36 results about "Terminal guidance" patented technology

The invention discloses a high-speed long-flight AI adaptive unmanned aerial vehicle guidance optimization method which comprises the following steps that multi-source sensing data of a target area are collected in real time through a multi-mode sensor, and the multi-mode sensor comprises a thermal imaging module, a visible light camera module, an inertial measurement unit and a global positioning system; performing dynamic target identification and tracking on the multi-source sensing data acquired in the step S100 based on a deep neural network; generating an optimal attack path in real time through a reinforcement learning algorithm according to a motion track prediction result of the target in combination with flight state data of the unmanned aerial vehicle and environmental constraint conditions; and when the target enters the preset strike range, triggering an AI terminal guidance function. The method has the following advantages and effects: the autonomous combat ability of the unmanned aerial vehicle in a complex dynamic environment is greatly improved, real-time target identification, path planning and attack optimization of the unmanned aerial vehicle in a complex battlefield environment are realized, and the combat efficiency and precision are improved.

The invention discloses a high-speed maneuvering target interception collaborative surrounding guidance method based on a coverage theory. The method can ensure that the high-speed maneuvering target cannot escape after being surrounded by the intercepted projectile group, and the success rate of multi-projectile cooperative interception of the maneuvering target is greatly improved. According to the method, the maneuvering characteristics of the missile and the target are considered, the surrounding area is analyzed in combination with the concept of the Aboronius circle, and then the surrounding formation is designed to achieve redundant coverage of the target escape area. And then under the angle constraint, the time constraint and the limitation of missile acceleration of the missile encircling formation, a cooperative guidance optimization problem is constructed and solved, an optimal multi-missile guidance instruction is obtained, all missiles can reach the preset encircling formation at the same time in the flight stage, encircling of the target is completed, and the aim that the multi-missile multi-missile guidance instruction is obtained is achieved. Then, the maneuvering target is intercepted and hit in the terminal guidance stage, it is ensured that the target cannot escape after being surrounded by the interception bullet group, and cooperative surrounding, interception and hit of the high-speed maneuvering target are achieved.

The invention discloses a speed-constrained optimal guidance method for convergence of average maneuvering targets, and in the method, a speed-ratio-constrained guidance control method is provided in the tangential direction of the speed of an unmanned aerial vehicle according to an optimal control principle for terminal guidance requirements under average overload constraints. The speed of the interceptor is greater than the target speed within a certain time, so that the speed advantage of the interceptor in the terminal guidance section is ensured, and necessary conditions are provided for realizing accurate interception under the condition of average maneuvering capability; and meanwhile, the method minimizes performance indexes related to control consumption, so that the aircraft has relatively high control efficiency.

The present invention discloses a method for reconstructing the coordinated flight time of mid-terminal guidance. Using the method, after multiple hypersonic glide missiles encounter an interception penetration during the glide phase, the method can reconstruct and coordinate the flight times of the glide phase and the terminal guidance phase by changing the speed of the mid-terminal handover point, based on the premise that the parameters of the mid-terminal handover point remain unchanged, to achieve a coordinated attack on the target. To address the problem of still being able to fly in coordination during the glide phase when encountering a penetration, the method simultaneously adjusts the flight times of the mid-terminal and terminal guidance phases by changing the speed of the mid-terminal handover point, based on the premise that the parameter information of the mid-terminal handover point is the same as that during the original coordinated flight. Given the total flight time proposed by the present invention, the speed of the mid-terminal handover point corresponding to the time is calculated, and then the corresponding flight time distribution of the glide phase and the terminal guidance phase is obtained, ultimately achieving coordinated flight of multiple missiles and simultaneous target impact.

The invention discloses a fixed-wing unmanned aerial vehicle given target attack control method, and belongs to the technical field of fixed-wing unmanned aerial vehicles, and the method comprises the steps: a ground station adds the precise coordinates of a plurality of possible attack target points into a waypoint, and inputs the waypoint into a flight control system; when the unmanned aerial vehicle enters a cruising stage and flies along a predetermined route, the ground station sends an attack instruction to the flight control system; and the flight control system executes a target attack strategy after receiving the attack instruction. The method is not limited by space and time, and the fixed-wing unmanned aerial vehicle can meet terminal guidance control conditions, so that target attacks are smoothly completed.

