Interception guidance method and related device considering speed time variation and target maneuver

Abstract

The application discloses an interception guidance method and related device considering speed time variation and target maneuvering, relates to the technical field of missile guidance, and comprises the following steps: acquiring missile maneuvering parameters and target maneuvering parameters; establishing a three-dimensional guidance model under the condition of speed time variation and target maneuvering according to the missile maneuvering parameters and the target maneuvering parameters; constructing a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model; increasing a vertical acceleration component on the basis of the time-varying navigation ratio proportional guidance law to generate an improved augmented proportional guidance law; decomposing a missile normal acceleration vector in a missile total acceleration vector to a pitch plane and a yaw plane, and generating an attitude control instruction based on the conversion between an inertial coordinate system and a speed coordinate system; and adjusting the attitude control instruction based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameters to intercept the target. The application can effectively improve the interception precision and stability of the missile in a complex maneuvering environment.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of missile guidance, in particular to an interception guidance method considering time-varying velocity and target maneuver and a related device. BACKGROUND

[0002] In recent years, the field of guidance law design has been studying the guidance problem of maneuvering targets. Hongyan Li et al. consider the limitation of the target maneuvering ability similar to the interceptor, give the collision condition in the relative frame, then design the optimal guidance law based on the small-angle linearization method considering the upper limit of acceleration, divide the guidance process into two stages, the first stage is the convergence stage of the miss distance, the second stage is the maintenance stage of the zero-control miss distance, and analyze the upper limit of the acceleration command. Qiancheng Zhao et al. propose a pseudo-capture guidance strategy, give the guidance condition to achieve pseudo-capture, analyze the forward angle of the missile, the range of the normal acceleration and the lower bound of the guidance law parameters to achieve pseudo-capture in the traditional line-of-sight coordinate system, and give an improved idea to prevent acceleration chattering. Jianqing Li et al. study the pseudo-capture problem in the Euler-Hill system, first use the capture index to deduce the condition that can achieve target interception, then describe the target interception problem as an infinite-time domain quadratic programming problem, take the interceptor acceleration and target acceleration as disturbances, and solve the interceptor guidance law in the form of differential game.

[0003] Although these studies obtain some capture conditions from different control or guidance methods, they often do not consider the influence of time-varying velocity constraints on guidance, and the guidance assumptions are too many, the cases of no speed advantage of the interceptor or initial large lead angle error are not considered, and some guidance analysis conclusions depend on the known acceleration information of the target by the interceptor. SUMMARY

[0004] The purpose of the present application is to provide an interception guidance method considering time-varying velocity and target maneuver and a related device, which can effectively improve the interception accuracy and stability of the missile in a complex maneuvering environment.

[0005] To achieve the above purpose, the present application provides the following solutions:

[0006] In a first aspect, the present application provides an interception guidance method considering time-varying velocity and target maneuver, comprising:

[0007] obtaining missile maneuvering parameters and target maneuvering parameters; the missile maneuvering parameters include a missile total acceleration vector and a missile velocity vector; the target maneuvering parameters include a target total acceleration vector and a target velocity vector;

[0008] According to the missile maneuvering parameters and the target maneuvering parameters, a three-dimensional guidance model under the condition of time-varying speed and target maneuvering is established;

[0009] Based on the three-dimensional guidance model, a time-varying navigation ratio proportional guidance law is constructed;

[0010] On the basis of the time-varying navigation ratio proportional guidance law, a vertical acceleration component is added to generate an improved augmented proportional guidance law;

[0011] The missile normal acceleration vector in the missile total acceleration vector is decomposed into the pitch and yaw planes, and a posture control instruction is generated based on the conversion between the inertial coordinate system and the speed coordinate system;

[0012] Based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameters, the posture control instruction is adjusted to intercept the target.

[0013] Optionally, the formula expression of the three-dimensional guidance model is:

[0014]

[0015] Wherein, r is the missile-target distance, v is the relative speed between the missile and the target, σ is the lead angle, cosσ=e V ·e r , a || is the axial acceleration vector, a ⊥ is the speed normal acceleration vector, a M⊥ is the missile speed normal acceleration vector, a M|| is the missile axial acceleration vector, V M is the missile speed vector, V T is the target speed vector, a T⊥ is the target speed normal acceleration vector.

