737 results about "Missile" patented technology

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

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

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

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

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

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

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

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

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

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

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

The invention discloses an omni-directional transfer device for vertical recovery of a rocket body, and relates to the technical field of rocket recovery, the omni-directional transfer device for vertical recovery of the rocket body comprises a plurality of automatic positioning and locking mechanisms, and the automatic positioning and locking mechanisms are arranged on the circumferential outer side of the rocket body; the automatic positioning and locking mechanism comprises a vehicle body, wherein a sliding block is slidably arranged on the top surface of the vehicle body; the vertical electric cylinder is rotatably arranged on the top surface of the vehicle body, and the upper end part of the vertical electric cylinder is connected with the locking hook through a first piston rod; and the first end of the overturning electric cylinder is rotatably connected with the sliding block, and the second end of the overturning electric cylinder is rotatably connected with the locking hook through a second piston rod. The omni-directional transferring device for vertical recovery of the rocket body solves the technical problems that an existing rocket body locking process is tedious in operation, needs manual intervention, has high requirements for operation space, is high in potential safety hazard, is long in consumed time, cannot be flexibly adjusted according to the diameter of the rocket body, and is small in application range.

An electronic protection (EP) missile has its warhead unit replaced by an electronic warfare (EW) pod which functions to jam enemy air defenses to protect the other missiles in a long rage strike package throughout a mission. The EP missile is otherwise configured similar to armed missiles with which it is launched as part of the strike package.

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

Various examples are provided related to ultrathin polymer-derived ultra-high temperature ceramic (UHTC). In one example, a ceramic coating includes a polymer-derived ceramic (PDC); an UHTC; and a ceramic adhesive comprising alumina (Al2O3). For example, the PDC can include silicon carbide and the UHTC can include a titanium carbide. A thickness of the ceramic coating can be less than 1 mm. In another example, a method includes providing the ceramic coating and applying the ceramic coating to a surface of a structure. For example, the structure can be an aircraft, a ship, a missile, or a rocket.

The invention discloses a missile-borne integrated navigation method and a missile-borne integrated navigation device for adaptive switching and sequential updating. The method comprises the following steps: acquiring angular velocity data and specific force data of an IMU (Inertial Measurement Unit), and performing inertial navigation state prediction based on the angular velocity data and the specific force data to obtain a prediction error state vector and a prediction error covariance matrix at the current moment; based on the prediction error state vector and the prediction error covariance matrix, navigation mode adaptive discrimination is carried out, a navigation mode discrimination result at the current moment is obtained, and first measurement information is determined based on the discrimination result; performing sequential measurement updating based on the first measurement information to obtain a posterior error state vector and a posterior error covariance matrix at the current moment; and performing error compensation correction based on the posterior error state vector and the posterior error covariance matrix to obtain navigation information. According to the invention, the technical problems of navigation interruption of a high-dynamic missile-borne platform in a GPS denial environment and large calculated amount and poor robustness of traditional fusion filtering are solved.

Heavy inert gas insulation layers are provided for one or more components of a launch system for a plurality of missiles. The layer may be integrated into the walls of the components or provided in inserts attached to the components. An inert gas fills a sealed void space in the walls or the insert. The inert gas has a density of at least 1.5 Kg / m3 and a thermal conductivity (Tcond_gas) of no greater than two-thirds of a thermal conductivity of air (Tcond_air) to form the heavy inert gas insulation layer. The inert gas may be Argon, Krypton, Xenon or a synthetic gas and is suitably held at a pressure of 760 Torr (1 atmosphere) or greater at sea level. The heavy inert gas insulation layer delays desensitization or inhibits premature reaction of the energetic materials inside the missiles due to high external temperatures. The insulation layers allow for more compact and dense configurations of the launch system and missiles.

A missile launcher dustproof plate punch forming system comprises a forming die, an edge folding die and a punching die which are sequentially arranged and installed on a press machine, the forming die comprises a forming female die (1) and a forming male die (2), a dustproof plate profiling groove (3) is formed in the bottom of the forming female die (1), and a dustproof plate profiling protrusion (4) is arranged on the top of the forming male die (2); the through hole punching and bending device has the advantages that the through hole punching and bending speed of the dustproof plate of the missile launcher is high, product precision is high, and labor intensity of workers is relieved.

