-of-sight direction, referred to as lateral motion and longitudinal motion, respectively. Design attachment guidance laws in two directions: longitudinally satisfying terminal position and velocity constraints, and laterally satisfying attachment angle constraints.

[0075] The specific implementation method of step one is as follows:

[0076] First, establish the line-of-sight coordinate system O. P X L Y L Z L The origin is located at the spacecraft's center of mass O. P O P X L The axis coincides with the line of sight from the spacecraft to the target, with the line pointing towards the target spacecraft being positive; O P Z L Vertical tracking of the axis and the target are in the same plane, and with O P X L The axis is vertical, pointing upwards is positive, O P Y L The axis is given by the right-hand rule.

[0077] At the end of the attachment mission, the spacecraft and the target are in the same orbital plane, and a relative dynamic model is established in the line-of-sight coordinate system within this plane; the relative position between the target and the spacecraft is ρ, also known as the line of sight, which is the origin of the line-of-sight coordinate system. L X L Axis, O L X L O of the axis and orbital coordinate system T Y T The angle between the axes is the line-of-sight angle q, O L Y L Axis direction positive direction;

[0078] Introduce unit vectors ρ0 and q0 perpendicular to the line of sight, and the coordinate system O of the line of sight. L Z L The unit vector k along the axes together constitutes the three unit vectors of the orthogonal coordinate system. The derivative of ρ with respect to time is given.

[0079]

[0080] In the formula, ω represents the orbital angular velocity of the target, and ρ represents the relative distance between the target and the spacecraft.

[0081] From geometric relations we know

[0082]

[0083] In the formula, rT r P , representing the geocentric distance between the target and the spacecraft in the inertial coordinate system; μ = 3.986005 × 10 14 m 3 / s 2 , represents the gravitational constant; F P Indicates the control quantities applied to a spacecraft; m P Indicates the mass of the spacecraft; Δf d Let r be an unknown but bounded relative perturbation acceleration. When the distance between the two spacecraft is much smaller than the distance between their centers on Earth, there exists an approximation r. T ≈r P Substituting into equation (20) yields

[0084]

[0085] Combining equations (21) and (19), we obtain the lateral motion equations and longitudinal motion equations in the line-of-sight coordinate system.

[0086]

[0087]

[0088] In the formula, a q a ρ These represent lateral and longitudinal accelerations, respectively; f dq f dρ These represent the lateral and longitudinal perturbation accelerations, respectively; Δζ and Δυ are the system's uncertain disturbance terms.

[0089]

[0090] Equations (4) and (5) describe the changes in the line-of-sight angle and the relative distance, respectively. Based on equations (4) and (5), guidance laws are designed in both the lateral and longitudinal directions: the lateral guidance law designed based on sliding mode control theory is as follows:

[0091]

[0092] In the formula, q d For the target attachment angle, q is taken in this embodiment. d Simulations were performed using four values—-90°, 0°, 90°, and 180°—as attachment angle constraints to demonstrate that the guidance method can achieve attachment under different angle constraints. q =K q0 +K q1 Let K be the lateral sliding mode coefficient. q =30; K q0 K is the angle constraint coefficient. q1This represents the damping coefficient introduced to reduce system chattering, where K is taken as... q1 =1; η q Let η be the lateral switching coefficient. q =0.01; t go Remaining flight time for the spacecraft

[0093]

[0094] In the formula, ε1 and ε2 are small positive quantities introduced to avoid the numerical result of the sliding surface being always equal to zero.

[0095] Similarly, designing longitudinal guidance laws

[0096]

[0097] In the formula, K ρ Let K be the longitudinal sliding mode coefficient. ρ =10; η ρ Let η be the vertical switching coefficient. ρ =0.1.

[0098] Step 2: Reduce guidance terminal error caused by target maneuvering through parameter adaptive adjustment: Keep the longitudinal guidance parameters constant, and adjust the lateral attachment angle constraint parameter K. q0 The curve showing the change in relative distance allows the spacecraft to appropriately relax the attachment angle constraint and tilt its maneuverability towards longitudinal maneuvering in order to improve the accuracy of terminal position and velocity.

[0099] The specific implementation method for step two is as follows:

[0100] Parameter K q0 The size varies with the relative distance between the spacecraft and the target, thus there is

[0101] K q0 =Ψ(ρ,Δρ) (28)

[0102] In the formula, Ψ represents the parameter as a function of relative distance; The future distance represents the trend of the relative distance increasing or decreasing over a future period of time. In this embodiment, the time interval dt is selected to be numerically equal to the guidance period.

