37 results about "Missile guidance" patented technology

The invention discloses a design method of a multi-constraint guidance law driven by a depth deterministic strategy gradient, and belongs to the technical field of missile guidance. The method comprises the following steps: S1, establishing a three-dimensional relative kinematics model of a missile interception target; s2, based on the S1, designing a fractional order sliding mode control law and analyzing the stability of the fractional order sliding mode control law; s3, designing and optimizing an A3DPG algorithm; and S4, inducing a strategy network to learn an action strategy through a reward mechanism in a dynamic environment, and realizing a multi-constraint guidance law with shortest time, small miss distance, field of view FOV constraint, acceleration change suppression, sliding mode surface convergence reward, end attack angle constraint and total reward. According to the method, the guidance law with the convergent line-of-sight angle is designed by using the fractional order sliding mode theory, and then the sliding mode surface parameters and the reaching law parameters are dynamically learned by using the reinforcement learning algorithm, so that the guidance law with dynamic adjustment is realized. And inducing a strategy network to learn an optimal guidance law under the evaluation of the indexes through factors such as a reward mechanism in a dynamic environment.

The invention provides a missile guidance method based on hierarchical reinforcement learning, and belongs to the field of control strategies. The method comprises the following steps: firstly, introducing a missile-target kinematics model and a missile velocity attenuation model; secondly, designing a process reward and a result reward on the basis of the missile-target kinematics model to obtain a total reward function, and guiding the missile to follow the target; thirdly, after reward function design is completed, a flexible action-evaluation algorithm SAC is used for conducting preliminary training on the missile under the condition that the target adopts linear motion, and an initial guidance strategy is obtained; and finally, on the basis of the initial guidance strategy, the enemy plane adopts an evasion strategy, a flexible action-evaluation algorithm SAC is used for continuing to train the missile, and finally a high-quality guidance strategy is obtained. According to the method, the missile can autonomously adjust the flight path in the complex battle condition, interference and challenge in the complex environment are overcome, and therefore the guidance precision and the attack success rate are greatly improved, and the missile striking effect and survivability are improved.

The invention provides a target maneuvering overload tracking and guidance method and system considering measurement noise. The method comprises the following steps: S1, selecting a state tracking variable and a measurable quantity; s2, establishing a theoretical missile-target relative kinematics model; s3, establishing a state tracking model with feedback, the model being a target maneuvering tracker, referred to as a tracker; s4, setting an initial value of the state tracking model with feedback; s5, updating the measured value, and updating the state of the state tracker with feedback; s6, tracking result smooth filtering processing; s7, performing convergence judgment on the tracker, and determining whether a target maneuvering compensation item is added into the guidance law or not according to a judgment result; and S8, judging whether the missile target meets an encounter condition or not, and carrying out loop iteration on the steps S5 to S8 before the missile target meets the encounter condition. And if so, ending the whole tracking and guidance process. According to the method, the problem that the interception strike efficiency of the hypersonic maneuvering target is greatly reduced due to target maneuvering lacking in the interception missile guidance law is solved.

The application discloses a laser beam riding guidance missile roll angle de-rotation measurement system and method based on polarized light, and belongs to the technical field of missile guidance control. The system comprises a laser irradiator for emitting polarized laser to generate a polarized laser field; a polarized detector assembly is arranged on one side of a target missile body close to the laser irradiator and is located in the polarized laser field together with the target missile body; wherein the polarized detector assembly comprises a four-quadrant detector, an accelerometer, a detection circuit and a signal processing module; a rectangular coordinate system is established in the polarized laser field, and a target missile body coordinate system is established. The application can not only measure the real-time roll angle parameter of the missile relative to the laser information field coordinate system, but also saves the loading process of the missile body reference information, simplifies the system complexity and is highly applicable.

The application provides a missile body and seeker installation error compensation method and system, comprising: obtaining a first angle value between a missile body and a target source by using a satellite navigation receiver and a strapdown inertial navigation system, the target source being at a preset distance from a launching frame on which the missile is placed and being within a field of view range of the seeker; obtaining an installation error angle value between the missile body and the seeker according to a rotation angle value of a frame of the seeker and the first angle value; and performing error compensation on an installation error of the missile body and the seeker according to the installation error angle value. Compared with the prior art, the first angle value between the missile body and the target source is obtained, the rotation angle value of the frame of the seeker is compared with the first angle value to obtain the installation error angle value, the installation error is reduced by software compensation according to the obtained installation error angle value, the precision of a line-of-sight angle measured by the seeker is effectively improved, and the precision of a missile guidance control system is improved.

