57 results about "Terminal guidance" patented technology

The present disclosure discloses a method for designing a terminal guidance law based on deep reinforcement learning, and relates to the field of missile and rocket guidance. The method includes: establishing a relative kinematics equation between a missile and a target in a longitudinal plane of a target interception terminal guidance section of the missile; to adapt to a research paradigm of reinforcement learning, abstracting a research problem and modeling as a Markov decision process; building an algorithm network, and setting algorithm parameters, where a selected deep reinforcement learning algorithm is a deep Q-network (DQN); in a terminal guidance process of each round, obtaining a sufficient number of training samples through Q-learning, training a neural network and updating a target network at fixed frequencies respectively, and continuously repeating the above process before set learning rounds are reached.

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 terminal guidance control method based on an all-wheel-drive system, and the method comprises the steps: obtaining an expected acceleration of an aircraft based on a satellite signal and an attitude sensitive system at a mid-guidance section, controlling the aircraft to fly to a target, and in the process, if the aircraft is rejected by a satellite, controlling the aircraft to fly to the target; if yes, the expected acceleration of the aircraft is obtained based on the satellite signal applied at the last moment until the real-time satellite signal is obtained again, and the laser seeker is started after the aircraft is launched for a preset time; after a laser seeker captures a target, the expected acceleration of the aircraft is obtained in real time through a novel sight angle constraint guidance law, a rudder instruction is generated based on the expected acceleration of the aircraft, a steering engine is controlled to conduct rudder work, the aircraft is controlled to fly to the target, and the target is collided at an expected sight angle. The deviation caused by satellite denial is compensated and corrected through the novel sight angle constraint guidance law, and finally the aircraft hits the target.

The invention relates to the field of spacecraft guidance and control, and discloses a model-free external atmospheric interceptor guidance and control integrated method, which comprises the following steps of: firstly, constructing an axially uncontrollable interceptor guidance and control integrated model; in consideration of measurement noise and external interference existing in atmospheric external interception, a non-singular fixed-time sliding mode surface is provided, and a fixed-time terminal guidance law and an attitude tracking controller are respectively designed on the basis of the non-singular fixed-time sliding mode surface, so that the real-time performance and robustness of a control system are improved. And finally, the fixed time stability of the control system is proved based on the Lyapunov theory, the attitude and orbit coupling effect of the interceptor in the interception process can be considered, and therefore accurate striking on the target is achieved in the terminal guidance stage.

The invention discloses a guidance method and equipment with falling angle constraint suitable for a high-speed aircraft, and relates to the technical field of terminal guidance. The method comprises the following steps: establishing a kinematics model and a dynamics model of a terminal guidance section according to trajectory inclination angles of an aircraft and a target, speeds of the aircraft and the target and a relative distance between the aircraft and the target; based on a tail-section small-angle hypothesis, after the kinematics model and the dynamics model are processed, a linear relative motion equation convenient to analyze is established in the direction perpendicular to the sight line; and based on a linear relative motion equation, establishing performance indexes according to a combat target and requirements, and based on the established performance indexes, solving an optimal guidance law analytical solution according to an optimal theory. The method can be directly applied to the high-speed aircraft under the constraint condition, and the accurate striking capacity of the tail end of the high-speed aircraft is improved.

The invention relates to the technical field of unmanned aerial vehicle guidance trajectory planning, and discloses a missile-target relative distance-based terminal guidance trajectory planning method, which comprises the following steps of: S1, acquiring missile-target relative positions and three-axis speeds of an unmanned aerial vehicle body coordinate system according to the position and speed of an unmanned aerial vehicle pre-acquired through a GPS (Global Position System); s2, carrying out cross multiplication on the missile-target relative position and the three-axis speed of the unmanned aerial vehicle body coordinate system to obtain a relative angular speed under an inertial system; s3, acquiring a pitching or yawing angle based on an inertial system sight line system; further processing operation is carried out to obtain a sight line coordinate system direction cosine matrix; s4, in combination with the inertial system angular velocity relative angular velocity, obtaining an inertial system angular velocity, and converting the inertial system angular velocity into a sight line system angular velocity; obtaining the pitch or yaw angular velocity of the sight line system based on the angular velocity of the sight line system; and S5, proportional guidance is carried out by using the pitching or yaw angular velocity of the sight line system. According to the invention, sight angle rate information is obtained from missile-target relative position information evolved from miss distance under a pixel coordinate system, and direct adaptation to conventional flight control is realized.

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 present invention relates to a collaborative mid-course guidance law design method based on a shaping guidance law, belonging to the field of guidance and control technology. It includes establishing a state space equation by introducing the concept of zero-effect miss distance according to the relative motion relationship between the missile and the target; introducing the remaining flight time to design the missile time collaboration term; designing a shaping guidance law with terminal angle constraints according to the state space equation; and introducing the time collaboration term into the shaping guidance law to achieve the guidance law design with angle and time constraints. The present invention combines methods such as multi-agent consensus theory, shaping mid-course guidance algorithm, and offset proportional guidance law, realizes the design of a multi-missile collaborative mid-course guidance law with time collaboration and angle collaboration, and realizes reaching the target area at a fixed angle and simultaneously, forming a better mid-course to terminal guidance handover situation.