The application discloses a mid-terminal stage handover guiding method based on laser semi-active strapdown guidance, relates to the field of aircraft control and comprises the following steps: S1, according to the pitch misalignment angle and the azimuth misalignment angle output by a seeker, the target is quickly guided into a linear area of a field of view of the seeker; S2, a missile-target line-of-sight angle in an inertial space is calculated, and an initial value of the initial missile-target line-of-sight angle is recorded, the line-of-sight angle comprises a high-low angle and an azimuth angle; S3, a digital filter algorithm is used to filter the missile-target line-of-sight angle, a missile-target line-of-sight angle rate is obtained, and the filter initial value is set as the initial value of the missile-target line-of-sight angle, so that the filter convergence is accelerated; S4, a guidance instruction is generated; and S5, a linear transition method is used to complete smooth transition of a mid-terminal stage, a L1 guidance stage and a PN guidance stage. The method guarantees smooth transition of the guidance instruction of the missile in the mid-terminal stage, provides guarantee for accurate attack on the target in the terminal guidance stage, ensures that the missile accurately hits the target, and the design of the mid-terminal stage handover guiding algorithm based on the laser semi-active strapdown guidance is crucial.

The present application relates to the field of guidance method, disclose a plurality of missile specified time cooperative interception terminal guidance method, including the following steps, the guidance model of intercepting maneuvering target is constructed, the expression of the estimated remaining flight time is derived according to the guidance model of intercepting maneuvering target, the specified attack time error is defined according to the estimated remaining flight time expression, the expression of the first derivative of the estimated remaining flight time is derived according to the specified attack time error, the specified attack time error is derived and estimated by using fixed time neural network disturbance observer, the hypersonic target specified time interception guidance law is designed, the specified time interception of the hypersonic target is realized based on the proposed fixed time neural network disturbance observer and the specified time interception guidance law, the method of the present application effectively solves the problem of the calculation efficiency of the guidance law design of the specified time cooperative interception of the hypersonic target by multiple interception missiles.

Aiming at the problem of missile terminal guidance with no-fly zone and impact angle constraints, a model predictive static programming guidance law based on control barrier function is proposed. The proposed algorithm inherits the high efficiency of the traditional model predictive static programming while ensuring the no-fly zone constraints. The main reasons are as follows: first, the projection operator under the linear inequality constraints has an explicit expression, avoiding additional optimization solving; second, due to the forward invariance of the control barrier function, the no-fly zone constraints do not need to be imposed at all discrete nodes, but only at one or a few discrete nodes, reducing the scale of solving. The simulation results show that the proposed algorithm can meet the no-fly zone and impact angle constraints while ensuring high computational efficiency. Subsequent research will focus on trajectory optimization and guidance problems with multiple no-fly zone constraints, even dynamic no-fly zone constraints.

A kind of carrier platform multi-constraint fire control method based on shift matching, which can be applied to infrared image terminal guidance type precision guided weapon in complex battlefield environment. The launch azimuth angle, launch height and launch speed data information of guided weapon are obtained through the flight carrier platform as the input of fire control algorithm design. The shift matching algorithm is used to perform shift matching operation on the launch angle, launch height and launch speed at the launch time in the fire control software. The fire control solution under the multi-constraint condition of infrared imaging type precision guided weapon is obtained. Good operational effect of aerial guided weapon can be achieved in complex battlefield environment. The release restriction under the operational background of aerial guided weapon is fully considered.

The application relates to a multi-to-multi task allocation method based on minimum zero-control miss distance. The method comprises the following steps: establishing a three-dimensional missile-target relative motion equation, deducing an analytical expression of the zero-control miss distance and a dynamic equation; calculating the zero-control miss distance value between each interceptor missile and a target according to the initial condition of the terminal guidance and the analytical expression of the zero-control miss distance; constructing an intermediate matrix according to a pre-set interception task, filling a cost matrix of a Hungarian algorithm into the intermediate matrix as matrix elements to obtain a unified cost matrix; substituting the zero-control miss distance value into the unified cost matrix to generate a multi-to-multi task allocation scheme based on the Hungarian algorithm; and designing an improved real true proportional guidance law for controlling the zero-control miss distance, and adopting the improved real true proportional guidance law to realize the multi-to-multi task allocation scheme. The method can effectively eliminate the zero-control miss distance and increase the interception success rate.