The invention relates to a trajectory planning method based on improved sampling reinforcement learning. The method comprises the following steps: constructing a dynamic environment model; based on the dynamic environment model, defining a learning agent and a Markov decision process; performing algorithm parameter initialization based on the defined learning agent and a Markov decision process; based on the initialized algorithm, generating an original bulletproof trajectory through interaction between the intelligent agent and the environment; carrying out thinning processing on the original trajectory to generate a thinned trajectory; the thinning track is stored in an experience playback buffer pool, and strategy updating is carried out by using the thinning track; and based on the updated strategy, executing a defense penetration strike task. According to the method, the sampling precision and efficiency of deep reinforcement learning training in aerospace guidance control are effectively balanced, the strategy generalization ability and the training stability are improved, and a feasible scheme is provided for intelligent game strategy optimization in a complex dynamic scene; the method is especially suitable for engineering scenes with high requirements on real-time performance and reliability, such as missile penetration and target interception.

The invention discloses an electromechanical fuse security mechanism based on gravity and multi-group spring linkage, and belongs to the technical field of electromechanical fuses. Aiming at a narrow space of a cylindrical missile cabin of an unmanned aerial vehicle and a launching scene without strong environmental force, a pull ring is triggered to be separated through gravity, a U-shaped locking block spring, a baffle spring, a sliding block fixing pin spring and a sliding block spring are driven to work cooperatively, double insurance is relieved in sequence, and alignment of a booster tube and a detonating tube is achieved. The mechanism abandons the traditional dependence of recoil force and the like, is driven by the combination of the gravity of the projectile body and the spring, has the advantages of compact structure, high space utilization rate, strong environmental adaptability and the like, effectively improves the safety and damage efficiency of small unmanned aerial vehicle ammunition, and is suitable for ammunition fuze design of special mounting scenes such as coaxial unmanned aerial vehicles and the like.

The invention relates to an aircraft missile avoidance decision-making method based on LSTM and PPO, and the method comprises the following steps: 1, collecting data, and 2, carrying out the normalization processing and feature fusion of the observation information of an aircraft and a missile through a long and short term memory network according to the data collected in simulation, and building the time sequence features of the aircraft and the missile. And step 3, establishing a hierarchical model based on a depth deterministic strategy algorithm in deep reinforcement learning, and controlling the rod amount to realize specific flight control. 4, respectively establishing communication between the client side and the server side according to the defined model, and 5, repeating the steps 1-4 until a PPO algorithm training ending condition is reached, and terminating the training. According to the scheme, the flexibility and the real-time performance of decision making are improved, the dependence on traditional modeling and expert knowledge is reduced, and a more intelligent solution is provided for modern air war.

The invention discloses a fragile front cover of a launch canister, a preparation method and a test system comprising the fragile front cover, the fragile front cover comprises a cover body and a flange part annularly arranged on the periphery of the cover body, at least three radial weakening grooves are formed in the front face of the cover body, an annular weakening groove is formed in the position, close to the edge, of the back face of the cover body, and the section of the annular weakening groove is in a single-slope wedge shape. Comprising a vertical side wall and an inclined side wall, the vertical side wall is vertically connected with the end face of the flange part, an angle a is formed between the inclined side wall and the end face of the flange part, and the inclined side wall extends from the groove bottom of the annular weakening groove to the surface of the cover body and intersects with the vertical side wall to form a directional crack propagation area, so that cracks preferentially extend along the angle a under the impact load effect. Through the design that the integrated polyurethane foaming body is combined with the directional crack propagation area, the actually-measured bursting force peak value is smaller than or equal to 7 kN, and the low-impact requirement of the missile fairing is met. The radial weakening grooves in the front face and the wedge-shaped annular weakening grooves in the back face achieve the synergistic effect, so that the maximum size of the fragments is far smaller than a safety threshold value.

The utility model discloses a launcher structure for missile simulation training, which is applied to the field of missile simulation training and is characterized in that a motor rotates to drive a worm to rotate, the worm rotates to drive a worm gear to rotate, the worm gear rotates to drive a rotating shaft to rotate, the rotating shaft rotates to drive a second gear to rotate, and the second gear rotates to drive a first gear to rotate. The first gear rotates to drive the supporting column to rotate, and the supporting column rotates to enable the direction of the launching stand to be changed, so that the aim of conveniently adjusting the launching direction is achieved, the workload is reduced, and the working efficiency is improved. The first bevel gear rotates to enable the screw rod to drive the sliding block to move, the sliding block moves to drive the sliding barrel to move, the sliding barrel moves to enable the supporting rod to drive the launching stand to rotate, the angle of the launching stand is changed, and therefore the aim of adjusting the launching angle is achieved, the use cost is reduced, and practicability and applicability are improved.

