304 results about "Missile" patented technology

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 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.

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.

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.

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 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 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.

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.

1. The name of the design product: neutral pen (missile). 2. The use of the design product: for writing. 3. The design points of the design product: in shape. 4. The picture or photo that best indicates the design points: perspective view 1.

The present invention relates to a missile (10) with a firing device (20) having several barrels (11) extending in the direction of a longitudinal axis (A) of the missile (10) and arranged inclined or skew relative to the longitudinal axis (A), wherein projectiles (12) provided with propellant charges (13) for ejection from the barrels (11) can be loaded into the barrels (20) and a method for operating such a missile (10).

The utility model provides a kind of automatic stop's suspension general hanging beam device, fixedly connected on ground or vehicle-mounted launcher, missile and guide launcher form missile launcher combined system, missile launcher combined system is hung in suspension general hanging beam device, press tightly stopper, fixed guide launcher and electrical connector pass through suspension general hanging beam device, suspension general hanging beam device includes beam body, automatic stop mechanism, hook assembly and anti-swing stop mechanism, wherein, automatic stop mechanism, hook assembly and anti-swing stop mechanism are respectively installed on beam body, when guide launcher is hung, hook assembly hook can be automatically inserted into lifting lug hole, with safety pin stop hook, guide launcher can be hung, i.e. It is realized;When guide launcher is unloaded, pull handle, drive hook to be stopped by automatic stop mechanism, guide launcher can be conveniently unloaded, and operation is simple.

The application discloses a missile and unmanned aerial vehicle body separation device and an unmanned aerial vehicle, and relates to the technical fields of unmanned aerial vehicles and missiles.The device comprises a missile body flange, one side of the missile body flange being used for being connected with the tail of a missile; a body flange, one side of the body flange being used for being connected with the end of an unmanned aerial vehicle body; a wrapping tape combination, the wrapping tape combination comprising a wrapping tape and a plurality of clamping blocks, the wrapping tape being connected to form a ring-shaped structure through a cuttable structure, the plurality of clamping blocks being distributed in a circumferential direction on the inner side of the ring-shaped structure when the wrapping tape forms the ring-shaped structure, and a clamping groove being arranged on the inner circumferential surface of the clamping blocks; a separation spring, the separation spring being arranged between the missile body flange and the body flange, the clamping groove being capable of accommodating the missile body flange and the body flange after the separation spring is compressed; and a pulling spring, the pulling spring being arranged between the wrapping tape combination and the end of the unmanned aerial vehicle body; the separation device realizes the quick separation of the missile and the unmanned aerial vehicle body through the cuttable structure, and realizes the quick attack function of the missile and unmanned aerial vehicle integrated aircraft.

Aircraft multiple sensitive small micro-eye optical / IR (visible light and / or infrared) cameras distributed in many suitable places in the aircraft so that collectively they are able to visualize and scan all the surroundings. With AI analysis and processing, an integrated panoramic entire-scope view with some stereoscopic perception effect could gained. After few early trial-flights of each aircraft, AI should be able to unit images collected by all cameras into one integrated view of all surrounding universe and build up surrounding concentric virtual sensory shells. These cameras will act as micro eyes to detect all surroundings in space all the time. It spot any approaching anti-aircraft missile as a hot spot increasing rapidly in size and intensity so that AI will immediately activate appropriate protective maneuvers to escape the attaching missile(s) while avoiding accidents with any other surrounding aircrafts (objects). Suitable cable connections added for power supply and information sharing to central data controller processor controlled by AI. For existing aircrafts, it could added as add-on camera clusters with suitable cable connections. For future newly manufactured aircrafts, it should added to design in a better and proper way. This innovation could use also in many other military and civil industries.

This application provides a method, apparatus, and medium for optimizing ultra-low-altitude ballistics based on optimal initial value setting. The method includes: determining the optimal intercept angle in the environment; constructing a response surface model and solving for the optimal initial angle; and performing sample point calculations to complete the ballistic optimization. The initial value setting method is applicable to the initial guidance phase of a missile, and the trajectory and ground-grabbing angle can be corrected by setting the parameters of the launch tilt angle. When applying the launch angle setting optimization method to achieve ultra-low-altitude optimized ballistics, the guidance and control system itself is not changed, and it is not necessary to add information about the target measured by the missile's onboard guidance system. Only the missile launch angle needs to be set before launch, making the engineering implementation simple and not reducing the missile's anti-jamming performance. This application provides a new technical approach to improving the ultra-low-altitude interception capability of air defense missiles.

The invention provides a magnetoelectric angular displacement sensor, and belongs to the technical field of design of steering engine system sensors for airborne missiles. The sensor comprises a rotating shaft connected to an output shaft of the steering engine through threads, a magnet pressed at the rear end of the rotating shaft, a front shell fixed to a shell of the steering engine, a control panel assembly pasted to the front shell, a rear end cover fixed to the front shell and heat insulation cotton arranged between the rear end cover and the control panel assembly, one end of the front shell is provided with a cavity used for containing the rear end of the rotating shaft, the control panel assembly is used for detecting the magnetic force of the magnet, and the distance between the control panel assembly and the magnet is not larger than 0.5 mm. The sensor is high in precision, small in size, strong in anti-interference capability, good in environmental adaptability and long in service life.