Embodiments of the present application disclose a virtual object control method and device, electronic equipment and a storage medium; the method comprises: acquiring a direction of a sight line and a starting point of the sight line, and displacement and speed of a virtual object; calculating a deviation vector between the virtual object and the starting point of the sight line according to the starting point of the sight line and the displacement of the virtual object; decomposing the speed into speed components in a plurality of specified directions according to the deviation vector; calculating acceleration in the specified directions; fusing the acceleration in all the specified directions to obtain a target acceleration of the virtual object; and controlling the virtual object to move along the direction of the sight line based on the target acceleration. Based on simple speed decomposition and acceleration calculation, the control process of missile guidance is simplified, and the calculation amount in control is reduced, so that the missile guidance can be efficiently applied in games to realize control of the virtual object.

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 embodiment of the invention discloses a guidance instruction correction method and device for an anti-large maneuvering target and a storage medium. The guidance instruction correction method for the anti-large maneuvering target comprises the steps that the maneuvering acceleration of the target is obtained, and a missile-target relative kinematics model is established; obtaining a missile acceleration instruction based on an optimal feedback solution obtained by the optimal feedback method; in combination with a guidance control loop in the missile-target relative kinematics model, performing frequency domain analysis to determine and obtain instruction correction parameters, and in combination with the missile acceleration instruction, obtaining missile acceleration; a feedforward unit and a feedback unit are introduced into the guidance control loop, and a new missile acceleration instruction is obtained by combining an advanced link and a proportional link. The problems that in the prior art, a missile guidance method is high in sensitivity to residual flight time estimation errors, and time delay of a transfer function between needed overload and actual overload is obvious are solved.

The invention relates to the technical field of missile guidance, and discloses a construction method, device and equipment of a hypersonic speed target capture area and a medium. The method comprises the steps that a finite time convergence guidance law is designed according to a relative motion equation between an interceptor and a hypersonic speed target, so that the interception time is shortest when the interceptor intercepts the hypersonic speed target with the finite time convergence guidance law as the acceleration in the terminal guidance section; acquiring a zero-control interception line formed by a plurality of speed preposed angles when the interceptor meets the zero-control interception flow pattern at the terminal guidance section; combining a finite time convergence guidance law and a zero control interception line to construct an initial capture area of the hypersonic speed target; and according to the control parameters of the hypersonic target pulse engine, the field angle and the constraint condition of the intersection angle, the initial capture area is optimized, and the target capture area of the hypersonic target is obtained. According to the method, capture area construction under complex constraints can be realized, and rapid convergence can be realized.

The present invention provides a method for autonomous switching of distributed antennas, which includes: Step 1, four satellite receiving antennas are sequentially and evenly arranged on the missile body along the circumferential direction of the missile body; Step 2, any two adjacent antennas are taken as a group, and the remaining two are taken as another group. One group serves as the working antenna group for satellite searching, and the other group serves as the backup antenna group and is in a non-working state; Step 3, a switching strategy between the two antennas of the working antenna group and the two antennas of the backup antenna group is designed according to the number of frame-synchronized satellites in the corresponding tracking channels of the two antennas of the working antenna group and the time duration measured by the corresponding tracking channel timers. This method can ensure that satellite signals are still available when inertial information is unavailable during the rolling process of the missile, and can continuously provide the position and speed information of the missile for the missile guidance and control system.

The present invention is applicable to the field of missile guidance technology and provides a missile precision guidance method and system, comprising the following steps: obtaining a theoretical flight route and a target position point; detecting the three-dimensional coordinates of the missile in real time to obtain an actual flight route, and determining the engine thrust direction and the rudder deflection angle according to the actual flight route and the theoretical flight route at set intervals; determining the missile position, and when the distance between the missile position and the target position point is less than a set distance value, calculating a predicted flight route each time the engine thrust direction and the rudder deflection angle are changed; determining a prediction error between the predicted flight route and the target position point, and when the prediction error is within a set miss distance, no further adjustment of the flight route is made; otherwise, the engine thrust direction and the rudder deflection angle are continuously adjusted until the guidance accuracy meets the standard. In this way, even if the engine thrust direction and the rudder deflection angle cannot be adjusted in real time, the guidance accuracy of the missile can still be guaranteed.