The invention provides a jet pipe array missile pitching channel jet flow opening method, device and equipment and a medium. The method comprises the steps that corresponding total thrust is calculated according to pitching control torque needed by a missile; generating a plurality of spray pipe opening combinations capable of providing total thrust according to the thrust of a single spray pipe in the spray pipe array, and taking the spray pipe opening combinations as initial candidate modes; a sequential constraint rule from the middle to the two sides is applied to each nozzle opening combination in the initial candidate modes, and candidate modes conforming to the constraint rule are obtained; calculating a jet amplification factor of the candidate mode conforming to the constraint rule in the current flight state, and evaluating the influence degree of jet interference; the candidate modes with the jet amplification factors smaller than the preset minimum threshold value of the jet amplification factors and conforming to the constraint rule are removed, and the remaining candidate modes conforming to the constraint rule serve as jet opening modes, so that the problem that the control effect is poor due to poor control torque adjustability, control coupling, insufficient spray pipe utilization rate and jet interference influence is solved.

The invention belongs to the field of aircraft missile pylon ejection system control, and particularly relates to a switch array pilot aircraft missile pylon ejection valve and a control method. The ejection valve comprises a flow valve and a pilot stage for controlling the flow valve, and the pilot stage forms a hydraulic half-bridge through a plurality of electromagnetic switch valves connected in parallel, a fixed damping hole and a switch type damping valve controlled by pressure difference. During working, when the electromagnetic switch valve is fully closed, the switch type damping valve is normally opened, and the flow valve is closed; any electromagnetic switch valve is opened, the switch type damping valve is closed under the action of differential pressure, and the flow valve is rapidly opened to supply oil. The redundant design that the multiple electromagnetic switch valves are connected in parallel is adopted, a traditional single pilot electromagnetic valve low in reliability is replaced, normal work under the condition that the single electromagnetic switch valve breaks down can be achieved, and the overall reliability of the system is remarkably improved. During ejection work, the drift diameter of the switch type damping valve is decreased under the action of differential pressure, it is guaranteed that a valve element of the flow valve is rapidly opened, and the response speed of the ejection valve is increased.

A hybrid pad system positioned between a missile and a launch tube is provided herein. The hybrid pad system includes an outer sheet and an inner sheet spaced from the outer sheet. A web region can be positioned between the outer sheet and the inner sheet. A restoration region can be positioned between the outer sheet and the inner sheet and tangentially offset from the web region. A support can be operably coupled with the inner sheet.

The invention discloses a synthetic attack angle control method during aircraft-rocket separation, and relates to the technical field of spaceflight, and the method comprises the following steps: controlling a deviation value between a pitching program angle and a trajectory inclination angle to be smaller than a preset deviation threshold value, and adjusting the pitching program angle in real time according to the trajectory inclination angle, thereby reducing a trajectory attack angle; when the ballistic attack angle is within a preset ballistic attack angle threshold value, an overload difference value between the actual normal overload at the current moment and the normal overload at the corresponding current height is calculated, and the correction amount of the pitching program angle is calculated based on the overload difference value; the correction amount of the pitching program angle is compensated into the pitching program angle, and the synthetic attack angle of the rocket is converged, so that the wind attack angle is reduced; before device-rocket separation, transverse normal acceleration data are continuously collected, and if the transverse normal acceleration data exceed a change threshold value, device-rocket separation is carried out after the acceleration data are recovered to be normal. According to the method, potential safety hazards caused by consideration of a single factor can be effectively avoided, and the safety and reliability of the separation process are remarkably improved.

The invention belongs to but not limited to the technical field, and discloses a single-point training missile traction interception net launching device and method, the device comprises a base, a power device is arranged on the base, and a launcher is arranged at the front end of the power device; and the intercepting net is connected with the emitter through a traction rope. The folding mode of the intercepting net comprises four-corner star folding and parallel folding. And the intercepting net is provided with a mass block and a non-uniform buoy. The working environment of the underwater navigation body touch net associated body motion model is in seawater; a plurality of counterweights are suspended in sea water by the two buoys through the net body, the net body is vertically unfolded by tensile force between the buoys and the counterweights, and the net body is basically not affected by stress in the horizontal direction and moves along with sea waves; when the navigation body vertically touches the net body at a certain navigation speed, the net body drives the buoy and the counter weight to do associated motion.