[0016] Optionally, based on the three-dimensional guidance model, a time-varying navigation ratio proportional guidance law is constructed, which specifically includes:

[0017] According to the formula The missile speed normal acceleration vector in the three-dimensional guidance model is changed by the proportional guidance law;

[0018] The target speed normal acceleration vector and the missile speed normal acceleration vector are decomposed into the relative speed system (e A ,e B ,e C ), to obtain a M⊥ =A M e A +B T e B +C M eC and a T⊥ = A t e A + B t e B + C t e C ;

[0019] determine the lead angle according to the decomposed target speed normal acceleration vector and the missile speed normal acceleration vector;

[0020] determine the miss distance z according to the lead angle; the miss distance z is used to determine the deviation between the missile direction and the target movement direction;

[0021] determine the time-varying navigation ratio proportional guidance law according to the miss distance z and the lead angle.

[0022] Optionally, the formula expression of the time-varying navigation ratio proportional guidance law is:

[0023]

[0024] wherein is a preset distance, r0 is an initial distance, the control parameter a>1, and the control parameter b>2.

[0025] Optionally, the formula expression of the improved augmented proportional guidance law is:

[0026]

[0027] wherein, is the included angle between Ω r and V M , N1 and N2 are proportional coefficients, N(t) is a time-varying navigation ratio, Ω r is a missile-target distance vector, V M is a missile speed vector, is a unit vector of the proportional guidance law, e τ is an additional acceleration unit vector.

[0028] Optionally, the formula expression for calculating the unit vector of the proportional guidance law is:

[0029]

[0030] wherein, V M is a missile speed.

[0031] Optionally, the formula expression for calculating the additional acceleration unit vector is:

[0032]

[0033] wherein, is the unit vector of missile velocity.

[0034] In a second aspect, the application provides an interception guidance device considering time-varying velocity and target maneuver, comprising:

[0035] a parameter acquisition module, configured to acquire missile maneuver parameters and target maneuver parameters; the missile maneuver parameters comprise a missile total acceleration vector and a missile velocity vector; the target maneuver parameters comprise a target total acceleration vector and a target velocity vector;

[0036] a model construction module, configured to establish a three-dimensional guidance model under time-varying velocity and target maneuver according to the missile maneuver parameters and the target maneuver parameters;

[0037] a guidance law determination module, configured to construct a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model;

[0038] a guidance law improvement module, configured to increase a vertical acceleration component on the basis of the time-varying navigation ratio proportional guidance law, to generate an improved augmented proportional guidance law;

[0039] an instruction generation module, configured to decompose a missile normal acceleration vector in the missile total acceleration vector to a pitch plane and a yaw plane, and generate an attitude control instruction based on a conversion between an inertial coordinate system and a velocity coordinate system;

[0040] an updating module, configured to adjust the attitude control instruction based on a real-time updated missile time-varying navigation ratio, a missile normal acceleration vector and target maneuver parameters, to intercept the target.

[0041] In a third aspect, the application provides a computer device, comprising a memory, a processor and a computer program stored in the memory and executable on the processor, and the processor executes the computer program to implement the interception guidance method considering time-varying velocity and target maneuver according to any one of the above.

[0042] In a fourth aspect, the application provides a computer readable storage medium, which stores a computer program, and the computer program is executed by a processor to implement the interception guidance method considering time-varying velocity and target maneuver according to any one of the above.

[0043] According to the embodiments provided in the application, the following technical effects are disclosed:

[0044] The application provides an interception guidance method and related device considering speed time variation and target maneuvering. Firstly, by acquiring missile maneuvering parameters and target maneuvering parameters, the dynamic relationship between the missile and the target can be understood in real time, and accurate basic data is provided for the subsequent guidance process. Secondly, the three-dimensional guidance model under the condition of speed time variation and target maneuvering is established based on the missile maneuvering parameters and the target maneuvering parameters, which can more truly reflect the relative motion relationship between the missile and the target in a complex maneuvering environment, and provide a more accurate model basis for the design of the guidance law. Then, the time-varying navigation ratio proportional guidance law is constructed, which can adjust the flight trajectory of the missile in real time according to the relative motion state between the missile and the target, so that the missile is closer to the target, thereby improving the interception accuracy. The vertical acceleration component is added to the improved augmented proportional guidance law, which can further improve the maneuverability and interception ability of the missile. In addition, the missile normal acceleration vector in the total acceleration vector of the missile is decomposed into the pitch and yaw planes, and the attitude control command is generated based on the conversion between the inertial coordinate system and the speed coordinate system, which can realize accurate control of the missile attitude and ensure the stable flight attitude of the missile during flight. Finally, based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameters, the attitude control command is continuously adjusted, which can realize continuous tracking and accurate interception of the target. The application can enable the missile to maintain high interception accuracy and stability in a complex maneuvering environment. BRIEF DESCRIPTION OF DRAWINGS

[0045] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the drawings needed in the embodiments will be briefly introduced as follows. Obviously, the drawings in the following description are only some embodiments of the present application, and other drawings can also be obtained by those skilled in the art without creative labor.

