961 results about "Missile" patented technology

Some embodiments of the present disclosure relate to a method for operating a hydraulic fracturing system, the method comprising the steps of: establishing fluid communication between a pumping unit and a fracturing missile; determining if fluids are flowing between the pumping unit and the fracturing missile; if fluids are determined to be flowing, stopping the fluid communication between the pumping unit and the fracturing missile; and establishing fluid communication between the pumping unit and a bleed-off receiver. Other embodiments of the present disclosure relate to a system for performing a hydraulic fracturing operation.

The invention discloses a missile avoidance and rotary enemy locking maneuvering strategy generation method, which comprises the following steps: S1, establishing a kinematics model and a dynamics model of an airplane and a missile based on Newton's law of motion, and simulating flight tracks and states of the airplane and the missile; s2, based on the kinematics model and the dynamics model, establishing a 1v1 beyond-visual-range air combat environment; s3, designing a state space, an action space and a reward function of deep reinforcement learning for the quadruple of deep reinforcement learning, and integrating the three into an over-the-horizon air combat environment to complete the design of a state transfer function; s4, simulating a warplane by using an intelligent agent, and performing repeated simulation confrontation by using the deep reinforcement learning framework to optimize and adjust a strategy and improve the survivability and combat efficiency in an air combat; and through continuous iterative training of the reinforcement learning model, the intelligent agent autonomously learns an effective maneuvering strategy so as to cope with threats of enemy planes and missiles from different directions.

The invention relates to an additive manufacturing method and application of an aluminum-based silicon carbide composite material complex component, and the additive manufacturing method of the aluminum-based silicon carbide composite material complex component comprises the following steps: step 1, uniformly mixing silicon carbide powder and aluminum alloy powder, and drying; 2, establishing a structural model for the target part, and importing the structural model into selective laser melting forming equipment; thirdly, selective laser melting forming is conducted; fourthly, linear cutting and surface treatment are conducted; and fifthly, hot isostatic pressing treatment is conducted. The invention further provides application of the aluminum-based silicon carbide composite material complex component to precise structural components such as aviation structural components, guided missile and satellite components and optical instruments. According to the method, the problem of subsequent large-allowance removal of the aluminum-based silicon carbide composite material component is solved, the wettability between the aluminum alloy and the silicon carbide particles is improved, the mechanical property meets the application requirement, and the aluminum-based silicon carbide structural component with stable performance is obtained.

The invention discloses a missile erection filtering method, system and device and a readable storage medium, and relates to the technical field of inertial navigation and the field of initial alignment. Based on the reference inertial unit pitch angle before erection, the reference inertial unit pitch angle at the current moment, the inertial unit angular velocity on the missile and the trajectory launching inclination angle, whether the missile is in any one of the non-target states of the non-erection state, the just erection state, the to-be-erected in-place state and the erection completion state or not is judged; if yes, measurement updating is not carried out in the Kalman filtering calculation process of the current moment, and time updating is carried out based on the state transition matrix of the current moment, the state vector of the previous moment, the mean square error matrix of the previous moment and the process noise variance matrix; and attitude information output by the inertial navigation system at the current moment is used as a missile attitude result for calculation at the next moment. According to the method, shaking between the guided missile and the launching box in the erecting process before launching is considered, and the attitude precision of initial alignment of the guided missile is improved.

The invention discloses a strategy fusion and differential game-based decentration cooperative guidance method, which comprises the following steps of: aiming at the problem of cluster system collapse caused by missed missile leading, combining the differential game with a consistency formation theory to enable missiles to approach a target in a predetermined formation so as to realize decentration; aiming at the problem that an HJI equation coupled by a nonlinear system is difficult to solve, constructing a neural network by utilizing a self-adaptive dynamic programming technology to carry out online solving on an optimal control strategy of pursuit and escape parties; for the dilemma that multiple escape strategies are difficult to select due to the fact that pursuit control strategies always appear in pairs, an interest value function is introduced, and a strategy fusion mechanism is provided; for the problem of a local minimum value possibly occurring in an artificial potential field method, a dynamic adjustment factor is introduced, and collision avoidance among intelligent agents of the whole cluster system is achieved.

