Multi-missile space-time cooperative guidance method

Abstract

The invention discloses a multi-missile space-time cooperative guidance method. The method comprises the following steps: 1) establishing a guidance motion dynamics model of missile interception targets in a three-dimensional space; 2) calculating a guidance time consistency error through information interaction between missiles, and designing a control input aMr of the acceleration in the sight direction as a control quantity for coordinating the strike time of the missiles; and 3) designing a missile guidance law of a bottom layer, namely control input in the sight line normal direction, for controlling each missile to intercept a maneuvering target at an expected sight line angle. According to the method, time-space cooperative guidance control for intercepting the same maneuvering target by the multiple missiles in the three-dimensional space is established on the basis of a guidance time estimation method and sight angle constraints, and the relative motion relation between the missiles and the target and the terminal guidance remaining time condition of the missiles can be truly reflected.

Description

Technical Field

[0001] The invention relates to aircraft control technology, and in particular to a multi-missile space-time coordinated guidance method. Background Art

[0002] Different from traditional guidance, in the case of multi-missile coordinated guidance, each missile can communicate with each other and coordinate with each other during flight, and has a specific tactical movement law. Especially when attacking moving targets, higher requirements are placed on the performance of the coordinated guidance law. From the perspective of basic theory, through the analysis of existing literature on multi-missile coordinated guidance technology, the current research on coordinated guidance technology at home and abroad is based on the following two guidance theories:

[0003] (1) Open-loop control method based on independent guidance. The essential feature of independent collaborative guidance is that the coordination information relies only on its own state information, and the attack time or angle is set in advance for each missile to achieve a coordinated attack of multiple missiles. However, in this guidance process, there is no dynamic information exchange between the missiles, and their respective state information cannot be perceived and used by other missiles. Therefore, this is an open-loop control method.

[0004] (2) Closed-loop control method based on cooperative guidance. Compared with the first method, the coordination information of each missile incorporates not only its own state information, but also the state information of adjacent or all missiles participating in the cooperative operation. This guidance method does not require setting the guidance time in advance. Instead, during the guidance process, the missiles continuously exchange information with each other through the communication network and adjust their respective attack times according to the cooperative strategy, and finally reach a consensus on the attack time, which has the basis for autonomous coordination. However, there are modeling errors in the cooperative guidance model and the problem that external interference affects the stability of cooperative guidance. Summary of the invention

[0005] The technical problem to be solved by the present invention is to provide a multi-missile space-time coordinated guidance method in view of the defects in the prior art.

[0006] The technical solution adopted by the present invention to solve the technical problem is: a multi-missile space-time coordinated guidance method, comprising the following steps:

[0007] 1) Establish a guidance motion dynamics model of missile interception targets in three-dimensional space;

[0008]

[0009]

[0010]

[0011] Among them, r is the relative distance between the missile and the target, qε represents the line of sight inclination, q β represents the line of sight deflection angle, a T =[a Tr ,a Tε ,a Tβ ] T is the target acceleration vector projected in three directions in the line of sight coordinate system, a M =[a Mr ,a Mε ,a Mβ ] T is the missile acceleration vector projected in three directions in the line of sight coordinate system;

[0012] 2) Through information interaction between missiles, calculate the guidance time consistency error and design the control input a of the line of sight acceleration Mr Used as a control quantity to coordinate the missile's strike time;

[0013] 2.1) Establish the state equation of the missile three-dimensional coordinated terminal guidance model;

[0014]

[0015] Among them, t go is the remaining time of the missile's terminal guidance, which refers to the time required for the missile to hit the target at the current moment. It is estimated that the state variables in the state equation are Among them, q εd ,q βd They represent the desired missile-target line-of-sight inclination angle and line-of-sight deviation angle in the terminal guidance phase respectively; the subscript i represents the i-th missile;

[0016]

[0017]

[0018]

[0019] a Mq is the acceleration input in the line of sight normal direction, a Mr is the acceleration input in the line of sight direction, and represents the total interference in the line of sight, is the total disturbance caused by target acceleration and external disturbance in the line of sight direction;

[0020] 2.2) By controlling the line of sight, input a Mri Design and coordinate the remaining time of each missile to ensure that the remaining time of each missile is consistent, thereby achieving time coordination;