[0046] Figure 1 The application environment diagram of the interception guidance method considering speed time variation and target maneuvering in an embodiment of the present application;

[0047] Figure 2 The flowchart of the interception guidance method considering speed time variation and target maneuvering provided by an embodiment of the present application;

[0048] Figure 3 The three-dimensional guidance differential diagram provided by an embodiment of the present application;

[0049] Figure 4 The acceleration decomposition diagram provided by an embodiment of the present application;

[0050] Figure 5 The acceleration synthesis diagram provided by an embodiment of the present application;

[0051] Figure 6 A related diagram of a guidance angle is provided for an embodiment of the present application;

[0052] Figure 7 A three-dimensional trajectory diagram of a missile and a target is provided for an embodiment of the present application;

[0053] Figure 8 A two-dimensional trajectory diagram of a missile and a target is provided for an embodiment of the present application;

[0054] Figure 9 A miss distance curve is provided for an embodiment of the present application;

[0055] Figure 10 A lead angle curve is provided for an embodiment of the present application;

[0056] Figure 11 An acceleration component curve is provided for an embodiment of the present application;

[0057] Figure 12 A speed curve of a missile and a target is provided for an embodiment of the present application;

[0058] Figure 13 A related acceleration size curve of a missile and a target is provided for an embodiment of the present application;

[0059] Figure 14 A functional module diagram of an interception guidance device considering time-varying speed and target maneuver is provided for an embodiment of the present application;

[0060] Figure 15 A structural diagram of a computer device is provided for an embodiment of the present application. DETAILED DESCRIPTION

[0061] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative work fall within the scope of protection of the present application.

[0062] With the rapid development of aerospace technology, the design of aircraft is becoming more and more diversified, and the penetration ability is significantly improved. The new type of aircraft represented by the cruise missile not only has the normal acceleration adjustment ability of the traditional aircraft, but also has the active speed changing ability. At the same time, the development of modern guidance law makes a variety of advanced penetration strategies can be applied to hypersonic aircraft, ballistic missiles and other tactical weapons. Under this background, the rapid development of aircraft penetration technology poses a serious challenge to the existing interception system. Therefore, in-depth study of the nature of aircraft interception problem and the study of new interception guidance law have become key scientific problems to be solved in the current defense technology field, and have important theoretical value and practical significance.

[0063] The purpose of the present application is to provide an interception guidance method and related device considering speed time-varying and target maneuvering, which can effectively improve the interception accuracy and stability of the missile in a complex maneuvering environment.

[0064] In order to make the above-mentioned purposes, characteristics and advantages of the present application more obvious and easy to understand, the present application will be further described in detail below in combination with the drawings and specific embodiments.

[0065] The interception guidance method considering speed time-varying and target maneuvering provided by the embodiments of the present application can be applied to, for example Figure 1The application environment shown. Among them, the terminal 102 communicates with the server 104 through the network. The data storage system can store the data required by the server 104 to process. The data storage system can be set up separately, or integrated on the server 104, or placed on the cloud or other servers. The terminal 102 can send the missile maneuvering parameters and target maneuvering parameters to be processed to the server 104. After receiving the missile maneuvering parameters and target maneuvering parameters to be processed, the server 104 establishes a three-dimensional guidance model under the condition of speed time-varying and target maneuvering according to the missile maneuvering parameters and target maneuvering parameters. Based on the three-dimensional guidance model, a time-varying navigation ratio proportional guidance law is constructed. The vertical acceleration component is added to the time-varying navigation ratio proportional guidance law to generate an improved augmented proportional guidance law. The missile normal acceleration vector in the missile total acceleration vector is decomposed into the pitch and yaw planes, and the attitude control command is generated based on the inertial coordinate system and the velocity coordinate system conversion. Based on the real-time updated missile time-varying navigation ratio, missile normal acceleration vector and target maneuvering parameters, the attitude control command is adjusted to intercept the target. The server 104 can feed back the obtained attitude control command to the terminal 102. In addition, in some embodiments, the interception guidance method considering speed time-varying and target maneuvering can also be realized by the server 104 or the terminal 102 alone, such as the terminal 102 can directly process parameters for the missile maneuvering parameters and target maneuvering parameters to be processed, or the server 104 can obtain the missile maneuvering parameters and target maneuvering parameters to be processed from the data storage system and process parameters for the missile maneuvering parameters and target maneuvering parameters to be processed.