The invention discloses a decentralized differential game cooperative guidance method for a missile cluster under a switching communication topological structure, and the method comprises the steps: introducing a virtual leading missile-slave missile architecture, enabling each missile to construct a virtual leading missile from an initial position, enabling nodes in a simplified virtual leading missile group system to interact through the state information of a topological structure, coordinating a virtual state, and achieving the cooperative guidance of the missile cluster under the switching communication topological structure. And generating a virtual consensus state track so as to realize decentration control. Communication topological structure switching is considered, an optimal control scheme for tracking an expected trajectory is solved in combination with a differential game theory and self-adaptive dynamic programming, and a network weight updating mechanism is designed in the process of solving an optimal control strategy for a nonlinear system line. According to the scheme, the average residence time limiting condition of switching signals is put forward, hypotheses and conditions needing to be met by part of parameters and variables are given, under the scheme, the designed multi-Lyapunov function can achieve exponential convergence, and the stability performance of a missile cluster using the guidance scheme under the topology switching condition is guaranteed.

The invention discloses a device and a method for testing initial velocity and distribution of fragments of a warhead, and the device is characterized in that a missile rack is arranged in a light filtering protection device, and is used for fixedly installing the warhead; the light filtering protection device is arranged between the high-speed camera laying room and the anti-reflection background, located on one side of the lens axis of the high-speed camera and used for enabling fragments generated by warhead explosion to penetrate through the space between the high-speed camera laying room and the anti-reflection background to shield explosion fire light and detonation products. The phenomenon that the lens of the high-speed camera is covered in the light curtain area of the explosion flame to cause overexposure is avoided, so that fragment flying images can be shot at a short distance. According to the testing device, explosion fire light is weakened through the light filtering protection device, the concentration of detonation products flying along with fragments is reduced, and therefore the fragment scattering process can be shot at a short distance, axial distribution of the initial speed of the fragments is directly calculated, and the fragment scattering angle is obtained.

The invention provides an anti-detection camouflage method based on missile-borne tail end guidance features, and relates to the technical field of anti-detection camouflage, and the method comprises the steps: constructing a feature data set through the reflection feature parameters of radar transmission waves with different transmission features under different environment data, transmission angles and distances, and building a reflection feature calculation model; the method comprises the following steps: acquiring guided missile tail end guidance features, performing preprocessing according to the emission angle, distance and environmental data of guided missile radar, predicting reflection feature parameters through a model, modulating radar interference waves with the same reflection feature parameters as the reflection feature parameters to generate an interference blind area with reflection waves, and controlling the position and size of the interference blind area. Blind area coverage is conducted on the position where the missile is located, the blinding index of the missile in the time period is calculated by calculating the moving trajectory fluctuation coefficient of the missile after interference and the change coefficient of a landing area, whether an attack target is lost or not is judged, and whether a tempting signal is released or not is selected.

The invention relates to the technical field of metal additive manufacturing, in particular to a discontinuous dot matrix supporting structure design method for restraining additive manufacturing thermal deformation, and the method comprises the steps that a missile wing discontinuous dot matrix filling area is determined, and the design density of the filling area and the appearance precision index of missile wing additive manufacturing are determined; determining an L-D curve of the unit cell size and the unit cell diameter of the filling area; constructing a three-dimensional geometric model of the missile wing structure on the basis of an L-D curve, performing additive manufacturing thermal deformation analysis, and determining a spacing interval of a body-centered cubic dot matrix by combining the difference between the maximum deformation of the feature points on the outer surface of the missile wing and the appearance precision index of the additive manufacturing of the missile wing; and obtaining the optimal discontinuous spacing according to the spacing interval of the body-centered cubic dot matrix. Through the size design and the spacing design of the discontinuous dot matrix structure, the discontinuous dot matrix supporting structure meeting the missile wing appearance precision requirement and the lightweight requirement is obtained.

The invention relates to the technical field of ballistic calculation, and discloses a rapid ballistic calculation method and system, computer equipment and a storage medium, and the method comprises the steps: generating ballistic calculation parameters based on rasterization, carrying out calculation, obtaining multiple pieces of original ballistic data, carrying out the feature extraction and segmentation processing of the original ballistic data, and obtaining a sample data set; the sample data set is adopted to train a trajectory resolving model, a trained trajectory resolving model is obtained, and the trajectory resolving model comprises a front section trajectory sub-model, a middle section trajectory sub-model, a rear section trajectory sub-model and a trajectory feature sub-model; and inputting to-be-calculated parameters into the trained trajectory calculation model for calculation to obtain trajectory scatter data. According to the method, the deep learning technology and the actual ballistic segmentation characteristics are combined, the rapid ballistic prediction method based on the segmentation training strategy is provided, the rapid ballistic calculation at the millisecond level is achieved on the premise that the prediction precision is guaranteed, and the requirements for missile combat mission planning and decision real-time performance are met.