[0021] Let t be the current time, t goi (i=1,2,3...,n) is the remaining time of terminal guidance of the ith missile, then the predicted guidance time of the ith missile is:

[0022] t fi =t+t goi (5)

[0023] t fi Derivation gives the new state equation in the line of sight direction:

[0024]

[0025]

[0026] The state variables are q εi represents the line of sight inclination, q βi represents the line of sight deflection angle, a Tri is the target acceleration disturbance in the missile-to-target line of sight direction, a Mri is the missile acceleration input in the missile-target line of sight direction; t goi Represents the remaining guidance time of the missile;

[0027] Design the following disturbance observer to estimate the target acceleration a Tri ;

[0028]

[0029] Where:

[0030] z 1i Represents the estimated value of the missile-target relative speed

[0031] z 2i Represents the estimated value of the target acceleration in the line of sight direction

[0032] z ki Represents the estimated value of the target acceleration derivative in the line of sight direction Where, k = 3, ..., m + 1;

[0033] Parameters of the disturbance observer The selection rules are as follows:

[0034] (1)l k and The values ​​are l k = kl0-(k-1) and in is a small value;

[0035] (2) Gain value κ k and is a positive number and satisfies the polynomial y m+1 +κ1y m +…+κ m y+κ m+1 and is the Hurwitz polynomial;

[0036] (3)η is a normal number and satisfies

[0037] The convergence time range of the disturbance observer can be given by the following inequality:

[0038]

[0039] Where τ=1-l0, For any Q h >0, symmetric matrix P h Satisfies the Lyapunov equation P h B h +B h T P h =-Q h , (h=1,2), matrix B h It is given by the following formula:

[0040]

[0041] The guidance time consistency error between missiles is defined as:

[0042]

[0043] Based on the designed fixed-time disturbance observer and state equation, the control input instructions in the line of sight direction are designed as follows:

[0044]

[0045] in, is the target acceleration disturbance in the line of sight observed by the disturbance observer. Where α>0, β>0, 0<p<1, q>1.

[0046] 3) Design the underlying missile guidance law, i.e., the control input in the line of sight normal direction, which is used to control each missile to intercept the maneuvering target at the desired line of sight angle, and control the relative line of sight angle error between the missile and the target within the preset performance boundary, so that multiple missiles can complete the attack target under the action of the guidance law;

[0047] Select a performance function that converges in a fixed time, and use this function to make the system state reach the predetermined tracking performance within a fixed time. Its expression is as follows:

[0048]

[0049]

[0050] Where ρ0 is a constant, indicating the initial value of the performance function, ρ ∞ is a constant, representing the preset upper bound of the steady-state error. The decay rate of ρ(t) is the lower bound of the convergence rate of the tracking error e(t). f For ρ(t) to converge to ρ ∞ The maximum allowed convergence time of c affects the convergence speed of the performance function, and its value satisfies 0<ρ ∞ <|e i (0)|<ρ0,0<T f <∞,c>1. Therefore, for a performance function that converges in fixed time, it means that the tracking error will converge strictly within the performance envelope, and the performance function can converge to any expected tracking accuracy within a specified time.

[0051] The derivative of the performance function ρ(t) is:

[0052]

[0053] in

[0054] The error function after the design conversion is as follows:

[0055]

[0056] where k h >0,0≤δ≤1,ρ(t) is the performance function that converges in a fixed time. Because it satisfies the inequality 0<ρ ∞ <|e i (0)|<ρ0, so h(0) exists, and the derivative of the error function h(t) is:

[0057]

[0058] In order to ensure that the system tracking error meets the specified performance, it can be seen that because h(0) exists, it is only necessary to ensure that h(t) is bounded during the system convergence process. It can be seen that when h(t) is bounded, because |e(0)|<ρ(t), |e|≠ρ(t), |e|≠δρ(t), which means that the system tracking error e will not exceed the preset convergence curve. Therefore, the state constraint problem of nonlinear system tracking control can be transformed into a bounded problem of the error function h(t), and the following controller will also be designed based on this principle.