[0066] Among them, the terminal 102 can be, but not limited to, various desktop computers, notebook computers, smart phones, tablet computers, Internet of Things devices and portable wearable devices. The Internet of Things device can be a smart speaker, a smart television, a smart air conditioner, a smart vehicle device, etc. The portable wearable device can be a smart watch, a smart bracelet, a head-mounted device, etc. The server 104 can be realized by an independent server or a server cluster composed of multiple servers, and can also be a cloud server.

[0067] In an exemplary embodiment, as Figure 2 shown, an interception guidance method considering speed time-varying and target maneuvering is provided, which is executed by a computer device, specifically can be executed by a terminal or a server alone, or can be executed by a terminal and a server together. In the embodiment of the present application, the server 104 in the Figure 1 application environment is taken as an example for illustration, including the following steps 201 to 206. Among them:

[0068] Step 201, obtaining missile maneuvering parameters and target maneuvering parameters; the missile maneuvering parameters include missile total acceleration vector and missile speed vector; the target maneuvering parameters include target total acceleration vector and target speed vector;

[0069] Step 202, establishing a three-dimensional guidance model under the condition of speed time-varying and target maneuvering according to the missile maneuvering parameters and the target maneuvering parameters;

[0070] Step 203, constructing a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model;

[0071] Step 204, adding a vertical acceleration component to the time-varying navigation ratio proportional guidance law to generate an improved augmented proportional guidance law;

[0072] Step 205, decomposing the missile normal acceleration vector in the missile total acceleration vector into the pitch and yaw planes, and generating an attitude control instruction based on the conversion between the inertial coordinate system and the speed coordinate system;

[0073] Step 206, adjusting the attitude control instruction based on the real-time updated missile time-varying navigation ratio, missile normal acceleration vector and target maneuvering parameters to intercept the target.

[0074] In an exemplary embodiment, when steps 201-202 are performed, the following can be specifically implemented:

[0075] Obtaining missile maneuvering parameters and target maneuvering parameters;

[0076] Establishing a three-dimensional guidance model under the condition of speed time-varying and target maneuvering;

[0077] Due to the time-varying speed of the missile, the missile total acceleration vector can be expressed as M = a M|| + a M⊥ , wherein a M⊥ is the speed normal acceleration vector of the missile, and a M|| is the axial acceleration vector of the missile. Similarly, the target total acceleration vector can be expressed as T = a T|| + a T⊥ , wherein a T⊥ is the speed normal acceleration vector of the target, and a T|| is the axial acceleration vector of the target.

[0078] The relative speed vector between the missile and the target can be expressed as wherein V M is the missile speed vector, V T is the target speed vector, is the unit vector of the missile speed, V is the unit vector of target velocity M V is the magnitude of missile velocity T V is the magnitude of target velocity

[0079] The angular velocity vector can be expressed as

[0080]

[0081] According to the angular velocity vector , it can be seen that the angular velocity vector is independent of the time-varying velocity of the missile, and further , and

[0082] The acceleration vector of the relative velocity can be expressed as

[0083] Let a ⊥ = a M⊥ - a T⊥ and a || = a M|| - a T|| , where ||a || × V|| is not necessarily 0.

[0084] The angular velocity vector Ω V of V can be expressed as

[0085]

[0086] Then

[0087] Then, the angular velocity of the missile-target distance vector r can be obtained as

[0088]

[0089] Further, we can get σ is the lead angle, which satisfies cos σ = e V · e r . The controllable control of the missile is a M⊥ . Taking the derivative of cos σ = e V · e r , we get Then we get further:

[0090]

[0091] Let σ M be the angle between V M and r, and σ T be the angle between V T and r, then:

[0092]

[0093] It can be seen that the derivative of the lead angle is also irrelevant to the time-varying velocities of the missile and the target. Based on the above analysis, the three-dimensional guidance model under the time-varying velocities can be obtained as follows:

[0094]

[0095] According to the three-dimensional guidance model, the quantity that can be changed by the missile is a M⊥ , i.e., the control quantity of the interception problem. The guidance schematic diagram is shown in FIG. 2. Figure 3

[0096] In an exemplary embodiment, when step 203 is performed, the following can be specifically performed.