The invention discloses a decision-making and guidance integrated design method and system based on a Nash equilibrium search strategy, and relates to the technical field of guidance. The method comprises the steps that kinetic parameters of a target and an interception missile group are acquired, and an interception model is constructed; determining an optimal guidance law based on the model and the end collision angle constraint; setting an expected collision angle, and solving through a Nash equilibrium algorithm; the missile angle decision problem is solved according to the optimal collision angle and the expected angle formation; and making a decision-making and guidance integrated control protocol. According to the method, the expected angle decision and tail end angle constraint guidance problems of the missile for the maneuvering target are solved.

The invention discloses a coaxial rotor gun firework launching unmanned aerial vehicle system which comprises a missile-borne shell, an unmanned aerial vehicle and a foldable firework guiding launching system. The missile-borne shell mainly comprises a missile body, a missile tail and a hood; the projectile body and the projectile tail are connected through a fixing mechanism, a shell ejection mechanism is arranged in the fixing mechanism, and shell ejection and separation can be achieved when a separation signal is received. An integral structure formed by the unmanned aerial vehicle and the foldable firework guide launching system is placed in the bomb body in a folded state; the unmanned aerial vehicle is composed of a foldable nose and a fuselage. The foldable firework guiding launching system is composed of a guiding firework launching device and an airborne firework launching frame, the airborne firework launching frame comprises a firework launching tube and a foldable cross-shaped connecting rod, the firework launching tube is connected with the unmanned aerial vehicle through the foldable cross-shaped connecting rod, firework projectiles and propellant powder are contained in the firework launching tube, and the unmanned aerial vehicle carries the firework launching tube to the sky. Remote control is achieved, and fireworks are ignited and launched in an electric ignition mode.

The invention refers to a counter countermeasures device, for use in an anti-aircraft combat. In one embodiment the CCM device together with an anti-aircraft weapon system forms an air defense system. The counter countermeasures device is configured for generating and emitting electromagnetic radiation of appropriate wavelengths and is capable of operating in spoof and jam mode and switching between both modes. The mode of operation is dependent on the selection of the output power of the EM source unit and the distance to the target. The selection of the output power may be done with a mechanical cover or with an electronic solution, decreasing driving current of the EM sources. With the CCM device the anti-aircraft missile launch warning and detection system installed on aircrafts are deceived and the countermeasures are released. With the use of the CCM device the efficiency of the anti-aircraft combat is highly improved, i.e. the probability of the target hit is highly improved.

The invention discloses a device and method for testing the cross-medium water entry of a navigation body under the wave condition, and relates to the technical field of missile body water entry testing equipment.The device for testing the cross-medium water entry of the navigation body under the wave condition comprises an experimental water tank, and a wave maker is arranged in the experimental water tank; the adjusting module comprises a launching platform capable of adjusting the inclination angle; the emission module comprises an emitter and a driving source, the emitter is arranged on the emission platform, the driving source is in butt joint with one end of the emitter, the other end of the emitter is provided with a rectification structure, and high-pressure gas provided by the driving source enables a carbon body loaded on the emitter to be released at a certain initial speed; the rectification structure is used for reducing interference generated by gas turbulence in the radial direction of the launcher when the projectile body leaves the launcher. According to the invention, a stable wave page is generated through the cooperation of the motion phases of the two wave generators, the process that a projectile body enters water under a real sea condition is simulated, and more accurate experimental data is provided for the design and optimization of a cross-medium motion body.

The invention discloses a decoy projectile simulation flow field modeling method based on PINN-U-Net and local-global feature fusion, and the method comprises the steps: constructing a simulation data set of a decoy projectile flow field through numerical simulation, and building a physical loss function meeting a decoy projectile temperature field control equation and boundary conditions; designing a Transform-based global feature extraction module, splicing continuous multi-frame time sequence physical quantity sequences in the data set, adding space coordinate information to form an input vector, inputting the input vector into the global feature extraction module, and extracting and outputting a feature vector corresponding to each physical point; replacing the main body network of the PINN with the U-Net network structure to construct a PINN-U-Net network, and splicing the frame data with the corresponding feature vector of each point of the corresponding frame to form an input tensor of the PINN-U-Net network; under the constraint of a physical loss function, outputting a physical quantity in a decoy projectile flow field; according to the method, the modeling speed and calculation precision of a decoy missile combustion scene are remarkably improved, various limitations of traditional numerical calculation are avoided, and a high-speed and accurate scene modeling function can be provided.