The invention relates to the technical field of missile guidance, and discloses a missile-to-target guidance method. The method comprises the following steps of: tracking a target party according to a proportion guidance coefficient of a missile party, carrying out a motion process of intercepting and striking the target party, and establishing a motion model of a missile-target motion problem; wherein the proportional guidance coefficient has a function relationship with the speed of the missile party, the speed of the tracking party and the relative distance between the missile party and the target party; the speed of the missile party, the speed of the tracking party and the relative distance between the missile party and the target party serve as states input into the deep Q network, the proportional guidance coefficient serves as an action output by the deep Q network, and the deep Q network is trained; and the trained depth Q network outputs different proportional guidance coefficients for guidance of the missile party to the target party according to different speeds of the missile party, the speed of the tracking party and the relative distance between the missile party and the target party, so that the guidance precision of the missile is improved.

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.

The application discloses a missile angle guidance law determination method, device and equipment and a storage medium. Firstly, according to the relative motion relationship between the missile and the maneuvering target, a state equation is established. Secondly, according to the established state equation, a non-singular terminal sliding mode surface is designed, and a double-power fast reaching law is designed to reach the sliding mode surface. Thirdly, the disturbance of the maneuvering target is expanded into a new state quantity, and an expanded state observer is designed to estimate the state quantity. Finally, considering the singularity of the state and the estimation of the disturbance of the maneuvering target by the expanded state observer, the guidance law of the missile is determined based on the double-power fast reaching law of the non-singular terminal sliding mode surface. The missile guidance law method provided by the application has fast convergence speed and high robustness.

The present invention provides an omnidirectional anti-inclined surface missile based on thrust redirection, which includes a reverse wing assembly, an electric servo, an on-board control module, a warhead, a thrust reversal engine, a guidance component, a connecting cylinder, a flight engine, a tail wing assembly and a launch engine; the reverse wing assembly, the electric servo, the on-board control module, the warhead, the thrust reversal engine, the connecting cylinder, the flight engine, the tail wing assembly and the launch engine are sequentially connected from head to tail to form the missile; the guidance component is arranged at the tail of the thrust reversal engine and is located in the connecting cylinder at the same time; the electric servo executes actions according to control instructions to make the missile fly along the planned ballistic trajectory until it hits the target. The present invention can achieve omnidirectional strikes on buildings and has the ability to precisely strike and suppress targets in urban warfare.

The invention discloses a multi-machine multi-missile collaborative air combat decision-making method based on SM-HAPPO deep reinforcement learning, and relates to the technical field of air combat intelligent decision-making. Starting from a multi-aircraft multi-missile collaborative air combat decision-making method, a dynamic model of an aircraft, a missile guidance model and an air combat situation model are established, and simulation of a complex environment is achieved. For collaborative decision-making of isomorphic or heterogeneous aircrafts, a PPO algorithm is expanded to the field of heterogeneous agents, a GRU-based state memory module is introduced, and an SM-HAPPO algorithm suitable for deep reinforcement learning of the heterogeneous agents is provided. In the strategy network, a loss function with a mask mechanism and a strategy entropy regular term is introduced, and the stability and exploratory performance of the algorithm are improved. Finally, a simulation experiment of aircraft collaborative air combat decision making is completed in a python environment, and it is proved that compared with a traditional intelligent algorithm, the collaborative decision making capacity and the air combat winning rate of the unmanned aerial vehicle can be further improved.

The application discloses a high-precision micro gyroscope based on noise compression effect and a working method thereof, mainly relates to a noise compression light source applied to a micro gyroscope based on a silicon-based photonic crystal microcavity, and can be applied to high-precision inertial navigation, missile guidance and other military fields. The application comprises (1) a noise compression light source generation structure; (2) a micro gyroscope structure based on a silicon-based photonic crystal microcavity; (3) a coupling structure of the noise compression light source and the micro gyroscope based on the silicon-based photonic crystal microcavity; and (4) a micro gyroscope test structure based on the silicon-based photonic crystal microcavity. The application has the advantages that the optical scattering noise and other noise sources of a cavity optical force structure in the micro gyroscope based on the silicon-based photonic crystal microcavity can be reduced, and super-high-precision angular velocity detection can be realized.

