1033results about "Self-propelled projectiles" patented technology

An apparatus includes a structure configured to receive thermal energy and to reject the thermal energy into an external environment. The structure includes a lid and a body. The structure also includes (i) multiple inline and interconnected thermomechanical regions and (ii) one or more oscillating heat pipes embedded in at least some of the thermomechanical regions. Different portions of at least one of the lid and the body form the thermomechanical regions. The one or more oscillating heat pipes are configured to transfer the thermal energy between different ones of the thermomechanical regions. At least one of the thermomechanical regions includes one or more shape-memory materials configured to cause a shape of the structure to change. Each of the one or more oscillating heat pipes includes at least one channel in the structure.

The invention provides a high-precision optical pose measurement method based on tower recovery, which adopts a close-range bidirectional optical cooperation mode, measures a mark point of a recovery tower through an optical detector on a rocket, and determines the distance and the pose of the rocket relative to the recovery tower; and the optical measurement system of the recovery tower synchronously detects and identifies the rocket and controls a capture arm and a buffer mechanism of the recovery tower to move, and optical measurement data of the recovery tower are fed back to the rocket and are integrated with optical measurement data on the rocket to control landing of the rocket. According to the invention, optical measurement and navigation are introduced into rocket tower type recovery, and the method has the advantages of high measurement precision, high data frequency, interference resistance and the like. The relative relation between the rocket and the recovery tower can be directly obtained through optical pose measurement, compared with inertial navigation, the coordinate conversion process can be reduced, and the advantage of no accumulated error is achieved. Compared with satellite navigation, optical pose measurement does not depend on external signals, and higher independence and electromagnetic resistance are achieved.

The invention provides a rocket power anchor. The rocket power anchor comprises a bullet of a spiral structure. The anchoring structure is detachably connected with the warhead, and the anchoring structure can stretch; the second-stage engine is detachably connected with the anchoring structure, a second-stage delay igniter capable of driving the anchoring structure to stretch is arranged at the joint of the second-stage engine and the anchoring structure, and the second-stage engine is filled with a second-stage propelling grain; the first-stage engine is detachably connected with the second-stage engine, a middle partition bin is arranged at the joint of the second-stage engine and the first-stage engine, a first-stage delay igniter is arranged on the middle partition bin, and the first-stage engine is filled with a first-stage propelling grain; the exhaust nozzle can be obliquely arranged at a certain angle, so that the rocket power anchor can rotate during movement; the lifting ring is rotationally connected to the first-stage engine; the total mass of the first-stage propelling grain is smaller than that of the second-stage propelling grain; and a pressure-bearing diaphragm is arranged on the middle partition bin.

The invention discloses an aerodynamic configuration structure of a precision guidance assembly and a design method, and belongs to the technical field of precision guidance, an assembly precursor adopts continuous three-section curve design: a supersonic speed section adopts a parabola profile, is provided with ribs on the surface, and is used for optimizing shock wave control under a supersonic speed working condition; the transonic speed section adopts a hyper-ellipsoid curved surface, ribs are arranged on the surface of the transonic speed section, and metamaterials are attached to the ribs and used for restraining transonic speed separation; the subsonic speed section adopts arc transition, and ribs are arranged on the surface and used for improving the lift force efficiency of the control surface; wherein the curvatures at the joints of the three sections of curves are continuous. According to the method, supersonic velocity-transonic velocity-subsonic velocity three-section collaborative design is realized, and meanwhile, a boundary layer is improved from a microscopic level by combining a collaborative drag reduction design of a microscopic rib and metamaterial technology, and the performance of a guidance assembly is optimized and improved.

The present invention relates to projectiles and munitions, and more specifically to such in flight. More particularly the present invention relates to projectiles and munitions in flight equipped with one or more image sensors adapted for acquiring image data of the environment surrounding the projection or munition in flight. The present invention further relates to systems and methods for correcting or stabilizing motion effects and artifacts present in the image data related to the movement or motion of the projectile or munition in flight, including spin or rotation of the projectile or munition.

The invention discloses a heat-proof skirt which is of a single-layer or multi-layer structure, each single-layer material is formed by compositing quartz fiber cloth and coating when multiple layers are arranged, the heat-proof skirt is made of 6-8 single-layer materials, and the skirt edge is in a wave shape. The coating is prepared from the following raw materials in parts by weight: 4 to 6 parts of vinyl polysiloxane, 15 to 20 parts of hydrogen-containing silicone oil, 15 to 20 parts of ammonium polyphosphate, 10 to 15 parts of aluminum hydroxide, 5 to 8 parts of silicon dioxide, 0.005 to 0.02 part of a platinum catalyst and 30 to 40 parts of a solvent. The heat-proof skirt has contractility, and the influence on normal swinging resistance of the engine is reduced to the maximum extent.