The invention relates to the technical field of guided projectiles, and discloses a data calculation method, device and equipment of a guided projectile and a medium, the method comprises the following steps: constructing an initial particle swarm, and carrying out initialization processing on the initial particle swarm, the initial particle comprises at least one variable of an emission pitch angle and an attack angle; initial particle information of the initial particle swarm is obtained, and the initial particle information comprises speed information and position information of a plurality of initial particles; substituting the initial particle information into the trajectory model, and calculating an objective function value of each initial particle in the initial particle swarm; the target function value is used for representing terminal guidance entering time after the initial particles move according to the trajectory model; and based on the target function value of each initial particle, determining a target particle to obtain data information of the target particle. According to the method, the terminal guidance entering time of the guided projectile can be optimized.

The present invention relates to the field of spacecraft guidance and control, and discloses a model-free integrated guidance and control method for an exoatmospheric interceptor. First, an integrated guidance and control model for an axially uncontrollable interceptor is constructed. Taking into account the measurement noise and external interference present in exoatmospheric interception, a non-singular fixed-time sliding mode surface is proposed, and on this basis, a fixed-time terminal guidance law and an attitude tracking controller are designed to improve the real-time and robustness of the control system. Finally, the fixed-time stability of the control system is proved based on the Lyapunov theory. The present invention can take into account the attitude and orbit coupling of the interceptor during the interception process, thereby achieving precise strikes on the target in the terminal guidance stage.

The present invention discloses a lightweight recognition and tracking method for terminal guidance of ground targets, which relates to the fields of artificial intelligence and unmanned aerial vehicle control technology. S1: Obtain a first-person perspective image through an onboard camera; S2: Input the obtained image into a recognition network, which infers the first-person perspective image to obtain ground target information; S3: Construct a value function based on the confidence level of the integrated object attributes and the object's motion state, and compare it with a preset threshold. If the value function is greater than the set threshold, the identified target is considered to be a strike target, and the tracking module is activated; S4: The tracking module tracks the strike target and calculates the relative line of sight angle of the strike target relative to the position of the fuselage based on the position of the strike target in the image, and transmits the result to the guidance module to control the aircraft to fly toward the strike target, thereby achieving a terminal-guided strike on the target. The present invention solves the problems of the large computing power requirements of the UAV visual terminal guidance neural network, which is difficult to deploy on edge devices and causes misidentification.

The application discloses a simulation solving method for the reachable domain of the terminal guidance section of a space-to-surface missile, and comprises the following steps: obtaining the initial state and the original target state of the missile at the time when the missile enters the terminal guidance, and establishing a missile body coordinate system and an inertial coordinate system; selecting the search angle of a new target in the yaw direction and the pitch direction of the missile body coordinate system, and determining the initial position of the new target in the inertial coordinate system; performing integral iteration on the motion models of the missile and the new target, obtaining the final hit result, and recording the boundary of the reachable domain in the direction; updating the search angle of the new target in the yaw direction and the pitch direction of the missile, and forming the reachable domain of the terminal guidance section of the space-to-surface missile after traversal. The method considers the field of view range constraint of a seeker of the space-to-surface missile, narrows the search range of the reachable domain, and improves the solving efficiency. The solving process is irrelevant to the types and motion characteristics of the space-to-surface missile and the target, has good compatibility for different types of missiles and targets, and is a high-efficiency solving method with certain expansibility.

The application discloses a guidance control method and system for a remote multi-mode composite guided aircraft, in the system, an active radar / infrared multi-mode composite guidance strategy is used in the terminal guidance section, the advantages of the two are complementary, the information acquisition capability of the detection system is enhanced, the detection aspect guarantees the capability of the aircraft to accurately hit the target under the influence of various interference factors, in the middle guidance section, the inertial guidance is adopted for the aircraft, and the over-weight compensation proportional guidance law control is used to make the aircraft to carry out the range-increasing gliding method, the accurate hitting range of the aircraft is improved from 20km to 40km, and in the terminal guidance section, the new terminal sliding mode guidance law is adopted to control the terminal guidance section of the aircraft, on the basis of guaranteeing the accurate hitting, the landing angle error is controlled to be within 2° from 10°.

The application belongs to the field of radar target detection and recognition, and is suitable for solving the problem of detecting low-speed targets by missile-borne radar with large incident residual angle and high sea conditions in the terminal guidance phase of hypersonic vehicles. Firstly, the sea clutter spectrum is approximated by a Gaussian curve to obtain the prior spectrum under different incident residual angles and different sea conditions. The block median method is used for real-time spectrum measurement of sea clutter. The KL divergence is used for sea clutter spectrum estimation and normalization suppression. The spectrum of sea clutter is whitened. After short-time Fourier transform, a new two-dimensional matrix signal is obtained. The support vector machine is used for sea clutter discrimination, so as to realize the anti-sea clutter of terminal guidance radar. The application combines the prior and posterior of sea clutter spectrum, improves the stability and accuracy of detection under clutter, and selects the shape, duration and RCS of sea clutter as characteristic quantities, so that the intelligibility of intelligent identification is strong.