[0059] Since the guidance law in the line of sight direction is the local guidance law of each missile, all members are equal, for the convenience of writing, the subscript i is not considered, and a single missile is selected to design the acceleration command. Based on the model established above, the state variable e1=q is defined ε -q εd , e2=q β -q βd , where q εd and q βd are the expected terminal line of sight inclination angle and line of sight deflection angle respectively. Then the guidance state equation of a single missile can be obtained as:

[0060]

[0061] Combining the missile guidance model and the preset performance control, it can be seen that the guidance law based on the preset performance mainly controls the relative line of sight angle error between the missile and the target within the preset performance boundary by designing the control input in the line of sight direction, so that multiple missiles can complete the attack target under the action of the guidance law. Combined with the previous analysis, the control purpose can be expressed by the following inequality:

[0062]

[0063] The sliding surface is designed using the designed error transfer function as follows:

[0064]

[0065] Among them, k1>0, k2>0, λ1>0, λ2>0 are the parameters to be designed, h1, h2 are the error functions based on the designed state quantities e1, e2 respectively;

[0066] Taking the derivative of the sliding surface, we can get:

[0067]

[0068] Based on the sliding mode control method, the preset performance guidance law and adaptive law are designed as follows:

[0069]

[0070] u ε ,u β The equivalent control input is:

[0071]

[0072] The adaptive law is:

[0073]

[0074] Where k3>0, k4>0, μ≥1 are the selected adaptive constants. Interference upper bound The estimated value of .

[0075] The beneficial effects produced by the present invention are:

[0076] 1. Based on the guidance time estimation method and the line of sight angle constraint, the present invention establishes the time-space coordinated guidance control of multiple missiles intercepting the same maneuvering target in three-dimensional space, which can truly reflect the relative motion relationship between each missile and the target and the remaining time of the missile's terminal guidance.

[0077] 2. A fixed-time convergence interference observer is designed, which can make the remaining guidance time of multiple missiles converge to the same within a preset fixed time, ensuring that the flight control system has higher control accuracy, faster convergence speed and stronger robustness, effectively enhancing the control performance of the system. BRIEF DESCRIPTION OF THE DRAWINGS

[0078] The present invention will be further described below with reference to the accompanying drawings and embodiments, in which:

[0079] Figure 1 is a method flow chart of an embodiment of the present invention;

[0080] Figure 2 Schematic diagram of the three-dimensional strike relationship of missile targets according to an embodiment of the present invention;

[0081] Figure 3 It is a schematic diagram of the geometric relationship of the three-dimensional collaborative guidance model of an embodiment of the present invention. DETAILED DESCRIPTION

[0082] In order to make the purpose, technical solution and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention.

[0083] like Figure 1 As shown, a multi-missile space-time coordinated guidance method comprises the following steps:

[0084] 1) Establish a guidance motion dynamics model of missile interception targets in three-dimensional space;

[0085]

[0086]

[0087]

[0088] Among them, r is the relative distance between the missile and the target, qε represents the line of sight inclination, q β represents the line of sight deflection angle, a T =[a Tr ,a Tε ,a Tβ ] T is the target acceleration vector projected in three directions in the line of sight coordinate system, a M =[a Mr ,a Mε ,a Mβ ] T is the missile acceleration vector projected in three directions in the line of sight coordinate system;

[0089] Regarding the problem of missiles striking maneuvering targets in three-dimensional space, the geometric relationship between the missile and the target is as follows: Figure 2 As shown in the figure, missile represents the center of mass of the missile, target represents the center of mass of the target, Oxyz is the reference inertial coordinate system, Ox4y4z4 is the line of sight coordinate system, and the coordinate origin O is located at the center of mass of the missile.

[0090] 2) Through information interaction between missiles, calculate the guidance time consistency error and design the control input a of the line of sight acceleration Mr Used as a control quantity to coordinate the missile's strike time;

[0091] Consider the scenario where multiple missiles attack the same maneuvering target in three-dimensional space. The geometric relationship diagram is as follows: Figure 3 As shown;

[0092] Establish the state equation of the missile three-dimensional coordinated terminal guidance model;

[0093]

[0094] Among them, t go is the remaining time of the missile's terminal guidance, which refers to the time required for the missile to hit the target at the current moment. It is estimated that the state variables in the state equation are x3=[q ε -q εd ,q β -q βd ] T , Among them, q εd ,q βd They represent the desired missile-target line-of-sight inclination angle and line-of-sight deviation angle in the terminal guidance phase respectively;

[0095]

[0096]

[0097] is the acceleration input in the line of sight normal direction, a Mr is the acceleration input in the line of sight direction, and represents the total interference in the line of sight,