This invention relates to the field of unmanned aerial vehicle (UAV) interception technology, specifically to a UAV swarm for nuclear missile interception and its interception method, comprising: UAV modules, a communication module, and a coordination module; multiple UAV modules constitute a UAV swarm; the communication module is connected to each UAV module to achieve encrypted information transmission; the coordination module is constructed based on multi-agent reinforcement learning, and is connected to the UAV modules through the communication module, and is used to control the UAV swarm to cooperate in interception; this application achieves stable operation in a nuclear environment through the UAV modules, and constructs a protective shield to intercept nuclear missiles through the coordination module.

This invention discloses a multi-level interlocking security device, including a base, a primary inertial safety mechanism, a secondary launch safety mechanism, a tertiary electromechanical safety mechanism, a first security seat, a second security seat, a PCB circuit board, and a release signal switch. The primary inertial safety mechanism can only release when the missile reaches a predetermined acceleration within the launch tube, achieving the required inertial force. The primary inertial safety mechanism can only release within the launch tube. If the missile does not reach the predetermined acceleration value within the launch tube, the secondary launch safety mechanism locks the primary inertial safety mechanism, preventing further release and preventing the secondary launch safety mechanism from completing its release action. The tertiary electromechanical safety mechanism can only release after the primary inertial safety mechanism and the secondary launch safety mechanism have released; otherwise, it will not release under any circumstances, ensuring the absolute safety of the missile.

A missile telemetry antenna array based on feed phase difference and screw coupling, comprising: inverted F antenna, missile cabin, rectifier block, RF cable and transmitter; a plurality of inverted F antennas are evenly distributed around the central axis of the missile cabin, and the inverted F antennas are fixedly installed on the outer wall of the missile cabin; the inverted F antennas are connected with the transmitter through the RF cable; a rectifier block is arranged above the head of each inverted F antenna; the rectifier block is fixedly installed on the outer wall of the missile cabin; and the rectifier block is used for flow guiding, so that no stagnation point of aerodynamic heating appears on the inverted F antenna. The present application optimizes the gain depression of the axis direction of the missile body, realizes the protection of the aerodynamic heating and force of the antenna by installing the rectifier block in front of the antenna, generates electromagnetic resonance and induced current through the installation screw of the rectifier block, realizes the secondary radiation of electromagnetic waves and the reflection guiding effect.

This invention discloses a method for correcting positioning errors of missile-borne SAR, specifically relating to the field of satellite positioning technology. It is used to solve the problem of unstable positioning error correction caused by the non-rigid deformation between the antenna phase center and the inertial navigation unit not entering the continuous solution chain during high-dynamic flight. By synchronously acquiring satellite positioning observation residuals and raw acceleration data, an acceleration gradient sequence is generated. In the first update cycle, the reference non-rigid deformation vector is extracted and a model degradation command is generated. In the second update cycle, the dynamic spatial offset of the antenna phase center is extrapolated, and then a Doppler frequency shift compensation sequence is generated by projecting along the instantaneous line-of-sight direction of the synthetic aperture radar and fed forward into the radar processor.

The invention relates to a storage and transportation integrated box-opening-free test packaging box which comprises a box body and a supporting protective cover, and the box body and the supporting protective cover are buckled in a sealed mode and can be fixedly connected. More than two supporting seats are arranged on the inner side of the box body, at least one supporting seat is fixed in the box body in a suspended mode, the supporting seat is provided with a first pressing plate used for pressing the guided missile into the supporting seat, and integrated conductive buffer rubber is arranged on the abutting face of the first pressing plate and the guided missile; the in-box ground wire is installed on the first pressing plate and connected with the grounding pile, and after the guided missile is pressed, the guided missile releases static electricity through the integrated conductive buffer rubber, the in-box ground wire and the grounding pile in sequence. The box body is provided with a testing seat used for being connected with testing equipment in an inserted mode, and the testing seat is connected with a missile electrical interface. The box body is provided with a storage cylinder embedded in the box body and used for storing and taking files and tools independently. The packaging box serves as an integrated packaging box for storage, transportation and testing operation and has the function of completing missile testing and detection under the condition that the box is not opened.

The application discloses a kind of image guide head unblock post-missile safety control method, first improvement tracking algorithm, enhance the ability of target recapture after unblock, improve tracking reliability, reduce the risk of missile out of control;In launch after regardless of mode, shooter does not carry out any intervention to missile, such as appears short-term unblock, then the pitch, yaw line-of-sight angular velocity rate output by guide head after unblock keeps the state before unblock, if more than a certain time fails to recapture target, then it is regarded as tracking failure;If tracking failure and autonomous identification failure and last for a certain time, then start down pressure instruction, control missile to explode on the trajectory line near target.In missile self-borne image transmission mode, shooter uploads instruction operation guide head by image transmission system, directly corrects aiming point in guide head automatic tracking state, relocks or changes attack target, and missile automatically flies to target;The application does not additionally increase hardware, so it is simple to operate, low in cost.