The invention provides an electric pump pressure type liquid rocket engine and a carrier rocket. The electric pump pressure type liquid rocket engine comprises a gas-driven generator system, an electric pump pressure type liquid rocket engine body and an energy management system, and the gas-driven generator system is electrically connected with the energy management system; and the electric pump pressure type liquid rocket engine is electrically connected with the energy management system. According to the technical scheme, the rocket is prevented from carrying a large-mass battery through the generator system, and the highest-efficiency operation of the generator under all working conditions is realized through the energy management system. According to the technical scheme, the energy loss is small, the structural mass is light, the development and application cost of the carrier rocket is reduced, and the reliability of the carrier rocket is improved.

The invention provides a carrier rocket engine stage-switching motor redundancy control system, which is combined with a multi-input instruction redundancy voting method, a voting path software and hardware arbitration method and the like to complete high-reliability carrier rocket engine stage-switching motor control. According to the system, two-out-of-three voting is carried out by using a triple-redundancy input control instruction and interactive heartbeat, software and hardware joint arbitration is carried out on voting results of three modules, and an instruction control channel is determined to complete phase sequence control of the stage conversion winding. The method has strong fault capability, has no single-point failure link, and can ensure the normal operation of the engine conversion motor in the flight process of the carrier rocket.

The invention discloses a multi-missile self-learning differential game cooperative guidance method based on disturbance feed-forward compensation, and the method comprises the steps: providing a finite-time convergence sliding-mode nonlinear expansion state observer for solving the problems of unknown partial parameters and external disturbance in a multi-missile cooperative guidance system; a sliding mode control method is combined with an extended state observer, a novel fixed-time convergence sliding mode surface and a sliding mode reaching law are constructed, and real-time estimation and efficient compensation of unknown disturbance in a system are achieved. Strictly proving the fixed convergence time of the observer by using a Lyapunov function, and analyzing a disturbance estimation error; compared with a traditional ESO, the method has the advantages that the speed and precision of state tracking are remarkably improved, and the anti-interference capability of the multi-missile cluster in a complex interference environment is effectively improved.

The invention provides a missile guidance method based on hierarchical reinforcement learning, and belongs to the field of control strategies. The method comprises the following steps: firstly, introducing a missile-target kinematics model and a missile velocity attenuation model; secondly, designing a process reward and a result reward on the basis of the missile-target kinematics model to obtain a total reward function, and guiding the missile to follow the target; thirdly, after reward function design is completed, a flexible action-evaluation algorithm SAC is used for conducting preliminary training on the missile under the condition that the target adopts linear motion, and an initial guidance strategy is obtained; and finally, on the basis of the initial guidance strategy, the enemy plane adopts an evasion strategy, a flexible action-evaluation algorithm SAC is used for continuing to train the missile, and finally a high-quality guidance strategy is obtained. According to the method, the missile can autonomously adjust the flight path in the complex battle condition, interference and challenge in the complex environment are overcome, and therefore the guidance precision and the attack success rate are greatly improved, and the missile striking effect and survivability are improved.

The utility model relates to an integrated light-weight launch vehicle tail section and launch vehicle, the integrated light-weight launch vehicle tail section is formed by welding two semicircular wall plates, the wall plates comprise a first arc plate and second arc plates located on the two sides of the first arc plate, the first arc plate and the second arc plates are both aluminum honeycomb plates, and the first arc plate and the second arc plates are arranged in parallel. The first arc plate comprises a first arc part and second arc parts located on the two sides of the first arc part, the first arc part and the second arc parts are tangent and opposite in direction, and the circle center of the second arc parts coincides with the circle center of the second arc plate. According to the utility model, the technical problem that the development requirements of light weight, modularization and low cost cannot be met in the prior art can be solved.

The invention discloses liquid rocket system polarity test equipment and a test method, and the method comprises the steps: carrying out the analysis of the working principles of a rocket system polarity test servo system, a final correction attitude control correction system and an engine power system, selecting a corresponding data collection device, and installing the data collection device at a to-be-detected position. During testing, a corresponding data acquisition device is used for carrying out data acquisition on parts with relative position changes or electromagnetic valve actions in the servo system, the final correction attitude control correction system and the engine power system. The data transmission system transmits the collected test data to the control and data processing system. And the control and data processing system compares the received test data with the instruction data to realize automatic interpretation of the system polarity test. The liquid rocket polarity test data and the system data are effectively and deeply fused, the accurate closed loop of the data is realized, the requirements of systematization and automation of carrier rocket test are met, the product reliability is improved, and the production efficiency is also improved.