[0098] is the total disturbance caused by target acceleration and external disturbance in the line of sight direction;

[0099] By controlling the sight direction input a Mri Design and coordinate the remaining time of each missile to ensure that the remaining time of each missile is consistent, thereby achieving time coordination;

[0100] Let t be the current time, t goi (i=1,2,3...,n) is the remaining time of terminal guidance of the ith missile, then the predicted guidance time of the ith missile is:

[0101] t fi =t+t goi (5)

[0102] It can be seen that to achieve simultaneous interception, it is necessary to ensure the predicted guidance time t of all missiles fi Before the missile hits the target, it converges to unity. So for t fi Derivation gives the new state equation in the line of sight direction:

[0103]

[0104]

[0105] The state variables are q εi represents the line of sight inclination, q βi represents the line of sight deflection angle, a Tri is the target acceleration disturbance in the missile-to-target line of sight direction, a Mri is the missile acceleration input in the missile-target line of sight direction. goi Represents the remaining guidance time of the missile. For the above system, the following disturbance observer is designed to estimate the target acceleration a Tri ;

[0106]

[0107] Where:

[0108] z 1i Represents the estimated value of the missile-target relative speed

[0109] z 2i Represents the estimated value of the target acceleration in the line of sight direction

[0110] z ki Represents the estimated value of the target acceleration derivative in the line of sight direction k=3,…,m+1;

[0111] Parameters of the disturbance observer The selection rules are as follows:

[0112] (1)l k and The values ​​are l k = kl0-(k-1) and in is a small value;

[0113] (2) Gain value κ k and is a positive number and satisfies the polynomial y m+1 +κ1y m +…+κ m y+κ m+1 and is the Hurwitz polynomial;

[0114] (3)η is a normal number and satisfies

[0115] The convergence time range of the disturbance observer can be given by the following inequality:

[0116]

[0117] Where τ=1-l0, For any Q h >0, symmetric matrix P h Satisfies the Lyapunov equation P h B h +B h T P h =-Q h , (h=1,2), matrix B h It is given by the following formula:

[0118]

[0119] Next, in order to design the time coordination law of multiple missiles, the guidance time consistency error between missiles is defined as:

[0120]

[0121] Because t fi =t+tgoi , the guidance time consistency error can also be given by get.

[0122] Based on the designed fixed-time disturbance observer and state equation, the control input instructions in the line of sight direction are designed as follows:

[0123]

[0124] in, is the target acceleration disturbance in the line of sight observed by the disturbance observer. Where α>0, β>0, 0<p<1, q>1.

[0125] 3) Design the underlying missile guidance law, i.e., the control input in the line of sight normal direction, which is used to control each missile to intercept the maneuvering target at the desired line of sight angle;

[0126] Select a performance function that converges in a fixed time, and use this function to make the system state reach the predetermined tracking performance within a fixed time. Its expression is as follows:

[0127]

[0128]

[0129] Where ρ0 is a constant, indicating the initial value of the performance function, ρ ∞ is a constant, representing the preset upper bound of the steady-state error. The decay rate of ρ(t) is the lower bound of the convergence rate of the tracking error e(t). f For ρ(t) to converge to ρ ∞ The maximum allowed convergence time of c affects the convergence speed of the performance function, and its value satisfies 0<ρ ∞ <|e i (0)|<ρ0,0<T f <∞,c>1. Therefore, for a performance function that converges in fixed time, it means that the tracking error will converge strictly within the performance envelope, and the performance function can converge to any expected tracking accuracy within a specified time.

[0130] The derivative of the performance function ρ(t) is:

[0131]

[0132] in,

[0133] The error function after the design conversion is as follows:

[0134]

[0135] where k h>0,0≤δ≤1,ρ(t) is the performance function that converges in a fixed time. Because it satisfies the inequality 0<ρ ∞ <|e i (0)|<ρ0, so h(0) exists, and the derivative of the error function h(t) is:

[0136]

[0137] In order to ensure that the system tracking error meets the specified performance, it can be seen that because h(0) exists, it is only necessary to ensure that h(t) is bounded during the system convergence process. It can be seen that when h(t) is bounded, because |e(0)|<ρ(t), |e|≠ρ(t), |e|≠δρ(t), which means that the system tracking error e will not exceed the preset convergence curve. Therefore, the state constraint problem of nonlinear system tracking control can be transformed into a bounded problem of the error function h(t), and the following controller will also be designed based on this principle.