In a rocket motor in which burn rate enhancement (BRE) wires are used to accelerate the burn rate of the solid propellent, embedded charge assemblies (ECAs) are configured as support structures for the BRE wires. Each ECA includes an energetic material that is configured to burn along with the solid propellent to produce thrust and, upon detonation, to break up the solid propellent to terminate thrust. The detonation may also be initiated as a part of process to prevent a higher-order reaction, such as in reaction to heating from a fire or other cause. By being located inside the casing, the energetic material and ECAs do not adversely affect aerodynamics of the flight vehicle of which the rocket motor is a part, such as a missile.

The invention discloses a control method of an attitude-orbit composite lower interceptor, and relates to the technical field of automatic control. Obtaining a strong coupling state space equation according to an angular acceleration calculation formula of the sight angle and a kinetic equation of the interceptor; constructing a high-order multi-dimensional sliding mode matrix according to the strong coupling state space equation, and designing a strong coupling control law of an attitude control engine and an orbit control engine of the interceptor according to the high-order multi-dimensional sliding mode matrix; according to a stable condition and a fixed time convergence condition of an interceptor guidance control system under attitude and orbit combination, solving an interference term in the strong coupling control law to obtain an estimated value of the interference term; and substituting the estimated value of the interference item and the state data of the interceptor into a strong coupling control law of the interceptor to obtain lateral thrust which is generated by an attitude control engine and an orbit control engine and is vertical to a projectile body shaft so as to control the interceptor. The method can improve the control precision of the interceptor.

The invention discloses a space-time guidance-decision integrated control method and system based on distributed optimization, and relates to the field of space-time guidance, and the method comprises the steps: building a dynamic model of each missile, and determining the remaining flight time of each missile; determining a space-time constraint distributed guidance law of each missile based on a preset convergence time of cooperation of the kinetic model, the remaining flight time and the missile group time; determining a distributed optimization problem of the expected end line-of-sight angle based on the remaining flight time and the expected end line-of-sight angle; determining an angle decision protocol of each missile based on the preset convergence time of angle decision, the distributed optimization problem of the expected end line-of-sight angle, the remaining flight time and the expected end line-of-sight angle; and determining a space-time guidance decision integrated control protocol of each missile based on an angle decision protocol and a space-time constraint distributed guidance law, and controlling the normal acceleration and the expected tail end sight angle of each missile based on the control protocol. The attack efficiency of the missile group is improved.

The invention relates to the technical field of active phased-array antennas, in particular to a high-integration low-profile missile-borne phased-array antenna, which comprises an integrated front-end module, an integrated composite cold plate, a power supply and wave control module and a radio frequency comprehensive processing module, the integrated front-end module, the integrated composite cold plate, the power supply and wave control module and the radio frequency comprehensive processing module are arranged in a stacked mode from top to bottom in a tile type framework, the integrated front-end module and the power supply and wave control module conduct heat to the integrated composite cold plate, and the radio frequency comprehensive processing module is installed on the power supply and wave control module. And the heat is indirectly conducted to the integrated composite cold plate. The design of isomorphic integration and heterogeneous layering is adopted, the system integration degree is improved, and the electromagnetic compatibility risk is reduced; a structural and thermal control multifunctional integrated design is adopted, so that low profile, high reliability and high stability are realized; meanwhile, an integrated composite cold plate capable of passing liquid and storing energy in a phase change mode is adopted, and the cold plate has the active heat dissipation capacity and the passive heat dissipation capacity at the same time.

The invention provides an intelligent simulation training bomb for ground attack of an unmanned aerial vehicle. The training bomb comprises a bomb body, a thrust reverser, a fairing, a mounting frame, a smoke tank and a control module. Wherein a gas cabin is arranged in the bomb body, and the thrust reverser is composed of an electromagnetic valve and a multi-way pipe. The mounting frame integrates a fuming tank, a control module and a motion sensing module, and the control module calculates the falling speed, the grounding time and the position coordinates in real time based on data of all sensors of the motion sensing module so as to control the electromagnetic valve to be opened, so that compressed gas in the gas cabin is ejected from a reverse thrust gas opening in the circumferential direction of the fairing through the multi-way pipe; in combination with deformation buffering of the fairing made of an elastic material, nondestructive recovery of the training bomb is achieved, and repeated use is facilitated. And meanwhile, the fuming tank is controlled to trigger and release tracing smoke, so that an impact point is marked, and parameters such as latitude and longitude during landing are sent to a ground guide training system to serve as an evaluation basis of throwing precision.