The invention provides a carrier rocket ground test system and method supporting visual process configuration. The system mainly comprises: (1) a visual process configuration module, which configures a test process file through a graphical interface; (2) a process analysis and scheduling module for analyzing a process file and converting the process file into a task sequence; the test execution engine module (3) is used for driving bottom hardware and executing a test instruction sequence; (4) the test resource library module is used for storing and managing various information of test elements and various information of test equipment; the execution monitoring and data management module (5) is used for displaying a test execution state in real time and recording whole-process data; according to the method, the test logic and the execution program are separated, and a user can quickly construct, modify and reuse an automatic test process in a graphical mode, so that the test flexibility is greatly improved, and the test adaptability and efficiency are remarkably improved.

The embodiment of the invention provides a rocket vertical recovery method and system and a rocket, and the method comprises the steps: after a first-stage rocket is separated, the thrust of a central engine of the first-stage rocket is adjusted to a preset proportion of rated thrust so as to provide positive overload for a propellant in a storage tank; the propellant can be gathered to an outlet in the bottom of the storage tank through positive overload; the flying direction of the first-stage rocket deviates from the current flying direction by performing bidirectional swinging on the central engine, and the posture of downward diving of the first-stage rocket towards the earth direction is kept; tangential force in two different directions tangent to the current flight direction of the first-stage rocket can be generated at the same time by swinging the center engine in two directions, the posture of the first-stage rocket is adjusted to the upward flight posture from the downward diving posture, and speed reduction is conducted till the first-stage rocket lands on the earth. The thrust of the center engine of the first-stage rocket is adjusted to be small, and attitude control and propellant management of first-stage rocket recovery are achieved.

The invention of system of observation and monitoring the earth warming and changes in polar ice behavior using LoRa to satellite technology is a system for measuring, observing and monitoring weather and environmental conditions in order to determine the effects of environmental conditions and climate changes on the way ice masses change in the polar regions of the earth. This invention is related to the rocket equipped with several modules for measuring the conditions of the surrounding environment as well as data transmission through the LoraWan protocol to a fixed or deployable station. Also, in this invention, if there is a need to calibrate any of the modules in the meteorological rocket, the central server structure, which is equipped with an artificial intelligence analyzer, can calibrate the device by comparing the arrays received from the surrounding rockets and analyzing the data received from these systems.

The invention discloses an artificial influence weather rocket projectile three-explosion self-destruction body assembly detection system, and relates to the technical field of multivariable comprehensive measurement, and the system comprises a multi-dimensional positioning measurement module which is used for measuring the coaxiality and fitting pressure of a self-destruction body and an engine housing; the multi-parameter measuring module is filled with the medicament to measure the weight, density and internal defects of the medicament; the connection strength dynamic measurement module is used for measuring torque, tensile strength and vibration displacement; the igniter performance parameter measuring module is used for measuring resistance, time delay and safety reliability; the measurement data integration analysis module is used for integrating a data generation report and an encryption log; and the system control and exception handling module is used for coordinating module time sequences, monitoring data and handling exceptions. According to the system, multi-parameter integrated measurement of the three-explosion self-destruction body is realized, the measurement precision of positioning, medicament filling, connection strength and igniter performance is improved, the remain after self-destruction is ensured to be compliant, and the data is traceable.

The invention relates to the technical field of functional materials, in particular to phenolic aldehyde hybrid SiO2 mesoporous microspheres and application thereof. Mesopores with the pore diameter of 2-50 nm are formed in the surface and the interior of the phenolic aldehyde hybridized SiO2 mesoporous microsphere; the average diameter of the phenolic aldehyde hybrid SiO2 mesoporous microspheres is 10 to 100 microns; the phenolic hybrid SiO2 mesoporous microspheres can be applied to the fields of thermal protection, wave absorption, heat insulation and coatings. According to the phenolic hybrid SiO2 mesoporous microsphere disclosed by the invention, by virtue of a unique mesoporous structure and an organic-inorganic hybrid characteristic, excellent heat insulation performance and multifunctionality are realized, and meanwhile, the problems of insufficient protection performance, single function, poor processability, high cost and the like of an existing material are solved.