[0138] Since the guidance law in the line of sight direction is the local guidance law of each missile, all members are equal, for the convenience of writing, the subscript i is not considered, and a single missile is selected to design the acceleration command. Based on the model established above, the state variable e1=q is defined ε -q εd , e2=q β -q βd , where q εd and q βd are the expected terminal line of sight inclination angle and line of sight deflection angle respectively. Then the guidance state equation of a single missile can be obtained as:

[0139]

[0140] Combining the missile guidance model and the preset performance control, it can be seen that the guidance law based on the preset performance mainly controls the relative line of sight angle error between the missile and the target within the preset performance boundary by designing the control input in the line of sight direction, so that multiple missiles can complete the attack target under the action of the guidance law. Combined with the previous analysis, the control purpose can be expressed by the following inequality:

[0141]

[0142] The sliding surface is designed using the designed error transfer function as follows:

[0143]

[0144] Among them, k1>0, k2>0, λ1>0, λ2>0 are the parameters to be designed, h1, h2 are the error functions based on the designed state quantities e1, e2 respectively;

[0145] Taking the derivative of the sliding surface, we can get:

[0146]

[0147] Based on the sliding mode control method, the preset performance guidance law and adaptive law are designed as follows:

[0148]

[0149] u ε ,u β The equivalent control input is:

[0150]

[0151] The adaptive law is:

[0152]

[0153] Where k3>0, k4>0, μ≥1 are the selected adaptive constants. Interference upper bound The estimated value of .

[0154] Due to the existence of the sign function, the designed guidance law is not continuous, and there will be a chattering phenomenon caused by discontinuity, which will have an adverse effect on practical engineering applications. In order to solve this problem, the boundary layer method is usually used, that is, a continuous saturation function is used to replace the sign function. The specific form of the saturation function is as follows:

[0155]

[0156] Among them, σ>0.

[0157] It should be understood that those skilled in the art can make improvements or changes based on the above description, and all these improvements and changes should fall within the scope of protection of the appended claims of the present invention.

Claims

1. A multi-missile space-time coordinated guidance method, characterized in that: The following steps are involved: 1) Establish a guidance motion dynamics model of missile interception targets in three-dimensional space; Among them, r is the relative distance between the missile and the target, q ε represents the line of sight inclination, q β represents the line of sight deflection angle, a T =[a Tr ,a Tε ,a Tβ ] T is the target acceleration vector projected in three directions in the line of sight coordinate system, a M =[a Mr ,a Mε ,a Mβ ] T is the missile acceleration vector projected in three directions in the line of sight coordinate system; 2) Through information interaction between missiles, the guidance time consistency error is calculated, and the control input of the line of sight acceleration is designed as the control quantity to coordinate the missile's strike time; 2.1) Establish the state equation of the missile three-dimensional coordinated terminal guidance model; Among them, t go is the remaining time of the missile's terminal guidance, which refers to the time required for the missile to hit the target at the current moment. It is estimated that the state variable in the state equation is x1=r, x3=[q ε -q εd ,q β -q βd ] T , Among them, q εd ,q βd They represent the desired missile-target line-of-sight inclination angle and line-of-sight deviation angle in the terminal guidance phase respectively; the subscript i represents the i-th missile; a Mq is the acceleration input in the normal direction of the line of sight, a Mr is the acceleration input in the line of sight direction, and represents the total interference in the line of sight, is the total disturbance caused by target acceleration and external disturbance in the line of sight direction; 2.2) By controlling the sight direction input a Mri Design and coordinate the remaining time of each missile to ensure that the remaining time of each missile is consistent, thereby achieving time coordination; 3) Design the underlying missile guidance law, i.e., the control input in the line of sight normal direction, which is used to control each missile to intercept the maneuvering target at the desired line of sight angle; The design error function is as follows: Among them, the parameter k h >0,0≤δ≤1,ρ(t) is the performance function that converges in fixed time, The derivative of the error function h(t) is: Define state variable e1 = q ε -q εd , e2=q β -q βd , where q εd and q βd are the desired terminal line of sight inclination angle and line of sight deflection angle, respectively, and the guidance state equation of a single missile is obtained as follows: By designing the control input in the line of sight direction, the relative line of sight angle error between the missile and the target is controlled within the preset performance boundary, so that multiple missiles can complete the attack target under the action of the guidance law. The control objective constraint is expressed by the following inequality: The sliding surface is designed using the designed error transfer function as follows: Among them, k1>0, k2>0, λ1>0, λ2>0 are the design parameters of the sliding surface, h1, h2 are the error functions of the designed state quantities e1, e2 respectively; Taking the derivative of the sliding surface, we can get: Based on the sliding mode control method, the preset performance guidance law and adaptive law are designed as follows: u ε ,u β Equivalent control inputs for sight inclination and sight deflection are: The adaptive law is: Among them, k3>0, k4>0, μ≥1 is the selected adaptive constant, Interference upper bound The estimated value of .