A missile for delivering a payload and having a longitudinal axis. The missile comprises a nonfragmenation multi-directional munition having a hollow collar. The hollow collar has an inner surface and an outer surface opposed thereto and a first plurality of ports extending therebetween in a substantially radial direction. An ammunition is disposed in each port of the first plurality of ports, oriented outwardly from the longitudinal axis with an ignition source for expelling such ammunition from the collar. The missile also has at least one flight control surface for controlling a trajectory of the missile during flight and a controller for controlling the flight of the missile and the ignition source.

The utility model relates to a missile-borne low-profile, miniaturized and low-frequency-division-ratio dual-frequency antenna, and belongs to the technical field of communication. The dual-frequency antenna comprises a metal bottom plate, an antenna housing fixed with the metal bottom plate, and a dielectric substrate fixed between the metal bottom plate and the antenna housing. Wherein the dielectric substrate comprises an upper-layer dielectric substrate and a lower-layer dielectric substrate which are bonded with each other; metal patches are loaded on the surfaces of the upper-layer dielectric substrate and the lower-layer dielectric substrate; the upper-layer dielectric substrate and the lower-layer dielectric substrate are electrically connected through a metal column located on the central axis of the dielectric substrate, and the lower-layer dielectric substrate and the metal bottom plate are electrically connected through a feed probe located on the central axis; and the metal patches on the surfaces of the upper-layer dielectric substrate and the lower-layer dielectric substrate are horizontally symmetrical along the central axis. According to the utility model, the dual-frequency antenna with low profile, small size and low frequency division ratio is realized, the frequency division ratio is low under the condition that the basic radiation efficiency is guaranteed, and two narrowband working frequency bands can be provided in the s wave band.

The invention discloses a missile-borne navigation receiver rotation compensation method, and relates to the technical field of rotation compensation. The method comprises the following steps: acquiring a complex intermediate frequency satellite signal sequence of a missile-borne navigation receiver, extracting an instantaneous total power sequence through short-time Fourier transform, identifying a rotation period according to the instantaneous total power sequence, and dividing an availability window and a sensitive window; constructing a double-layer rotation frequency spectrum model based on the divided windows, and carrying out online calibration on the double-layer rotation frequency spectrum model; quantizing the monitored abnormal signal into a rotation consistency and interference possibility score by taking the calibrated model as a reference, and determining a discrimination result of the abnormal signal; and performing rotation compensation on the receiver according to a judgment result to generate continuous navigation measurement information. According to the invention, rotation compensation of the missile-borne navigation receiver is realized based on continuous navigation measurement information.

The present invention relates to a vertically arranged nozzle reaction thrust structure with an array distribution, which includes a cabin section housing, a reaction thrust engine, a connecting rod, and an ignition system; the reaction thrust engines are symmetrically distributed in a circumferential array along the inner wall surface of the cabin section housing, and are connected to the cabin section housing through a connecting rod and a nozzle housing flanging structure; the ignition system is a non-electrically transmitted explosion ignition system, which includes an igniter and detonating cords. The non-electrically transmitted explosion ignition system is connected to the cabin section cable through the igniter end and is connected to the reaction thrust engines one-to-one through multiple detonating cords; the axis of the nozzle assembly is perpendicular to the axis of the combustion chamber; an adiabatic ring is bonded to the inner wall surface of the combustion chamber housing, and a converging ring is bonded to the interface of the combustion chamber housing; a throat liner is bonded to the inlet of the nozzle assembly. The present invention can adapt to the axial layout space of the projectile body, increase the effective payload, and improve the missile range under the strict limitation of the axial dimension.

Embodiments of a payload attachment system secure a missile, a suicide drone, or other ordinance to a wing or fuselage of a combat drone (or another type of aircraft) using one or more vacuum mounting modules, wherein each vacuum mounting module comprises at least one vacuum pump controllably coupled to a microcontroller, a vacuum cup fluidly coupled to the at least one vacuum pump, and a transceiver that receives an instruction corresponding to one of a vacuum cup actuation signal or a vacuum cup release signal.