The application discloses a kind of hyperbolic-like front angle remodeling terminal guidance strategy based on relative distance, including the following steps: establishing the kinematic model of missile in two-dimensional plane intercepts stationary target;Design guidance strategy based on hyperbolic equation, solve the geometric parameters of hyperbolic according to the initial conditions of missile and target, obtain the analytical equation of guidance strategy about design parameter;According to the kinematic relationship inverse solution specified angle of different launch angle parameters.The guidance strategy in the application uses equation constraint to front angle information, meets the attack angle constraint under the condition of zero miss distance strike, and end view angle rate converges to 0, and guidance parameters are obtained by numerical integration calculation, without linearization to model, simplify the design of guidance system.The application improves the applicability of guidance system, and the design method is simple and applicable.

The present application belongs to the technical field of guided shell trajectory optimization, and particularly relates to a trajectory design method suitable for radar terminal guided shells, which is suitable for a shell with high initial velocity, small overload capacity and radar seeker to achieve large angle of fall attack on short range targets, and is used for providing good trajectory conditions for the radar seeker to capture targets, and increasing the angle of fall to improve the damage performance. The method takes the "anti-inclined plane attack" as the guiding ideology, comprehensively considers the overload capacity of the shell, the pitch angle at the handover time of the mid-terminal guidance and the angle of fall when hitting the target, and uses the GPOPS optimization toolbox of matlab as the design tool to optimize the design of the minimum range trajectory, which can effectively improve the minimum range capacity of the terminal guided shell, widen the attack range of the shell, avoid the problem of small terminal angle of fall and unfavorable radar seeker capture conditions caused by the conventional low stretch trajectory, and effectively improve the combat effectiveness of the weapon system, and provide a reference for the trajectory optimization design of similar terminal guided shells, and has a wide application space.

The application provides a high-speed impact terminal guidance shift method and device for asteroid defense, and belongs to the technical field of asteroid defense, and comprises the following steps: state confirmation before shift, target asteroid discovery and identification by a terminal guidance system guidance camera, target asteroid identification state setting and sending to the ground by the terminal guidance system, shift instruction sending to the high-speed impact spacecraft by the ground, on-duty spacecraft state setting, stop control and sending of stop control time, corresponding time attitude and orbit data to the terminal guidance system by the middle guidance system, and start control by the terminal guidance system. The application has comprehensive consideration factors and strong engineering practicability. The shift method of the application can realize smooth shift of the high-speed impact spacecraft from the middle guidance system to the terminal guidance system, ensure smooth transition of the high-speed impact spacecraft attitude and orbit before and after the shift, and ensure that the high-speed impact spacecraft can finally successfully complete accurate impact on the target asteroid.

The invention provides a cluster patrolling device flight control system which is composed of a power distribution unit, a main control unit, a navigation unit and a task unit which are connected through inter-board connectors. The power distribution unit comprises an initiating explosive device activation loop and a power-on and power-off control power supply loop; the initiating explosive device activation loop consists of an initiating explosive device battery, a switch switching module, a dual-control switch holding circuit and a time sequence power distribution module; the power-on and power-off control power supply loop consists of a main battery and a control holding switch circuit; the main control unit is configured for health management, flight control and terminal guidance before and after shooting; the navigation unit is configured to carry out data processing on navigation data of the patrol flight device; and the task unit is configured to perform image recognition and task route collaborative planning on the patrolling device and support air-to-air and air-to-ground data interconnection communication. The system is highly integrated, through the high-integration design, the control precision is improved, the data transmission delay is reduced, the equipment size is reduced, and the requirement for stable flight of the patrol aircraft is met.

The application discloses a multi-missile time-space cooperative guidance method, which comprises the following steps: 1) establishing a guidance motion dynamics model of a missile intercepting a target in a three-dimensional space; 2) calculating a guidance time consistency error through information interaction among the missiles, and designing a line-of-sight direction acceleration control input as a control quantity for coordinating the attack time of the missiles; and 3) designing a bottom-layer missile guidance law, i.e. a line-of-sight normal control input, for controlling the missiles to intercept a maneuvering target at an expected line-of-sight angle. The application establishes a time-space cooperative guidance control of multiple missiles intercepting the same maneuvering target in the three-dimensional space based on a guidance time estimation method and a line-of-sight angle constraint, and can truly reflect the relative motion relationship among the multiple missiles and the target and the remaining time of the terminal guidance of the missiles.