2. The multi-missile space-time coordinated guidance method according to claim 1 is characterized in that: In step 2.2), the control input a in the line of sight direction is Mri Design as follows: Let t be the current time, t goi (i=1,2,3...,n) is the remaining time of terminal guidance of the ith missile, then the predicted guidance time of the ith missile is: t fi =t+t goi t fi Derivation gives the new state equation in the line of sight direction: The state variable is x 1i =r i , q εi represents the line of sight inclination, q βi represents the line of sight deflection angle, a Tri is the target acceleration disturbance in the missile-to-target line of sight direction, a Mri is the missile acceleration input in the missile-target line of sight direction; t goi Represents the remaining guidance time of the missile; Design the following disturbance observer to estimate the target acceleration a Tr i; Where: z 1i Represents the estimated value of the missile-target relative speed z 2i Represents the estimated value of the target acceleration in the line of sight direction z ki Represents the estimated value of the target acceleration derivative in the line of sight direction Where, k = 3, ..., m + 1; The guidance time consistency error between missiles is defined as: Based on the designed fixed-time disturbance observer and state equation, the control input instructions in the line of sight direction are designed as follows: you ri =a x i p +vx i q in, is the target acceleration disturbance in the line of sight direction observed by the disturbance observer, where α>0, β>0, 0<p<1, q>1.

3. The multi-missile space-time coordinated guidance method according to claim 2 is characterized in that: In step 2.2), the parameter κ of the disturbance observer k , l k , (k=1,2,…,m+1) The selection rules are as follows: (1)l k and The values ​​are l k = kl0-(k-1) and in is a small value; (2) Gain value κ k and is a positive number and satisfies the polynomial y m+1 +κ1y m +…+κ m y+κ m+1 and is the Hurwitz polynomial; (3)η is a normal number and satisfies 4. The multi-missile space-time coordinated guidance method according to claim 2 is characterized in that: In step 2.2), the convergence time range of the interference observer is as follows: Among them, τ=1-l0, For any Q h >0, symmetric matrix P h Satisfies the Lyapunov equation P h B h +B h T P h =-Q h , (h=1,2), matrix B h It is given by the following formula: The guidance time consistency error between missiles is defined as: Based on the designed fixed-time disturbance observer and state equation, the control input instructions in the line of sight direction are designed as follows: you ri =a x i p +vx i q (12) in, is the target acceleration disturbance in the line of sight observed by the disturbance observer. Where α>0, β>0, 0<p<1, q>1.

5. The multi-missile space-time coordinated guidance method according to claim 1, characterized in that: In step 3), the performance function is designed as follows: Among them, ρ0 is a constant, which represents the initial value of the performance function, ρ ∞ is a constant, which represents the preset upper bound of the steady-state error. The decay rate of ρ(t) is the lower bound of the convergence rate of the tracking error e(t). f For ρ(t) to converge to ρ ∞ The maximum allowed convergence time of c affects the convergence speed of the performance function, and its value satisfies 0<ρ ∞ <|e i (0)|<ρ0,0<T f <∞, c>1; The derivative of the performance function ρ(t) is: in, 6. The multi-missile space-time coordinated guidance method according to claim 1, characterized in that: In step 3), a continuous saturation function is used to replace the sign function in the preset performance guidance law. The specific form of the saturation function is as follows: Among them, σ>0.

7. An electronic device, characterized in that: include: one or more processors; as well as a storage device for storing one or more programs, When the one or more programs are executed by the one or more processors, the one or more processors are caused to execute the method according to any one of claims 1 to 6.

8. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method according to any one of claims 1 to 6 is implemented.

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