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.

The present invention belongs to the research field of radar weak target detection and is applicable to solving the problem of long-time accumulation detection of targets by hypersonic missile-borne radars. Aiming at the problem of difficult detection of low-observable targets by hypersonic missile-borne terminal guidance radars, a target detection method based on micro-motion sliding window and agile long-time coherent accumulation is provided. The computational complexity of long-time coherent accumulation is improved through a hybrid accumulation method of short-time coherent accumulation within a frame and long-time compensation coherent accumulation between frames. Data is updated through the mode of sliding window micro-motion, solving the problem of high data rate output in long-time coherent accumulation and ensuring the real-time performance of target detection. By iteratively using the stationary data within the sliding window, the computational amount is reduced. This method extends the long-time coherent accumulation of early warning radars to missile-borne radars, solving the contradiction between long-time accumulation and high data rate requirements for weak target detection by missile-borne radars. The algorithm has strong real-time performance and is convenient for engineering implementation.

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.

The present invention relates to a technology based on time coordination and belongs to the field of guidance technology. That is, the terminal guidance algorithm can perform time constraints to solve the problem of terminal cooperative strike. By decoupling the time constraint and the impact angle constraint, time control is realized in the lateral plane. And on this basis, taking the remaining time as the coordination variable, a cooperative terminal guidance strategy using a two-layer architecture is used to achieve the cooperative strike of multiple cruise missiles on the target. Thus, the problems of easily reducing the impact angle and time constraint control capabilities of the guidance algorithm and easily causing mutual interference between impact angle control and time control, which affect the miss distance, impact angle and time control accuracy, are solved.

The invention belongs to the technical field of guided weapons, and particularly relates to an image terminal guided rocket projectile weapon system based on ground auxiliary recognition. The system comprises an image terminal guided rocket projectile, a ground image control system, a box-type rocket gun, a command vehicle and an unmanned aerial vehicle reconnaissance relay system. The image terminal guidance rocket projectile adopts a satellite / inertial combination guidance + data link target image return ground target auxiliary identification + image terminal guidance mode, and is used for accurate striking and damage of a target. According to the system, under the condition that a satellite is disturbed, the precise striking capability on a fixed target and a land or sea surface moving target can be realized.

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 technical field of aircraft guidance and control, and specifically discloses a cooperative guidance method for a cluster of multiple aircraft, including: Step 1: Analyze and explore the advantageous positions of terminal guidance to obtain the velocity angle and position angle when having better position advantages; Step 2: Re-input the current missile state parameters into the ballistic model to calculate the remaining flight time to reach the position node; Step 3: Design an adaptive proportional navigation law; Step 4: Verify the simulation results. By proposing a new type of adaptive terminal guidance law, the present invention enters the terminal guidance process through an adaptive proportional navigation coefficient, suppresses the maneuver of the aircraft in the initial stage of terminal guidance, maintains the position advantage and reduces the velocity loss. Finally, it realizes the restoration of normal proportional navigation at the end of terminal guidance to ensure the convergence of the ballistic trajectory.

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 invention relates to a seeker target capturing method and device, computer equipment, a medium and a product, and the method comprises the steps: obtaining a unit missile target line-of-sight vector under a navigation system of each time frame before a missile enters terminal guidance; obtaining a pitching theoretical frame angle and a yaw theoretical frame angle of the missile based on the unit missile target sight vector and the missile flight attitude under the navigation system; fusing the pitching theoretical frame angle, the yaw theoretical frame angle and the triangular wave scanning signal under each time frame to obtain first navigation information; acquiring a first deviation value between an actual pitching frame angle and a pitching theoretical frame angle corresponding to a previous frame of a current time frame when the missile exits terminal guidance and a second deviation value between an actual yaw frame angle and a yaw theoretical frame angle; and acquiring a missile-target distance in the current time frame, and fusing the first deviation value, the second deviation value, the pitching theoretical frame angle and the yaw theoretical frame angle of the current time frame and the triangular wave scanning signal according to the missile-target distance to obtain second navigation information.

The invention discloses a guidance control method and system for a remote multi-mode composite guidance aircraft, in the system, an active radar / infrared multi-mode composite guidance strategy is used in a terminal guidance section, and the information acquisition capability of a detection system is enhanced by utilizing complementary advantages of the active radar and the infrared multi-mode composite guidance strategy; the ability of the aircraft to accurately hit a target under the influence of various interference factors is guaranteed from the aspect of detection, the aircraft is subjected to inertial guidance in the middle guidance section, and the aircraft is controlled through an overweight complementary proportion guidance law to perform range-extending gliding, so that the accurate hit range of the aircraft is increased from 20 km to 40 km; the terminal guidance section of the aircraft is controlled by adopting a novel terminal sliding mode guidance law at the terminal guidance section, and the falling angle error is controlled to be within 2 degrees from 10 degrees on the basis of ensuring accurate hit.