[0103] In the lateral guidance law (25), the parameter K q The magnitude of K affects the chattering of the control curve. q When the value decreases, the switching coefficient η q The increased proportion of the sign function and its corresponding sign function in the calculation of maneuvering acceleration leads to increased oscillations. When K q If the value is too small, the system will oscillate significantly on both sides of the sliding mode, making it difficult to apply in actual attachment tasks. Therefore, the design parameter K is necessary. qWhen K is a constant, the parameter K is... q1 It is subject to change, therefore there is

[0104] K q1 =K q -K q0 =K q -Ψ(ρ,Δρ) (29)

[0105] Based on this, adjustment coefficients χ1 and χ2 are introduced, and a function is designed.

[0106] K q0 =Ψ(ρ,Δρ)=χ1g k (ρ)+χ2h k (Δρ) (30)

[0107] In the formula, χ1+χ2=K q Furthermore, χ1 < χ2, making the function more susceptible to future trends; g k with h k These represent the curves for the ρ term and the Δρ term, respectively.

[0108] To improve sensitivity to initial changes in the relative distance between the spacecraft and the target, and to ensure the guidance law reacts as soon as unsuitable parameters emerge, the curve is designed as a quadratic function with an upward opening, such as... Figure 2 As shown, the mathematical expressions are as follows:

[0109]

[0110]

[0111] In the formula, ρ s , Δρ s ρ represents the ideal relative distance and the ideal future distance, respectively. max , Δρ max Let represent the maximum permissible relative distance and the maximum permissible future distance, respectively. For Δρ, when it is positive, the spacecraft begins to move away from the target, and the system state begins to diverge; therefore, we take Δρ as... s =0.

[0112] Taking the guidance period Δt = 0.5s, and combining the above-mentioned adaptive adjustment method with the guidance law described in step one, guidance commands for the spacecraft are generated in each period, the states of the spacecraft and the target are updated, and the state variables are recorded to obtain the flight trajectory. When the relative distance between the target and the spacecraft exceeds ρ... max If the distance between the spacecraft and the target continues to increase rapidly, proceed to step three. When the relative distance between the spacecraft and the target enters within 0.1m, guidance command generation stops, state variable updates and recording end, and the multi-constraint attachment guidance mission for non-cooperative maneuvering targets in space is completed.

[0113] Step 3: When the relative distance between the target and the spacecraft exceeds ρ max When the error rate continues to increase rapidly, this situation is considered an extreme case. A circuit breaker strategy is proposed for this extreme case: when the extreme case occurs, the angle constraint coefficient is reduced to zero, and the spacecraft's maneuverability is quickly concentrated on satisfying the terminal position and velocity constraints; when the reset condition is met, the angle constraint coefficient is adjusted using the adjustment method in step two. By designing an adaptive adjustment curve for parameters and a circuit breaker strategy for the angle constraint coefficient, the autonomous and rational allocation of the spacecraft's maneuverability is achieved. This effectively reduces the terminal position and velocity errors caused by target maneuvering while taking into account the attachment angle constraints, thereby improving the multi-constraint attachment guidance accuracy for non-cooperative maneuvering targets in space.

[0114] The specific implementation method for step three is as follows:

[0115] When ρ and Δρ satisfy the condition of equation (15), K q0 =0.

[0116]

[0117] At this point, the relative distance between the spacecraft and the target is large and still tends to increase rapidly. Therefore, this situation is considered unsuitable for considering attachment angle constraints and is called an extreme case.

[0118] At this time, the lateral sliding mode system s q for

[0119]

[0120] Lateral guidance of spacecraft does not consider angle constraints, but only ensures that the spacecraft's flight direction points to the target, thereby helping the relative position and velocity converge in longitudinal motion. In other words, all of the spacecraft's maneuverability is used to satisfy the terminal position and velocity constraints.

[0121] When K q0 When K = 0, q0 The magnitude of the fuse no longer changes with relative distance and velocity; this is called a fuse. q0 Keep it at 0 until the reset condition is met, then recalculate K according to formula (30) in step two. q0 .

[0122] Define σ k To determine the acceptable terminal error, σ is taken. k =0.1m, reset condition is

[0123]

[0124] Circuit Breaker Strategy

[0125]

[0126] In the formula, i represents the guidance cycle number.