The application discloses a beyond-visual-range air combat decision-making method based on dynamic fire field and artificial potential field fusion, and relates to the field of aerospace. The method comprises the following steps: establishing a beyond-visual-range air combat simulation environment, modeling enemy and friendly camps in the air combat; grouping the enemy camp in real time; constructing a time-varying killing performance model of a missile to determine a three-view dynamic fire field of an aircraft-missile-target so as to establish an artificial potential field of combat groups, and determining the potential field gradient generated by each combat group in the grouping result; improving a particle updating formula in a particle swarm optimization algorithm by using the potential field gradient generated by the combat groups; solving a target function to determine a maneuvering decision of a friendly aircraft. The improved particle swarm optimization algorithm is obtained by improving the particle updating formula in the particle swarm optimization algorithm by using the potential field gradient generated by all the combat groups in the grouping result. The application can improve the coping ability, autonomous decision-making ability and coping ability in a complex battlefield environment of the beyond-visual-range air combat.

The utility model provides a gliding guidance kit based on an unmanned aerial vehicle. The gliding guidance kit comprises a shell, a reducing sleeve, a gliding wing, an inertial navigation assembly, a flight control system, a battery and a rudder tail, the reducing sleeve is installed at the front end of the shell, the rudder tail is installed at the tail end of the shell, the hang glider, the inertial navigation assembly, the flight control system and the battery are all installed in the shell, and the inertial navigation assembly and the flight control system are both electrically connected with the battery; under the air-drop scene, the lift-drag ratio is increased by unfolding the large hang glider of the missile, the air-drop strike distance is increased, meanwhile, the deviation rectifying capacity is achieved, and the precise strike effect is achieved.

The utility model discloses a single foot twist break type transducing element. Including electrode plug, shell, fixed glue, solder and bridge silk. The utility model has designed single foot twist break type initiating explosive device structure in the country first, is a novel structure's transducing element for missile gyro, has higher firing reliability, twist break reliability and security, can be applicable to various gyro ignition, provides new train of thought for gyro design and structure optimization. Product adopts single foot structure electrode plug, double bridge parallel firing design, constitutes firing unit after filling detonating agent, integrates firing and working function, realizes product 700ma firing, 35MPa output pressure bearing, foot line 3.92N the twist break of cm, 59kPa vacuum ignition, damp heat, medium withstand voltage, electromagnetic compatibility, firing reliability 0.999 etc. technical index requirement. The utility model product reliability, security, versatility are high, and environmental adaptability is strong, and the technology reaches the leading level in the country, and the application prospect is wide, and military, economic benefit is remarkable.

The application provides an internal embedded bomb cabin honeycomb configuration aerodynamic layout for improving flow field stability, and the internal embedded bomb cabin cross section body is a hexagonal honeycomb configuration layout, wherein the top surface of the honeycomb configuration is opened, the other five surfaces are closed wall surfaces, the top surface is parallel to the bottom surface, the included angle alpha between the bottom surface and the adjacent surface is 135 degrees, the longitudinal depth ratio beta of the honeycomb configuration is L / H=1.5, L is the maximum transverse distance of the bomb cabin cross section, and H is the maximum longitudinal height of the bomb cabin cross section. Under the condition of meeting the requirements of structural strength, sound and heat insulation, weight reduction, internal structure design of the fuselage, capacity and other constraints, the internal flow flow field stability of the internal embedded bomb cabin can be greatly improved, the flow field characteristics are improved from the physical root, the evolution of fine vortex and secondary vortex is effectively inhibited, and the process of the flow field from the steady state to the unsteady state is greatly delayed. Because the optimization design effectively inhibits the unsteady flow, especially the premature appearance of the turbulent flow, the aerodynamic efficiency of the internal flow of the bomb cabin is improved.

The invention provides a fuse monorail type lead-free reflow soldering machine, and relates to the technical field of reflow soldering. A fuze monorail type lead-free reflow soldering machine comprises a reflow soldering cavity, a floor cabinet, a servo motor and a soldering product and further comprises a driving piece, a conveying rail, a product base and a tensioning piece, and the driving piece is arranged at the top of the floor cabinet and used for providing autorotation driving force; the conveying rail is arranged on the driving part and is used for continuously providing a conveying position; the product seat is arranged on the transportation track and is used for providing annular limitation for welding products; the tensioning piece is arranged on the lower portion of the conveying rail and used for tensioning the conveying rail and gathering hot air flow to the welding products. And through the arrangement of the conveying rail and the product base, annular limitation can be provided for the missile fuze product, the stability of the missile fuze product is guaranteed, and the reflow soldering quality is improved.