[0097] Step 301, the normal acceleration vector of the velocity of the missile in the three-dimensional guidance model is changed according to the proportional navigation law;

[0098] Step 302, the normal acceleration vector of the velocity of the target and the normal acceleration vector of the velocity of the missile are decomposed into the relative velocity system (e A ,e B ,e C ), to obtain a M⊥ =A M e A +B T e B +C M e C and a T⊥ =A t e A +B t e B +C t e C ;

[0099] Step 303, the lead angle is determined according to the decomposed normal acceleration vector of the velocity of the target and the normal acceleration vector of the velocity of the missile.

[0100] Step 304, the miss distance z is determined according to the lead angle; the miss distance z is used to determine the deviation between the direction of the missile and the direction of the target.

[0101] Step 305, the time-varying navigation ratio proportional navigation guidance law is determined according to the miss distance z and the lead angle.

[0102] Specifically, the basic form of the proportional navigation law is adopted by the missile, so:

[0103]

[0104] ​​Again because a ⊥ = a M⊥ . T⊥ In order to achieve the guidance purpose, it is necessary to decompose a M⊥ into relative velocity system (e T⊥ , e A , e B ). Among them, e C = e A , V , M⊥ . According to the formula :

[0105]

[0106] The decomposition of a M⊥ is shown in the schematic diagram Figure 4 .

[0107] That is, a M = A A e T + B B e M + C C e T⊥ , and similarly a t = A A e t + B B e t + C C e ⊥ .

[0108] a M⊥ = a T⊥ = (A M - A t ) e A + (B M - B t ) e B + (C M - C t ) e C .

[0109] Substituting the differential formula of the pre-angle :

[0110] Among them, C T = -a || · e C + C t .

[0111] Among them, the traditional miss distance can be expressed as z = rsinσ, in order to deal with the influence of the initial large pre-angle, the miss distance is described as z = rσ. Then:

[0112]

[0113] Let z = x + iy 2 Then:

[0114]

[0115] In order to make We have:

[0116]

[0117] On the basis of implementing the PNG, in order to make the above formula hold, the time-varying navigation ratio proportional guidance law is specifically:

[0118]

[0119] Wherein is a preset distance, satisfying r0is an initial distance, and the control parameters a > 1 and β > 2.

[0120] Wherein, in an exemplary embodiment, when step 204 is performed, it can be specifically as follows:

[0121] According to the generated time-varying navigation ratio proportional guidance law, the improvement of the time-varying proportional guidance law is realized on the basis of retaining the performance of the guidance law, and specifically:

[0122] If the lead angle tends to 0, the normal acceleration of the missile can realize the requirement of small acceleration upper limit, but when the lead angle is relatively large, the missile may need relatively large control energy to realize the convergence task of the miss distance. A vertical acceleration component of the proportional guidance law can be constructed, and then a multi-dimensional acceleration direction is added to minimize the normal acceleration of the missile in the lead angle convergence process. Let:

[0123]

[0124] Wherein, the unit vector of the proportional guidance law is And the unit vector of the additional acceleration is is the angle between Ω r and V M , and N1 and N2 are related proportional coefficients. The schematic diagram of acceleration synthesis is shown in Figure 5 .

[0125] According to the above formula, the unit vector of the proportional guidance law can be obtained as

[0126] a M⊥Decomposed into the relative velocity (e A , B , C ) coordinate system, where e A = e V , ,

[0127]

[0128] Substituting into the differential of the lead angle, we get:

[0129]

[0130] and

[0131]

[0132] Let Lyapunov function V = z 2 / 2, then:

[0133]

[0134] where,

[0135] If αN1+βN2=κ, then the convergence of the miss distance can be guaranteed.

[0136] where,

[0137] Since , From the above analysis, we get When: , M⊥ ||a

[0138]

[0139] The complete improved proportional guidance law can be expressed as:

[0140]

[0141] In this embodiment, in the implementation process of the improved proportional guidance law for the interceptor missile, the following can be specifically implemented:

[0142] (1) Select the parameters α, β related to the convergence of the miss distance, where α>1, β>2,

[0143] (2) Select the preset distance where

[0144] (3) Load r, V into the navigation device M T , and then obtain and other information.