The application discloses a missile guidance control integration method based on a neural network mechanism modeling, which comprises the following steps: in a continuous time system, a vertical plane missile guidance control integration model is established; according to the dynamic equation of the vertical plane missile guidance control integration, a state transition equation in a discrete time system is established; a data set for fitting the dynamic model of a neural network is established, and a neural network parameter model obtained through neural network regression calculation is used as a non-mechanism fitting model of a missile control system; and according to the neural network parameter model, a guidance control integration mechanism model used in a control cycle is obtained. The method uses the input and output data of a control process to perform self-modeling of a time-varying control system, meanwhile, the model is equivalent to a mechanism method, and the modeling result can be used for the control of a discrete time system, and the modeling principle has interpretability and reliability.

The invention relates to a complementary training method for improving the robustness of a reinforcement learning interception strategy, and the method comprises the steps: obtaining an average confrontation combat scene, describing the average confrontation combat scene through a first-order delay dynamics model, and defining a defense missile guidance process in the average confrontation combat scene as a Markov decision process; on the basis of a Markov decision process, a strategy is updated step by step by using a PPO algorithm based on a course learning mode, a basic interception strategy is obtained, full-scene performance evaluation is performed on the basic interception strategy, failure scenes with miss distance exceeding standards are screened out to construct a failure scene list, and mixed initial state distribution is established by using the failure scene list. And difference compensation training is carried out through a PPO algorithm to obtain an optimal interception strategy. Aiming at the problem that a specific scene is easy to fail in a complex air combat interception scene, the robustness and reliability of an interception strategy are improved by introducing a failure scene compensation training mechanism, and the task success rate in a diversified combat environment is ensured.

The invention relates to a missile guidance method, device and equipment based on a digital main line and a storage medium. A missile mechanism model is constructed by using development and production stage data on a digital main line, a missile error compensator is constructed by using test verification stage data on the digital main line, and a missile system hybrid model is constructed based on the missile mechanism model and the missile error compensator. And guided missiles are guided through the guided missile system hybrid model. Compared with the prior art, a large amount of missile flight test data can be provided for training through the data link integrated by the digital main line, the complexity of data fusion is reduced, error compensation is carried out on the missile mechanism model through the missile error compensator of the missile system hybrid model, and the problem of missile guidance error deviation is solved.

A computer implementation processor is provided for directing a missile from a radar towards an interception point of a moving target. The processor includes a missile simulator module; a pre-launch module; a trajectory curvature module; and a guidance module. The missile simulator module determines look angle time to the interception point from the radar. This module determines the look angle time remaining to intercept, at which the angular difference between the line-of-sight to the interceptor and the target are below a threshold value. The pre-launch module provides an initial curvature gain value. The trajectory curvature module updates the curvature gain value from the look angle time. The simulator module receives the gain value in feedback. The guidance module receives the gain value and physical characteristic information for the missile and provides acceleration command to the missile to intercept the target.

The present invention discloses an angle guidance method, device, medium, and product with optimal expected line-of-sight angle constraint, relating to the field of missile guidance technology. The method comprises: performing relative kinematic analysis of a coordinated guidance system to obtain a set of kinematic equations for the missile and target; the coordinated guidance system comprises multiple missiles and a target; designing an angle-constrained guidance law based on sliding mode control; designing a capture cost function, and determining an optimal capture terminal line-of-sight angle vector set based on the capture cost function; designing an angle error cost function, and determining an optimal terminal line-of-sight angle vector based on the angle error cost function and the optimal capture terminal line-of-sight angle vector set; and utilizing the angle-constrained guidance law and the optimal terminal line-of-sight angle vector to perform coordinated angle control on each missile in the coordinated guidance system. The present invention ensures both strong anti-interference performance and the optimality of the capture effect throughout the entire process.

This invention provides a missile guidance method based on hierarchical reinforcement learning, belonging to the field of control strategies. The steps are as follows: First, a missile-target kinematic model and a missile velocity decay model are introduced; second, based on the missile-target kinematic model, process rewards and outcome rewards are designed to obtain a total reward function, guiding the missile to pursue the target; third, after completing the reward function design, the Flexible Action-Evaluation Algorithm (SAC) is used to initially train the missile under the condition that the target is moving in a straight line, obtaining an initial guidance strategy; finally, based on the initial guidance strategy, the enemy aircraft adopts an evasion strategy, and the SAC is used to continue training the missile, ultimately obtaining a high-quality guidance strategy. This invention enables missiles to autonomously adjust their flight paths in complex combat situations, overcoming interference and challenges in complex environments, thereby significantly improving guidance accuracy and attack success rate, and enhancing missile strike effectiveness and survivability.