The invention provides an artificial influence weather operation equipment ammunition locking device and system and a use method. The device comprises a rocket projectile, a launching controller, an unlocking controller and a monitoring terminal, a shell of the rocket projectile is provided with a projectile body code, a lock control chip, a storage and an ignition element are arranged in the rocket projectile, and a UID code and a launching code corresponding to the projectile body code are stored in the storage; the launching controller is connected with the unlocking controller, the unlocking controller is connected with a lock control chip of the rocket projectile, and the monitoring terminal is connected with the unlocking controller and used for sending a UID reading instruction and an unlocking instruction with a launching code to the unlocking controller. By adopting the locking device provided by the invention, the safety of the artificial influence weather operation equipment ammunition in the transportation and use process can be greatly improved.

A payload module for an aerial vehicle stores and deploys one or more submunition assemblies, which provide a portion of the vehicle's OML when stored, in a manner that preserves the aerodynamics of the aerial vehicle post-deployment. Pre-deployment a skin is aligned to a strongback such that the submunition assemblies are exposed through openings in the skin. Post-deployment, the skin rotates to cover any openings in the strongback exposed by the deployment of the submunition assemblies to restore the OML.

The invention discloses a grid rudder folding and unfolding mechanism, a rocket interstage section and a rocket, and relates to the technical field of grid rudders.The grid rudder folding and unfolding mechanism comprises a rudder shaft, a grid rudder, a driving cylinder, a piston, a piston rod and a connecting assembly. According to the technical scheme, the piston rod is driven to move through sliding of the piston in the driving cylinder, and then unfolding and folding actions of the grid rudder are achieved through the connecting assembly. Meanwhile, the piston rod is matched with a key groove in the folding state of the driving cylinder, the integrated design of folding and locking can be achieved, and unfolding and locking are achieved through air pressure after unfolding. Through the arrangement mode, mechanism integration is achieved, two sets of locking mechanisms are reduced, meanwhile, use of traditional servo mechanisms, torsional springs and other elastic elements is reduced, cost is reduced, reliability is improved, meanwhile, enough driving force can be provided, and it is ensured that the grid rudder can be unfolded and folded rapidly and reliably.

The invention discloses a method for analyzing the testability of a marine launched rocket by light thunder equipment, which comprehensively considers factors such as radar performance indexes, tracking target characteristics and the like through radar tracking capability calculation, tracking angle calculation, visibility constraint analysis, testability constraint analysis and testability time determination. And a quantitative analysis result is provided for the testability of the sea-based light thunder equipment of the rocket launched on the sea, and the accuracy of a calculation result is high. The application of the method provided by the invention can assist the analysis of measurement and control support conditions during maritime emission, especially the development of maritime optical and radar equipment station distribution design and the verification of the rationality of a maritime station distribution method, and also facilitates the optimization of a scheme which meets the key measurement and control requirements and has a better overall measurement effect through the testability analysis of different station distribution schemes.

The invention relates to the technical field of aerospace, and discloses a multi-engine mounting cabin section and a carrier rocket. The multi-engine installation cabin section comprises a cabin section and a force transmission mechanism, the force transmission mechanism is located in a cabin section body and connected with a front end frame and a rear end frame of the cabin section, and thrust generated by a center engine is transmitted to a force transmission ring through a main cross beam and finally transmitted to the front end frame of the cabin section through a radial beam and a first force transmission beam. Thrust generated by the peripheral engine is transmitted to the radial beam, the rear end frame and the second force transmission beam through the outer cross beam and finally transmitted to the front end frame of the cabin section, the force transmission process is uniform in force transmission, high in efficiency and good in stability, and the carrying efficiency of the carrier rocket can be improved. The invention further provides a carrier rocket which comprises a rocket body and a plurality of engine installation cabin sections installed on the rocket body, and the rocket carrying requirement is met.

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.

The invention relates to the technical field of rocket guidance, and discloses a recoverable rocket power soft landing two-stage control method, system and medium, the method comprises the following steps: establishing a dynamical model of a first stage, introducing constraint conditions of power descent landing, controlling the rocket to reach the position of # imgabs0 # right above a landing point, and controlling the position of # imgabs0 # right above the landing point; establishing an optimization problem P0 by considering the fuel minimization of the recoverable rocket; converting a non-convex constraint in the optimization problem P0 into a convex constraint, and establishing an optimization problem P1; variable substitution is introduced into the P1, and a convex optimization problem P2 of the first stage is obtained; controlling the recoverable rocket to be vertical to a landing point according to the displacement and speed constraint of the first stage, and establishing a convex optimization problem P3 by considering the minimum fuel of the recoverable rocket; according to the method, the dynamic soft landing process of the rocket is divided into two stages, the constraint condition of the attitude of the rocket terminal is considered, it is guaranteed that the attitude of the rocket during landing is in a vertical or approximately vertical state, and therefore the reliability of vertical landing is improved.