The invention provides a surface target damage effect evaluation method and system, and the method comprises the steps: generating a target region through a preset boundary, and calculating the area of the target region; according to the aiming point in the target area and the weapon ammunition data, N cannonball (or missile) drop points are generated; according to the ammunition drop point and the damage radius, damage circles are generated in sequence; m random points are uniformly generated in each damaged circle; traversing the N * M random points, and counting the situation that the random points in the target area fall into the damage circle; calculating a target damage area according to the random points and the number thereof, the situation that the random points in the target area fall into the damage circle, the damage radius and the area of the damage circle; counting target damage areas of more than preset times, and calculating an average value as an effective damage area evaluation result; and processing according to the effective damage area evaluation result to obtain an evaluation result of the damage rate. According to the method, the technical problems of overlapping calculation, low evaluation accuracy and relatively high calculation resource consumption of the breadth are solved.

The invention provides an optimal throwing method for smoke screen jamming bombs in a multi-machine joint cooperation scene, and belongs to the field of applied mathematics and computational sciences, and the method comprises the following steps: building a missile trajectory equation according to a known enemy missile motion trajectory; constructing a single-machine single-missile shielding model; solving a maximum effective shielding duration putting strategy under a single-machine single-missile model based on a particle swarm algorithm; constructing a multi-machine joint cooperation shielding model; on the basis of a particle swarm algorithm, a genetic algorithm is fused, the maximum effective shielding duration is solved under the condition that the overlapping shielding duration is reduced as much as possible, and an optimal putting strategy is obtained. According to the method, under the constraint conditions of smoke concentration effectiveness, the shortest time interval of unmanned aerial vehicle launching and the like, the overall shielding duration of the multi-smoke-screen cloud cluster can be remarkably prolonged, and optimal configuration and utilization efficiency maximization of shielding resources are achieved.

The invention provides an air passive missile arresting net bullet, and belongs to the technical field of air missile arresting. The arresting net bullet comprises a bullet head, a bullet body and a bullet tail which are connected in sequence; wherein a propelling system, a delayed ignition controller, a gunpowder chamber and a separation chamber are arranged in the bullet tail, the propelling system is used for providing power for the arresting net bullet, the delayed ignition controller is used for controlling ignition of the gunpowder chamber, the gunpowder chamber is filled with separation fuel for separating the bullet body from the bullet tail, and the separation chamber is used for controlling separation of the bullet body and the bullet tail; a flexible blocking net which is in a preset shape after being unfolded is arranged in the projectile body, the vertex angle or the edge of the flexible blocking net is connected with a small traction ball, and the small traction ball is located in a separation chamber at the connecting position of the projectile tail and the projectile body.

The invention discloses a two-stage separation type projectile support for sub-caliber launching of an unstable projectile body. The two-stage separation type projectile support comprises an outer projectile support body, a projectile tail structure, an inner projectile support body and the projectile body. The missile tail structure is driven to decelerate by using the backward opening resistance of the clamping petals, and the missile tail structure is separated from a flying missile body, so that the problem that the missile tail structure with larger mass cannot be separated can be solved; the inner sabot slides backwards along the axis of the projectile body under the action of air resistance in the flying process, so that the flying stability of a combination formed by the inner sabot and the projectile body is improved, target collision at a small attitude angle at a long distance can be achieved, and meanwhile, due to the fact that the flying distance is prolonged, the projectile tail structure and the projectile body are separated more thoroughly; through the arrangement of the resistance table, the inner projectile support can be separated from the projectile body before collision and can also impact the target body together with the projectile body; due to the fact that the inner projectile support can slide in the axial direction of the projectile body and is small in mass, the penetration influence of the inner projectile support on the projectile body can be ignored.

The invention particularly relates to a switching guidance method for tracking a non-escapable area, which comprises the following steps of: firstly, calculating the non-escapable area taking an intercepted target as a center, and then dividing a guidance process into two stages: in the first stage, taking minimization of missile energy loss as an optimization target, selecting a point on the boundary of the non-escapable area as a virtual target point in real time by adopting an Adam optimizer; the virtual target point is tracked by adopting a proportional guidance law, so that the aircraft enters a non-escapable area of the target; in the second stage, a proportional guidance law is adopted to track a real target, and the interception effect on a high-maneuverability flying target is remarkably improved due to the fact that the real target is in a non-escape area.