The invention discloses a laser beam driving guided missile roll angle unwinding measurement system and method based on polarized light, and belongs to the technical field of missile guidance control. The system comprises a laser irradiator used for emitting polarized laser to generate a polarized laser field; the polarization detector assembly is arranged on the side, close to the laser irradiator, of the target projectile body, and the polarization detector assembly and the target projectile body are jointly located in the polarization laser field; wherein the polarization detector assembly comprises a four-quadrant detector, an accelerometer, a detection circuit and a signal processing module; a rectangular coordinate system is established in the polarization laser field, and a target projectile body coordinate system is established. The real-time roll angle parameter of the missile relative to the laser information field coordinate system can be measured, the loading process of missile body reference information is omitted, the system complexity is simplified, and the applicability is high.

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.

This invention provides a guidance method based on the missile-target line-of-sight inertial spatial angular velocity, belonging to the field of missile guidance calculation. The method includes: obtaining missile and target position information through the guidance radar of a guidance station; projecting the target and missile onto the horizontal and vertical planes, and calculating the missile-target line-of-sight angle q based on the position information and projection parameters; calculating the change dq of the missile-target line-of-sight angle using a total differential equation to obtain the missile-target line-of-sight inertial spatial angular velocity; and transmitting the missile-target line-of-sight inertial spatial angular velocity to the missile control system for guidance. This invention utilizes the guidance radar to measure the missile and target position information, calculates the missile-target line-of-sight inertial spatial angular velocity, and then transmits it to the missile. This eliminates system errors, minimizes the influence of errors when calculating the missile-target line-of-sight inertial spatial angular velocity, and significantly improves the calculation accuracy of the missile-target line-of-sight inertial spatial angular velocity.

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

The invention belongs to the technical field of missile guidance and control, and discloses a composite adaptive guidance gain regulation and control method based on a missile-target distance and a closed speed difference index, and the method comprises the steps: fusing the dynamic criterion of a far section based on the closed speed difference index and the robust criterion of a near section based on the distance; and the whole-course intelligent and fine management of the guide gain is realized. The missile can actively enhance maneuvering in a middle and long distance to cope with potential threats, and can automatically switch to a stable mode in a short distance to ensure the hit precision, so that the adaptability and interception success rate of the missile in various battle scenes are remarkably improved.

A method for designing a multi-constraint guidance law driven by deep deterministic policy gradients belongs to the field of missile guidance technology. The method comprises the following steps: S1: establishing a three-dimensional relative kinematic model of a missile intercepting a target; S2: designing a fractional-order sliding mode control law based on S1 and analyzing its stability; S3: designing and optimizing the A3DPG algorithm; and S4: inducing a policy network to learn action strategies in a dynamic environment through a reward mechanism, thereby achieving a multi-constraint guidance law with the shortest time, minimal miss distance, field of view (FOV) constraint, acceleration variation suppression, sliding surface convergence reward, terminal attack angle constraint, and total reward. The method utilizes fractional-order sliding mode theory to design a guidance law with line-of-sight angle convergence. Then, a reinforcement learning algorithm is used to dynamically learn the sliding surface parameters and convergence law parameters, thereby achieving a dynamically adjustable guidance law. Furthermore, in a dynamic environment, the reward mechanism and other factors are used to induce the policy network to learn the optimal guidance law evaluated by the aforementioned indicators.

The embodiment of the present invention discloses a collaborative game guidance method and system for multiple leader missiles and multiple follower missiles with interception angle constraints. The multiple leader missiles are modeled as a non-cooperative game guidance problem, and the influence of actuator failure and linearization error is considered. The game guidance law of the leader missile is designed by combining the function envelope idea and the near-optimal control principle. In the line of sight direction, the follower missile guidance law is designed based on the consistency theory and the preset time stability theory to ensure that the sliding mode surface converges within the preset time, thereby realizing a salvo attack of the missile group. At the same time, a finite-time super-helical sliding mode guidance law is designed in the line of sight normal direction to ensure the rapid convergence of the interception angle and effectively suppress the chattering of the sliding mode surface, thereby realizing all-round blockade and attack of the target. The guidance method disclosed by the present invention can achieve saturation attack on multiple targets through the master-slave collaboration of multiple missile groups. When facing interference from unknown factors, the leader missile can still achieve precise strike with a low miss rate, so that the robustness of the guidance system is guaranteed.