The invention provides a power measurement and control system and method for a liquid carrier rocket, and relates to the field. Comprising a to-be-tested real equipment system used for completing a rocket body test; the simulator is used for installing a simulation equipment system, and simulation equipment in the simulation equipment system is in one-to-one correspondence with real equipment in the real equipment system; the control machine is in communication connection with the real equipment system and / or the simulator and is used for issuing a target instruction to the real equipment system and / or the simulator in the rocket body testing process; the front-end machine is in communication connection with the control machine and the simulator and is used for forwarding first feedback information after the simulator executes the target instruction to generate the first feedback information; the monitoring machine is in communication connection with the front-end machine, the real equipment system and the simulator and is used for receiving first feedback information and / or second feedback information in the rocket body testing process and monitoring the first feedback information and / or the second feedback information; the second feedback information is feedback information generated after the real equipment system executes the target instruction.

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.

A dual-mode roll / nod gimballed seeker is configured to provide an unobstructed path for a forming-warhead penetrator. A nod assembly includes first and second seekers separately offset from the roll axis whose FOV overlap within a stand-off range to a target. A roll gimbal rotates the dome and nod assembly around the roll axis. A nod gimbal rotates the nod assembly about a nod axis over a nod FOR that causes the seekers to cross the roll axis and obstruct a path along the roll axis. At terminal, the nod gimbal rotates the nod assembly to a nominal nod angle at which the path is unobstructed to detonate the forming warhead and project a penetrator along the path through the dome. This alleviates the need for a precursor charge and saves precious volume in the airframe, which is of particular import for small diameter warheads, in particular those having a diameter of no more than 8 inches.

The invention discloses a spoiler device of a high-rotation tail control electric steering engine. The device comprises a shell used for being fixed to the tail of a projectile body; the executing mechanism comprises a plurality of spoilers, and the spoilers are symmetrically distributed on the outer side of the shell along the axis; the motor is fixedly arranged on the shell; the gear assembly is arranged between the motor and the connecting rod assembly and used for converting rotation output of the motor into displacement change of the connecting rod assembly; the connecting rod assembly is connected to the spoilers and used for responding to driving of the gear assembly to adjust the relative positions of the spoilers in the shell of the projectile body. The technical problem that in the prior art, when the flight direction and state need to be controlled in the flight process, the tail shape of the projectile body cannot be changed is solved.

The invention belongs to the technical field of rocket unlocking and separating, and particularly relates to an unlocking and separating device, a system thereof and a carrier rocket. The device comprises a plurality of unlocking and separating units. Each unlocking and separating unit comprises an air supply assembly, a controllable switch assembly and a pneumatic unlocking and separating assembly. The controllable switch assembly comprises a plurality of sub-controllable switches which are connected in parallel, the pneumatic unlocking and separating assembly comprises a plurality of unlocking and separating structures which are connected in parallel, and an air outlet of the air supply assembly communicates with an air inlet of each unlocking and separating structure through each sub-controllable switch; the unlocking and separating structures of every two adjacent unlocking and separating units are connected through a device to be separated. Therefore, the controllable switch assembly and the pneumatic unlocking separation assembly both adopt redundancy design. If a certain sub-controllable switch or an unlocking separation structure breaks down, the other parallel connection parts can still work normally, it is ensured that the overall function of the unlocking separation device is not affected, and therefore the failure rate in the unlocking separation process can be reduced, and the safety coefficient and reliability of the unlocking separation device are improved.

The application discloses a laser beam riding guidance missile roll angle de-rotation measurement system and method based on polarized light, and belongs to the technical field of missile guidance control. The system comprises a laser irradiator for emitting polarized laser to generate a polarized laser field; a polarized detector assembly is arranged on one side of a target missile body close to the laser irradiator and is located in the polarized laser field together with the target missile body; wherein the polarized detector assembly comprises a four-quadrant detector, an accelerometer, a detection circuit and a signal processing module; a rectangular coordinate system is established in the polarized laser field, and a target missile body coordinate system is established. The application can not only measure the real-time roll angle parameter of the missile relative to the laser information field coordinate system, but also saves the loading process of the missile body reference information, simplifies the system complexity and is highly applicable.