The application discloses a kind of air-to-air missile interception probability sensitivity analysis method based on agent model.Interception probability has important significance in missile composite guidance, determines the miss distance and hit probability of missile terminal guidance phase, to effectively improve the interception performance of missile during midcourse guidance and terminal guidance shift, it needs to carry out sensitivity analysis on each error that influences missile interception probability, through sensitivity analysis, the influence size of each error is quantified.However, the timeliness of traditional interception probability calculation method is poor, to solve this problem, the application establishes the support vector machine agent model of particle swarm optimization, to obtain the interception probability calculation model with higher timeliness and accuracy.Based on the agent model, the sensitivity coefficients of 18 kinds of errors are calculated by Sobol index method.The results show that radar angle measurement error and optical radar angle measurement error have the greatest influence on air-to-air missile interception probability.The method shortens the calculation time while ensuring the accuracy, and provides a basis for further sensitivity analysis.

The present application provides an active and passive combined centroid jamming method, which includes calculating the long-range and short-range demarcation values of the ship. The long-range and short-range demarcation values of the ship are determined, and a multi-stage control strategy is adopted to flexibly select different jamming styles according to the distance of the incoming missile from the ship. In the long-range stage, narrowband noise jamming of the phased array active jamming equipment is adopted. In the short-range stage, an active and passive combined centroid jamming method is adopted. The present application uses a multi-stage control strategy, comprehensively considering the characteristics of the single-pulse angle measurement and tracking of the terminal guidance radar, and flexibly selects different jamming styles according to the distance of the incoming missile from the ship, thereby achieving targeted jamming of the seeker in different working stages and effectively improving the success rate of the whole process jamming.

The present application discloses a trajectory prediction and coordinated interception method, system, equipment and medium for maneuverable targets, relating to the field of aircraft guidance technology. The method includes establishing a control system model for the maneuverable target, obtaining a target observation trajectory sequence and filtering it to obtain a target historical trajectory sequence; fitting the coefficients in the initial target increment function model based on the target historical trajectory sequence, and obtaining a target trajectory prediction sequence based on the fitted target increment function model; using the expected intercept point as the mid-range guidance target point of each interceptor missile, using MPSP to plan the mid-range guidance trajectory that meets the attack angle constraint, and determining the mid-range guidance trajectory; delaying the time for the delayed interceptor missile to enter terminal guidance through pulse maneuvers to achieve mid-range guidance time coordination; and adjusting the line of sight angular acceleration using a line of sight angular acceleration control amount to make the hit time of each interceptor missile consistent. The present application can improve the interception accuracy of maneuverable targets.

This invention discloses a multi-vehicle cooperative terminal guidance method and related products with cooperative detection configuration constraints. The master vehicle, based on the finally determined control sequence, interpolates and generates real-time control quantities at time t after the initial terminal guidance time t0 to regulate its motion state; and combines the offline calculated and online iteratively calculated state quantity sequences and their corresponding time sequences [t1; ...; t] N Terminal guidance time t f The master aircraft sends the control sequence to the slave aircraft. Based on the final determined control sequence, at time t after the initial guidance time t0, the slave aircraft interpolates and generates real-time control quantities to regulate its motion state. This involves the offline and online iterative calculations of the state quantity sequences and their corresponding time sequences [t0; t1; ...; t] sent by the master aircraft. N Terminal guidance time t f And the system detects configuration constraints to determine the sequence of control quantities.

The invention discloses a damage performance evaluation system and evaluation method, and belongs to the technical field of military intelligent equipment, and the system comprises a target characteristic perception layer which constructs a digital twinborn body of a target ship and achieves the time-space alignment of multi-modal data, and an intelligent guidance analysis layer which is used for extracting attack path optimization parameters and a terminal guidance radar locking strategy. The target characteristic sensing layer, the intelligent guidance analysis layer, the dynamic evaluation layer and the feedback optimization layer are connected through a high-speed data bus, the dynamic evaluation layer is used for simulating the damage propagation process after detonation of the warhead, the feedback optimization layer is used for dynamically optimizing the attack angle and the detonation opportunity, and the target characteristic sensing layer, the intelligent guidance analysis layer, the dynamic evaluation layer and the feedback optimization layer are interconnected through the high-speed data bus. A closed loop of perception-decision-evaluation-optimization is realized through the arrangement of a target characteristic perception layer, an intelligent guidance analysis layer, a dynamic evaluation layer and a feedback optimization layer, and the technical bottlenecks of excessive simplification of target modeling and deficiency of electromagnetic characteristic static processing and guidance-damage association in a traditional evaluation method are solved by combining the evaluation method.