[0127] The relative coordinate flight trajectories under four attachment angle constraints: -90°, 0°, 90°, and 180° are as follows: Figure 3 As shown, Figure 4 for Figure 3 A magnified view of the area within the dashed box. Taking an attachment angle constraint of 90° as an example, the relative distance curve is as follows. Figure 5 As shown, the relative velocity curve is as follows: Figure 6 As shown, the change in the line of sight angle is as follows: Figure 7 As shown, the terminal position error is 0.10m, the velocity error is 0.06m / s, and the angle error is 0.76°, satisfying the multiple constraints of the attachment. Simulation results show that the designed analytical guidance law can achieve attachment of maneuvering targets under various angle constraints, and this guidance method is suitable for multi-constraint attachment tasks of non-cooperative maneuvering targets.

[0128] The above detailed description further illustrates the purpose, technical solution, and beneficial effects of the invention. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A multi-constraint attachment guidance method for non-cooperative maneuvering targets in space, characterized in that: Includes the following steps, Step 1: Establish a relative dynamics model in the line-of-sight coordinate system and simplify the relative dynamics model by decomposing the spacecraft motion into motion perpendicular to the line-of-sight direction and motion along the line-of-sight direction, referred to as lateral motion and longitudinal motion, respectively; construct attachment guidance laws in two directions: longitudinal motion satisfies terminal position and velocity constraints, and lateral motion satisfies attachment angle constraints. Step 2: Reduce the guidance terminal error caused by target maneuvering through parameter adaptive adjustment method, keep the longitudinal guidance parameters constant, and design the lateral adhesion angle constraint parameter K. q0 The curve showing the change in relative distance allows the spacecraft to appropriately relax the attachment angle constraint and tilt its maneuverability towards longitudinal maneuvering in order to improve the accuracy of terminal position and velocity. Step 3: When the relative distance between the target and the spacecraft exceeds ρ max When the angle constraint coefficient continues to increase rapidly, this situation is considered an extreme case. A circuit breaker strategy is proposed for extreme cases: when an extreme case occurs, the angle constraint coefficient is reduced to zero, and the spacecraft's maneuverability is quickly concentrated to meet the terminal position and velocity constraints; when the reset condition is met, the angle constraint coefficient is adjusted using the adjustment method in step two; by designing the adaptive adjustment curve of the parameters and the circuit breaker strategy of the angle constraint coefficient, the autonomous and reasonable allocation of the spacecraft's maneuverability is realized, effectively reducing the terminal position and velocity errors caused by target maneuvering while taking into account the attachment angle constraints, and improving the multi-constraint attachment guidance accuracy of non-cooperative maneuvering targets in space. The specific implementation method of step three is as follows: When ρ and Δρ need to satisfy the equation Under certain conditions, K q0 =0; K q0 The angle constraint coefficient is ρ, which represents the current relative distance between the target and the spacecraft. , representing the future distance, indicating the trend of relative distance increasing or decreasing over a future period of time; ρ max , Δρ max These represent the maximum permissible relative distance and the maximum permissible future distance, respectively. In this case, the relative distance between the spacecraft and the target is large and still has a rapid increasing trend. Therefore, this situation is judged as not suitable for considering the attachment angle constraint and is called the extreme case. At this time, the lateral sliding mode system s q for q is the line-of-sight angle; the spacecraft’s lateral guidance does not consider angle constraints, but only ensures that the spacecraft’s flight direction points to the target, so that the relative position and velocity in the longitudinal motion converge, that is, all the spacecraft’s maneuverability is used to satisfy the terminal position and velocity constraints. When K q0 When K = 0 q0 The magnitude of the fuse no longer changes with relative distance and velocity; this is called a fuse. q0 Keep it at 0 until the reset condition is met, then recalculate K according to step two. q0 ; Define σ k To accept terminal error, the reset condition is: t go The remaining flight time of the spacecraft; The circuit breaker strategy is constructed as shown in formula (18). In the formula, i represents the guidance cycle number, and Ψ represents the parameter variation function with relative distance; By designing an adaptive adjustment curve for parameters and a fusion strategy as shown in formula (18), the autonomous and reasonable allocation of spacecraft maneuverability is achieved. While taking into account the attachment angle constraint, the terminal position and velocity error caused by target maneuvering is effectively reduced, and the multi-constraint attachment guidance accuracy of space non-cooperative maneuvering targets is improved.

