1381 results about "Missile" patented technology

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

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

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

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

The invention discloses a method for identifying aerodynamic parameters of an adaptive volume Kalman filter aircraft, which is used for improving a traditional Kalman filter algorithm so as to solve the common problem of negative determination of a covariance matrix in practical application. Specifically, singular value decomposition is adopted to replace traditional Cholesky decomposition to decompose a covariance matrix, and the improvement can improve the stability and precision of the algorithm, so that the aerodynamic parameters of the missile are more reliably estimated; in the filtering process, the noise variance is corrected according to real-time measurement information, so that the state estimation error is reduced; according to the method, the adaptive capacity to the dynamic change of the system can be improved, and the precision and robustness of parameter estimation are further improved.

The invention discloses a missile wing unfolding unsteady aerodynamic force measurement test method, and belongs to the technical field of aviation aerodynamic force wind tunnel special tests. The method comprises the following steps: 1, measuring the rotational inertia of the missile wing; 2, measuring blowing test data; 3, calculating an inertial load; 4, the inertial load is deducted, and a pure aerodynamic load is obtained; the invention aims to solve the problems that in the existing missile wing unfolding wind tunnel test, the unfolding track difference is large under the wind condition and the wind-free condition, and the error is large due to the existing data processing method, the method can be used for any unfolding track working condition, the application range is wide, meanwhile, the inertia force in the missile wing unfolding process can be accurately deducted, and the reliability is high. Pure unsteady aerodynamic force is obtained; the moment of inertia of the missile wing can be measured by directly utilizing the existing test conditions, the cost of transformation of the test device is reduced to the greatest extent, the overall cost of the improvement and implementation process of the method is low, and the test efficiency is greatly improved and the time is saved.

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

The invention discloses a variable gain inversion maneuvering target interception guidance method considering attack angle constraint, and belongs to the technical field of aircraft guidance control. The implementation method comprises the following steps: considering necessary conditions for interception success and end attack angle constraints, based on trajectory shaping, representing a sight angle reference profile by constructing a polynomial function taking a flied distance as an independent variable, and realizing reduction of initial and end constraints; the profile only utilizes missile-target relative information and does not depend on motion information of the target, so that the profile is suitable for intercepting targets in various motion forms. In order to accurately track a line-of-sight reference profile, adverse effects of aircraft control input saturation constraint and target maneuvering interference on a guidance system are considered, an input saturation compensator and a finite time interference estimator are introduced, and a variable gain interception guidance law is designed based on an inversion control architecture. And the aircraft can accurately track the profile and can intercept the maneuvering target at an expected angle at the same time. The method has the advantages of low required overload, high interception precision and good robustness.

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

The invention discloses a distributed cooperative guidance method and system for striking a maneuvering target and electronic equipment, and relates to the technical field of target guidance, and the method comprises the steps: obtaining the speed and normal overload of the maneuvering target and a current missile, and a missile-target sight angle between the current missile and the maneuvering target; based on the speed and the normal overload of the current missile and the maneuvering target, and the missile-target sight angle, determining the estimation of the residual flight time; calculating an estimation consistency error based on the estimation of the remaining flight time of the current missile and the estimation of the remaining flight time of all non-current missiles; and determining parameters of a fixed time jammer and a distributed time cooperative guidance law based on the speed, the normal overload and the estimation consistency error of the current missile, the missile-target sight angle, and the speed and the normal overload of the maneuvering target, obtaining a target fixed time jammer and a target distributed time cooperative guidance law, and guiding the current missile. According to the invention, the ambulance attack on the maneuvering target is realized.

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

According to the missile-target matching planning method and device based on deep reinforcement learning and the medium, the deep reinforcement learning algorithm based on the DQN is provided for the characteristics of large solution space, discretization, dynamic performance, nonlinearity and the like of a firepower-target distribution problem, the six-layer full-connection feedforward neural network is combined with the Q-learning algorithm, and the firepower-target distribution problem is solved. According to the method, the perception ability of deep learning and the decision-making ability of reinforcement learning are fully exerted, through model performance test comparison, the method is high in fitting ability, high in convergence speed and small in variance jitter, the distribution result conforms to combat expectation, and a certain reference can be provided for commander firepower attack problem decision making. The method has remarkable advantages in solving the problem of dynamic resource scheduling and distribution, the ammunition-target distribution efficiency can be effectively improved, and the overall combat benefit is maximized.

The invention relates to the field of missile interception, and particularly discloses a guidance and control integrated design method considering flight state constraints, which comprises the following steps: S1, establishing a three-dimensional guidance and control integrated model; s2, designing an algorithm for striking or intercepting a maneuvering target at an expected angle based on the model in S1; s3, designing a guidance control algorithm under a state constraint based on the algorithm in S2, and establishing a six-degree-of-freedom guidance and control integrated model according to the kinematic model, the kinetic model and the three-dimensional missile-target relative kinematic model of the aircraft, so as to avoid the influence of inconsistent time constants of the inner and outer ring kinetic models on guidance efficiency, and improve the guidance performance of the aircraft. According to the scheme, flight condition constraints such as the flight attack angle and the attitude angle rate are considered, a state-considered constraint guidance and control integrated method is designed based on the safety constraint control theory, and the flight safety and stability of the interception missile are effectively improved.

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

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

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

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

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

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

The invention discloses a semi-physical simulation system and method for a gliding extended-range guided aircraft. The simulation equipment comprises a rolling motion simulation turntable system, an electric steering engine load simulation system, a sensor measurement system, an information setting module, a gliding extended-range guidance aircraft trajectory calculation module, a data communication module, a data display and system evaluation module and the like. Missile-borne test parts comprise a missile-borne flight control computer and the like. The information setting module sets trajectory initial information and pulling deviation information, and the gliding extended-range guided aircraft trajectory resolving module calculates aircraft trajectory information and sends the aircraft trajectory information to the rolling motion simulation turntable system, the electric steering engine load simulation system and the missile-borne flight control computer; the missile-borne flight control computer performs trajectory control according to a preset control strategy to obtain a lift airfoil angle control instruction, and the lift airfoil angle control instruction is controlled to a target angle by a double-rotation electric steering engine; the data display and evaluation module collects related simulation and control data, and evaluates, detects and optimizes missile-borne tested parts.

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

The invention provides an RADP-based hypersonic-velocity interception missile adaptive mid-guidance method and system, and belongs to the technical field of hypersonic-velocity interception missiles. Firstly, calculation and judgment are carried out according to obtained motion information of an interception bomb body and a target, a lateral guidance instruction is solved through an RADP method after triggering conditions are met, and the problems of high strike precision and large calculation amount caused by uncertain external interference are solved through the method in combination with a robust self-adaptive dynamic programming method. According to simulation verification, the scheme can obviously improve the resolving efficiency, and the method has a good guidance effect.

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

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

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

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

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

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

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

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