The invention discloses a high-reuse multi-core heterogeneous parallel processing aircraft control software architecture which comprises a master control CPU (central processing unit), a first slave control CPU, a second slave control CPU and a third slave control CPU. The master control CPU is used as a master control execution unit, is used for performing data interaction with the peripheral, the first slave control CPU, the second slave control CPU and the third slave control CPU, and is responsible for controlling the overall data flow, the time sequence flow and the control flow of the system; the first slave control CPU is used for running a multi-mode navigation calculation task and calculating a position, a posture and a speed required by the control system; the second slave control CPU is used for running a visual calculation task and calculating a sight angle and a sight angle speed required by the terminal guidance instruction; and the third slave control CPU is used for running a path planning and guidance calculation task and calculating flight path planning and guidance instruction related data. According to the scheme, rapid and efficient design of aircraft software is achieved, and software system structure reuse, design reuse and code reuse are achieved.

The application discloses a generation method and equipment of a decoy jamming projectile discrimination model, and belongs to the technical field of infrared homing terminal guidance. The method comprises the following steps: obtaining time-series radiation sequences and attitude sequences based on original images output by an infrared focal plane array in a seeker and attitude data of a projectile output by an inertial measurement unit; coupling the radiation sequences and the attitude sequences by using a first encoder in a preset network to obtain space-time embedding vectors; adjusting a single-layer full connection classification head in the preset network according to the space-time embedding vectors; constructing an initial discrimination model according to the first encoder and the adjusted single-layer full connection classification head; obtaining a decoy sequence according to limit values of physical parameters of a decoy jamming projectile; obtaining an intermediate discrimination model according to the initial discrimination model and the decoy sequence; and training the intermediate discrimination model by using the decoy sequence, the time-series radiation sequences and the attitude sequences to obtain a target discrimination model, which can reduce the false alarm rate of the decoy jamming projectile.

The invention discloses an event-triggered aircraft press-down section self-adaptive guidance method which is characterized by comprising the following steps: S1, entering a terminal guidance press-down section working state; s2, line-of-sight angular velocity information is obtained, and a guidance controller adopts a proportional guidance method to solve a guidance law under an aircraft-target relative motion coordinate system; s3, guidance is carried out through aircraft attitude adjustment to judge whether the target enters the damage radius or not; s4, judging an event triggering mechanism; s5, carrying out strategy evaluation and strategy improvement on the value function constructed by the sight angular velocity and state information of the aircraft; s6, comprehensively evaluating parameters of the neural network and executing the neural network; s7, performing parameter updating on the neural network through a result of the approximate error; and S8, replacing the parameters of the guidance controller with the updated control parameters, and ending the workflow of the aircraft. According to the method, the guidance precision, the aircraft capability and the adaptability to a high-dynamic environment are effectively improved, and the effectiveness of training parameters is greatly improved.

The invention discloses a maximum field angle constraint analysis guidance method based on model prediction, and belongs to the field of guidance control. The implementation method comprises the following steps: defining a sight coordinate system in a longitudinal plane, and establishing a missile-target relative motion model; establishing a field angle change prediction model by taking an initial missile-target connecting line as a reference and combining a multi-constraint optimal terminal guidance expression based on an optimal control theory; initial guidance parameters are given, and parameters needed by a guidance instruction are obtained according to the missile-target relative relation; analyzing the field angle prediction model, deriving the dimensionless time t, finding the dimensionless time of which the derivative is a zero point, substituting the dimensionless time into the field angle prediction model to obtain a maximum value expression of field angle prediction, combining the maximum value expression with the maximum field angle phi maxc, and solving a parameter s2 corresponding to the maximum field angle in all parameters s2; and based on the calculated guidance parameter critical value s2max phi, the guidance parameter s2 is substituted, a longitudinal guidance instruction is obtained, and maximum field angle constraint accurate guidance based on model analysis is achieved according to the longitudinal guidance instruction.