[0145] (4) Execute the improved proportional guidance law where the time-varying navigation ratio N(t) is shown in the formula The time-varying parameters N1(t), N2(t) are solved by the formula

[0146] (5) Obtain After that, convert the three-dimensional acceleration into and then provide input for the attitude control link.

[0147] In an exemplary embodiment, when step 205 is executed, the following can be performed:

[0148] After obtaining a M⊥ , it is necessary to further decompose the normal acceleration vector into overloads in the pitch and yaw directions (i.e., a y and a z ), so as to provide expected information for the attitude control system.

[0149] In the implementation process, a M⊥ in the inertial coordinate system can be obtained first, and then a y and a z can be obtained through the conversion matrix between the inertial coordinate system and the trajectory coordinate system. In the guidance process, the related angles of the missile and the line of sight are defined as shown in Figure 6 where γ imp is the included angle between V M and V T , θ r and ψ r are the trajectory inclination angle and the trajectory deflection angle, respectively, are the trajectory inclination angle and the trajectory deflection angle of the missile, respectively. are the acceleration components in the pitch and yaw directions, respectively. Let (i I , j I , k I ) be the unit vectors of the coordinate axes in the inertial coordinate system, and V M , V T and r can be expressed as:

[0150]

[0151] where the definitions of a and a are the same as a .​

[0152] The normal acceleration vector of the missile in the inertial coordinate system can be calculated according to the formula Then, by using the conversion relationship between the inertial coordinate system and the velocity coordinate system, the following is obtained:

[0153]

[0154] Wherein

[0155] In the application, a simulation result of the interception guidance method considering time-varying velocity and target maneuvering is also provided, and is specifically as follows:

[0156] (1) The precise interception task of the maneuvering target is realized by improving the proportional guidance law and the implementation process of the guidance law.

[0157] In the simulation, the overload components of the target acceleration in the pitch and yaw directions (i.e. and ) can be set as:

[0158]

[0159] Wherein, the gravitational acceleration g = 9.81 m / s 2 . The initial position (X T0 , Y T0 , Z T0 ) of the target is (1.3 km, 1.0 km, 3.5 km), the initial velocity size V T = 200 m / s, and the initial velocity direction of the target is The initial position (X M0 , Y M0 , Z M0 ) of the missile is (0.2 km, 0.2 km, 0.6 km), and V M (t0) = 500 m / s. The initial velocity direction of the missile is set as The maximum acceleration of the missile is 15g. The missile velocity

[0160] (2) Decision of guidance parameter setting:

[0161] The related parameters in the guidance law are set as: α = 10, β = 10, The related simulation results are shown in Figures 7-13 .

[0162] (3) Result analysis:

[0163] The trajectory of the missile using the improved proportional guidance law and the target is shown in Figures 7-8 . The curve of the missile miss distance and the lead angle is shown inFigures 9-10 The normal acceleration vector generated by the missile is projected to the acceleration components in the pitch and yaw directions as shown in Figure 11 The time-varying processes of the velocities of the missile and the maneuvering target during the interception are shown in Figure 12 The time-varying processes of the velocities of the missile and the maneuvering target during the interception are shown in Figure 13 The time-varying processes of the velocities of the missile and the maneuvering target during the interception are shown in

[0164] From Figures 7-10 It can be seen that the improved guidance law can achieve convergence of the miss distance and the lead angle in the case of time-varying velocities of the target and the missile, and successfully hit the target. From Figure 11 and Figure 13 It can be seen that the improved guidance law can achieve convergence of the miss distance and the lead angle in the case of time-varying velocities of the target and the missile, and successfully hit the target. From Figure 12 It can be seen that the improved guidance law can achieve convergence of the miss distance and the lead angle in the case of time-varying velocities of the target and the missile, and successfully hit the target.

[0165] Based on the same inventive concept, the embodiments of the present application also provide a device for implementing the above-mentioned interception guidance method considering time-varying velocities and target maneuvers. The implementation scheme of the device for solving the problem is similar to the implementation scheme described in the above-mentioned method, and therefore the specific limitations in one or more device embodiments provided below can be referred to the limitations of the interception guidance method considering time-varying velocities and target maneuvers described above, which will not be described herein again.