2. The multi-constraint attachment guidance method for non-cooperative maneuvering targets in space as described in claim 1, characterized in that: The specific implementation method of step one is as follows: Establish line-of-sight coordinate system O P X L Y L Z L The origin is located at the spacecraft's center of mass O. P O P X L The axis coincides with the line of sight from the spacecraft to the target; The target aircraft is positive; O P Z L Vertical tracking of the axis and the target are in the same plane, and with O P X L The axis is vertical, pointing upwards is positive, O P Y L The axis is given by the right-hand rule; At the end of the attachment mission, the spacecraft and the target are in the same orbital plane, and a relative dynamic model is established in the line-of-sight coordinate system within this plane; the relative position between the target and the spacecraft is ρ, also known as the line of sight, which is the origin of the line-of-sight coordinate system. L X L The axis, the line of sight ρ, and the O coordinate system of the orbital coordinate system T Y T The angle between the axes is the line-of-sight angle q; the orbital coordinate system refers to a system where the origin is located at the target's centroid, O T X T The axis coincides with the target position vector, and its direction is from the Earth's center to the target; O T Y T The axis is within the target orbital plane; O T Z T A non-inertial coordinate system in which the axis is perpendicular to the target orbital plane, points in the direction of the target orbital angular velocity, and satisfies the right-hand rule; ρ0 is O L X L The unit vector of the axis, q0 is O L Y L The unit vector of the axis, k is O L Z L The unit vectors of the axes, ρ0, q0, and k together constitute the three unit vectors of the orthogonal coordinate system; the derivative of ρ with respect to time is expressed as: In the formula, ω represents the orbital angular velocity of the target; From geometric relations we know In the formula, , Represent the geocentric distance between the target and the spacecraft in the inertial coordinate system; μ represents the geocentric gravitational constant; F P This indicates the control quantities applied to a spacecraft; Indicates the mass of the spacecraft; The relative perturbation acceleration is unknown but bounded; when the distance between the two spacecraft is much smaller than the distance between the Earth's centers, there exists an approximate r. T ≈ r P Substitution get Combined With formula Obtain the lateral motion equations and longitudinal motion equations in the line-of-sight coordinate system. In the formula, , These represent lateral acceleration and longitudinal acceleration, respectively; f dq f dρ These represent lateral perturbation acceleration and longitudinal perturbation acceleration, respectively. and The uncertain interference term of the system is represented as follows: Equations (4) and (5) describe the changes in the line-of-sight angle and the relative distance, respectively. Based on equations (4) and (5), guidance laws are designed in both the lateral and longitudinal directions. A lateral guidance law is constructed based on sliding mode control theory. In the formula, q d Target attachment angle; K q =K q0 +K q1 K is the lateral sliding mode coefficient. q1 K represents the damping coefficient introduced to reduce system chattering. q1 > 0; η q η is the lateral switching coefficient. q > 0; In the formula, ε1 and ε2 are small positive quantities introduced to avoid the numerical result of the sliding surface being always equal to zero; Similarly, constructing a longitudinal guidance law In the formula, K ρ K is the longitudinal sliding mode coefficient. ρ > 1; η ρ η is the vertical switching coefficient. ρ > 0.

3. The multi-constraint attachment guidance method for non-cooperative maneuvering targets in space as described in claim 2, characterized in that: The specific implementation method for step two is as follows: Parameter K q0 The size should vary with the relative distance between the spacecraft and the target, therefore we have In the formula, Ψ represents the parameter as a function of relative distance; Lateral guidance law In the middle, K q =K q0 +K q1 Parameter K q The magnitude of K affects the chattering of the control curve; when K q When the value decreases, the switching coefficient η q The increased proportion of K and its corresponding sign function in the calculation of maneuver acceleration leads to increased oscillations; when K q If the value is too small, the system will oscillate significantly on both sides of the sliding mode, making it difficult to apply in actual attachment tasks; therefore, the design parameter K is necessary. q When K is a constant, the parameter K is... q1 It is subject to change, therefore Based on this, adjustment coefficients χ1 and χ2 are introduced, and a function is designed. In the formula, χ1 + χ2 = K q Furthermore, χ1 < χ2; χ1 < χ2 makes the function more susceptible to future trends; g k with h k Represent the curves for the ρ term and the Δρ term, respectively; The design curve is a quadratic function with an upward opening, and its shape is expressed as follows: In the formula, ρ s , Δρ s Δρ represents the ideal relative distance and the ideal future distance, respectively. s =0, ρ s ρ max , Δρ max All are positive values; Combining the above-mentioned parameter adaptive adjustment method and step one, a guidance law is constructed, spacecraft guidance commands are generated, and the spacecraft and target states are updated; When the relative distance between the target and the spacecraft exceeds ρ max If it still shows a rapid increasing trend, proceed to step three; When the relative distance between the spacecraft and the target enters the preset acceptable terminal error range, the guidance command stops being generated, and the multi-constraint attachment guidance mission for non-cooperative maneuvering targets in space is completed.

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