19 results about "Terminal guidance" patented technology

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

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

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

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

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

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

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

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

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

The application provides an attack angle control method using a neural network and a high-precision analytical solution, and comprises the following steps: step 1, a perturbation model is established; the terminal guidance of a missile is considered as a large earth model, therefore, a dynamic model of the missile is established and small quantities therein are analyzed, a perturbation equation is established, and is used for subsequent solution of a trajectory analytical solution; step 2, zero-order solution solving; after the perturbation equation is obtained, an analytical solution of a zero-order term in the perturbation equation is obtained by using a perturbation method, approximate fitting and mathematical derivation, and is used for subsequent solution of a first-order solution; step 3, first-order solution solving; after the solution of the zero-order term is obtained, a new perturbation equation is further constructed by using the perturbation method, and an analytical solution of the first-order term is solved, and is used for calculation of a guidance law; step 4, resistance coefficient fitting using a neural network; given a simulation state, different resistance change values are obtained through Monte Carlo target simulation, fitting is performed through a quadratic function, fitting coefficients are obtained, and the fitting coefficients are fitted through a neural network; and step 5, attack angle control guidance law; after the above analytical solution formula is obtained, the guidance law needs to measure a current state of the missile, predict a terminal flight deviation, finally, correction is performed through a bias proportional guide, and the attack angle is constrained.

This invention relates to a hardware-in-the-loop (HIL) simulation method for rollover and terminal control. The method designs corresponding control logic based on the control requirements of the aircraft at different stages. In the initial launch phase, the heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle are inertially smoothed. In the terminal guidance phase, the roll angle is exponentially smoothed according to a certain rule based on the missile-target distance. In the terminal attack phase, the heading and pitch line-of-sight angles are reassigned based on the changes in these angles to control the heading and pitch line-of-sight angles of the five-axis turntable. This invention effectively solves the problem in hardware-in-the-loop simulation experiments where the aircraft cannot perform rollover maneuvers and terminal attack simulations due to the limitations of the five-axis turntable's roll frame and the acceleration limitations of each frame angle.

This invention belongs to the field of guided weapon technology, specifically relating to an image-guided terminal-guided rocket weapon system based on ground-assisted identification. The system includes: an image-guided terminal-guided rocket, a ground image control system, a box-type rocket launcher, a command vehicle, and an unmanned aerial vehicle (UAV) reconnaissance relay system. The image-guided terminal-guided rocket employs a satellite / inertial combined mid-course guidance + data link target image transmission for ground target-assisted identification + image-guided terminal guidance method, used for precise strike and destruction of targets. Even when satellites are jammed, the system can achieve precise strike capability against fixed targets and moving targets on land or sea.

This invention discloses a multi-vehicle cooperative terminal guidance method and related products with line-of-sight separation angle constraints. Both the lead and follower aircraft generate the lead aircraft's line-of-sight tilt tracking error, line-of-sight tilt velocity tracking error, line-of-sight deflection tracking error, line-of-sight deflection velocity tracking error, sliding mode surface function related to the line-of-sight tilt tracking error, and sliding mode surface function related to the line-of-sight deflection tracking error based on the command polynomial coefficients and set equations; thereby determining the normal acceleration command and the lateral acceleration command. This method can achieve line-of-sight separation angle constraints.