[0166] In an exemplary embodiment, as shown in Figure 14 A device for interception guidance considering time-varying velocities and target maneuvers is provided, which comprises:

[0167] A parameter acquisition module 1401 is configured to acquire missile maneuvering parameters and target maneuvering parameters; the missile maneuvering parameters comprise a missile total acceleration vector and a missile velocity vector; the target maneuvering parameters comprise a target total acceleration vector and a target velocity vector;

[0168] A model construction module 1402 is configured to establish a three-dimensional guidance model in the case of time-varying velocities and target maneuvers according to the missile maneuvering parameters and the target maneuvering parameters;

[0169] A guidance law determination module 1403 is configured to construct a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model;

[0170] A guidance law improvement module 1404 is configured to increase a vertical acceleration component on the basis of the time-varying navigation ratio proportional guidance law, to generate an improved augmented proportional guidance law;

[0171] The instruction generation module 1405 is configured to decompose the missile normal acceleration vector in the missile total acceleration vector into the pitch plane and the yaw plane, and generate an attitude control instruction based on the inertial coordinate system and the velocity coordinate system conversion;

[0172] The updating module 1406 is configured to adjust the attitude control instruction based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameter, and intercept the target.

[0173] In an exemplary embodiment, a computer device is provided, which can be a server or a terminal. An internal structure diagram of the computer device can be as shown in FIG. 1. Figure 15 The computer device includes a processor, a memory, an input / output interface (I / O) and a communication interface. The processor, the memory and the input / output interface are connected through a system bus, and the communication interface is connected to the system bus through the input / output interface. The processor of the computer device is configured to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program and a database. The internal memory provides an environment for the operating system and the computer program in the non-volatile storage medium to run. The database of the computer device is configured to store parameter processing data. The input / output interface of the computer device is configured to exchange information between the processor and external devices. The communication interface of the computer device is configured to communicate with external terminals through network connection. The computer program is executed by the processor to implement an interception guidance method considering time-varying velocity and target maneuvering.

[0174] Those skilled in the art can understand that Figure 15 The structure shown in FIG. 1 is only a block diagram of part of the structure related to the scheme of the present application, and does not constitute a limitation on the computer device to which the scheme of the present application is applied. Specifically, the computer device can include more or fewer components than those shown in the figure, or combine certain components, or have a different component arrangement.

[0175] In an exemplary embodiment, a computer device is provided, which includes a memory and a processor. The memory stores a computer program, and the processor executes the computer program to implement the steps in the above method embodiments.

[0176] In an exemplary embodiment, a computer readable storage medium is provided, which stores a computer program. The computer program is executed by a processor to implement the steps in the above method embodiments.

[0177] It should be noted that the user information (including but not limited to user equipment information, user personal information, etc.) and data (including but not limited to data for analysis, stored data, displayed data, etc.) involved in the present application are all information and data authorized by the user or authorized by all parties, and the collection, use and processing of related data need to comply with relevant regulations.

[0178] It can be understood by those skilled in the art that all or part of the processes in the above-mentioned embodiment methods can be completed by instructing related hardware through a computer program, and the computer program can be stored in a non-volatile computer readable storage medium. When the computer program is executed, it can include the processes of the above-mentioned embodiments of each method. In the embodiments provided in the present application, any reference to memory, database or other medium can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (Read-Only Memory, ROM), magnetic tape, floppy disk, flash memory, optical storage, high-density embedded non-volatile memory, resistive memory (ReRAM), magnetoresistive random access memory (Magnetoresistive Random Access Memory, MRAM), ferroelectric memory (Ferroelectric Random Access Memory, FRAM), phase change memory (Phase Change Memory, PCM), graphene memory, etc. Volatile memory can include random access memory (Random Access Memory, RAM) or external cache memory, etc. As an illustration but not limitation, RAM can be in various forms, such as static random access memory (Static Random Access Memory, SRAM) or dynamic random access memory (Dynamic Random Access Memory, DRAM), etc.

[0179] The database involved in the embodiments provided in the present application can include at least one of a relational database and a non-relational database. The non-relational database can include a distributed database based on a blockchain, etc., without being limited thereto. The processor involved in the embodiments provided in the present application can be a general-purpose processor, a central processing unit, a graphics processing unit, a digital signal processor, a programmable logic device, a data processing logic device based on quantum computing, etc., without being limited thereto.

[0180] The technical features of the above embodiments can be combined arbitrarily. In order to make the description simple, not all possible combinations of the technical features in the above embodiments are described, but as long as the combination of the technical features does not exist contradictory, it should be considered as the scope of the present application.

[0181] The principles and implementations of the present application are described in the specific examples herein, and the above examples are only used to help understand the method of the present application and its core idea; at the same time, for those skilled in the art, according to the idea of the present application, the specific implementation and application range will be changed. In summary, the content of the specification should not be understood as a limitation of the present application.

Claims

1. An interception guidance method that takes into account the time variation of the speed and the target maneuver, characterized in that, The method comprises the following steps: acquiring missile maneuvering parameters and target maneuvering parameters; the missile maneuvering parameters comprise a missile total acceleration vector and a missile velocity vector; the target maneuvering parameters comprise a target total acceleration vector and a target velocity vector; establishing a three-dimensional guidance model under the conditions of velocity time variation and target maneuvering according to the missile maneuvering parameters and the target maneuvering parameters; constructing a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model; increasing a vertical acceleration component on the basis of the time-varying navigation ratio proportional guidance law to generate an improved augmented proportional guidance law; decomposing a missile normal acceleration vector in the missile total acceleration vector into a pitch plane and a yaw plane, and generating an attitude control instruction based on the conversion between an inertial coordinate system and a velocity coordinate system; adjusting the attitude control instruction based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameters to intercept the target; a formula expression of the three-dimensional guidance model is as follows: ； where r is the missile-target distance, v is the relative velocity between the missile and the target, is the lead angle, , is the axial acceleration vector of the missile, is the velocity normal acceleration vector of the target, is the velocity normal acceleration vector of the missile, is the axial acceleration vector of the missile, is the velocity vector of the missile, is the velocity vector of the target, is the velocity normal acceleration vector of the target; constructing a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model, specifically comprising the following steps: According to the formula The proportional navigation law is changed for the velocity normal acceleration vector of the missile in the three-dimensional guidance model. is the unit vector of the missile velocity. The velocity normal acceleration vector of the target and the velocity normal acceleration vector of the missile are decomposed into the relative velocity system wherein and ; determining a lead angle according to the decomposed target velocity normal acceleration vector and the missile velocity normal acceleration vector; determining a miss distance z according to the lead angle; the miss distance z is used to determine the deviation between the missile direction and the target motion direction; determining a time-varying navigation ratio proportional guidance law according to the miss distance z and the lead angle; a formula expression of the time-varying navigation ratio proportional guidance law is as follows: ； wherein , is a preset distance, control parameter , control parameter ; a formula expression of the improved augmented proportional guidance law is as follows: ； wherein is and the angle between and is a proportionality factor, N ( t ) is a time-varying navigation ratio, is a missile-target distance vector, is a missile velocity vector, is a unit vector of the proportional navigation law, is an additional acceleration unit vector; a formula expression for calculating a unit vector of the proportional guidance law is as follows: ； wherein is the missile velocity vector; a formula expression for calculating an additional acceleration unit vector is as follows: 。 2. An intercept guidance device for implementing the method of intercept guidance taking into account the time variation of the speed and the target maneuver according to claim 1, characterized in that, The method comprises the following steps: a parameter acquisition module is configured to acquire missile maneuvering parameters and target maneuvering parameters; the missile maneuvering parameters comprise a missile total acceleration vector and a missile velocity vector; the target maneuvering parameters comprise a target total acceleration vector and a target velocity vector; a model construction module is configured to establish a three-dimensional guidance model under the conditions of velocity time variation and target maneuvering according to the missile maneuvering parameters and the target maneuvering parameters; a guidance law determination module is configured to construct a time-varying navigation ratio proportional guidance law based on the three-dimensional guidance model; a guidance law improvement module is configured to increase a vertical acceleration component on the basis of the time-varying navigation ratio proportional guidance law to generate an improved augmented proportional guidance law; an instruction generation module is configured to decompose a missile normal acceleration vector in the missile total acceleration vector into a pitch plane and a yaw plane, and generate an attitude control instruction based on the conversion between an inertial coordinate system and a velocity coordinate system; an update module is configured to adjust the attitude control instruction based on the real-time updated missile time-varying navigation ratio, the missile normal acceleration vector and the target maneuvering parameters to intercept the target.

3. A computer device comprising: A memory, a processor and a computer program stored in the memory and executable on the processor, characterized in that the processor executes the computer program to implement the method for intercept guidance considering velocity time variation and target maneuvering according to claim 1.

4. A computer-readable storage medium having stored thereon a computer program, characterized in that, The computer program is executed by the processor to implement the method for intercept guidance considering velocity time variation and target maneuvering according to claim 1.

Images