898results 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 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 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 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 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 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.

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 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.

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.

The invention provides a hybrid power system for a dual-mode patrolling bomb and a use mode of the hybrid power system, and belongs to the technical field of aviation power. The problem that an existing patrolling bomb power system is difficult to meet the requirements for long-time cruising and instantaneous explosion acceleration at the same time is solved. The system comprises a hydrogen supply mechanism, a hydrogen peroxide supply mechanism, a fuel cell, a motor and a combustion chamber, the hydrogen supply mechanism is connected with an anode of the fuel cell, the fuel cell is connected with the motor, the motor is connected with a propeller, and the hydrogen peroxide supply mechanism is provided with two supply lines. The first supply line is connected with a cathode of the fuel cell through the hydrogen peroxide decomposition unit, the second supply line is connected with the combustion chamber, a first flow regulating valve is arranged on the first supply line, a second flow regulating valve is arranged on the second supply line, a grain is arranged in the combustion chamber, and the grain is spontaneous combustion fuel with a metal framework. The combustion chamber is connected with the exhaust nozzle. The hybrid power device is mainly used for providing hybrid power for the dual-mode patrol ammunition.

An airframe for a projectile having a cylinder having a bottom spring and a piston within the cylinder is provided. The cylinder has a locking assembly having a housing. A leaf spring is in the housing and coupled thereto. The housing has a limiter where the leaf spring operatively couples with the limiter and the limiter biases the leaf spring. The piston is moveable relative to the cylinder between a retracted configuration and expanded configurations. The piston has a housing where the piston moves between the retracted configuration and the expanded configurations when the piston housing receives a fluid. The piston has a groove that receives the leaf spring. The groove can align with the leaf spring in one expanded configuration where the piston groove receives the leaf spring thereby locking the piston with cylinder. In another expanded configuration, the leaf spring is outside the piston groove.

A guided munition and associated method, which may be implemented by a computer program product, is provided. The guided munition may have an annular folded optic (AFO), wherein at least a first portion of the AFO may be, at least partially, on a wing of the guided munition. The AFO may be transversely segmented to enable the AFO to move between a collapsed position and a deployed position. There may be a first segment and a second segment that selectively interface in response to at least one instruction to move the first segment and the second segment closer to each other. There may be an interface plane, wherein the interface lies along the interface plane, that intersects a central axis at a location outside the body of the AFO. This allows for a dual mode seeker with both the SAL and an image sensor located behind a warhead.

The invention provides a solid sounding rocket landing point control guidance method which comprises the following steps: calculating a geocentric angle between a rocket depletion point and a target landing point, the speed of the rocket depletion point and the to-be-increased speed of the target landing point of the rocket according to current real-time parameters of the rocket; taking the target drop point of the rocket as a constraint, and calculating a theoretical program angle of the rocket according to the depletion point speed of the rocket and the to-be-accelerated speed of the target drop point of the rocket; the actual pneumatic influence and the earth rotation influence are considered, the theoretical program angle of the rocket is corrected, and the actual program angle of the rocket is obtained; and the actual program angle is issued, and guidance calculation is completed. According to the method, the theoretical program angle is iteratively calculated in real time on the rocket computer to complete flight control of the rocket, it is guaranteed that the drop point precision meets the requirement, the problems that a traditional perturbation guidance method is large in calculation amount and poor in adaptability are solved, and high-precision and low-calculation-complexity real-time guidance is achieved.

The invention discloses a modular foldable grid rudder device driven by a single driving link and applied to attitude adjustment of a cylindrical launching navigation body, and relates to the technical field of underwater launching navigation bodies. The problems that an existing grid rudder structure is complex in control, the reliability of the variant process is poor, a traditional mechanism is prone to damaging the power appearance, and the existing grid rudder structure cannot be applied to a cylindrical navigation body are solved. Base assembly spaces vertically communicated with an inner cavity of a mounting base are arranged on the lower portions of grid rudder storage spaces, each grid rudder storage space is provided with a grid rudder folding module, grid rudders of the grid rudder folding modules are matched with the grid rudder storage spaces in shape, and folding and unfolding and variable attack angle driving mechanisms are embedded in the base assembly spaces. The folding, unfolding and variable attack angle driving mechanism is movably connected with the grid rudder so as to achieve the functions of folding, unfolding and variable attack angle of the grid rudder. The grid rudder follows the modular design principle and can serve as a middle section module to be mechanically and electrically connected with a navigation body to bear the action execution function.

The invention relates to a preform of an aircraft bulkhead and a manufacturing method thereof. According to the invention, the preform of the bulkhead has edge strips located at both ends thereof and a web located in the middle of the preform and connecting the edge strips, the web has a radian in the longitudinal direction of the web, and a 0-degree yarn extending at an angle of 0 degree relative to the longitudinal direction of the web without wrinkles is laid on the web. According to the invention, the 0-degree yarn can be arranged on the web plate of the bulkhead preform in the longitudinal direction without wrinkles, so that the 0-degree yarn has tension in the longitudinal direction of the web plate, thereby meeting the stress requirement on the bulkhead in the longitudinal direction, ensuring the smoothness of preform fibers, improving the adaptation degree of the bulkhead preform to a curved surface, and improving the production efficiency of the bulkhead preform. The manufacturing efficiency of the bulkhead preformed body is improved, and the manufacturing quality and reliability are ensured.

Composite projectiles include various material compositions, diameters, and cavities within the projectiles having a variety of cavity diameters, sidewalls, and bottoms selected and formed to induce different levels of penetration and disintegration of the composite projectiles upon impact with targets.

The invention provides a self-adaptive control system of a small recoverable rocket, the rocket comprises a rocket body, a parachute opening mechanism, a rocket wing mechanism and a power mechanism, and each module realizes modular quick disassembly and replacement design to meet different launching requirements; the rocket body is made of a light high-strength material, so that the internal structure of the rocket is protected; the parachute opening mechanism is unfolded in the rocket returning stage, and the falling speed is reduced; the rocket wing mechanism is used for adjusting the rolling angle and the pitch angle of the rocket; the flight control system is used for instruction receiving, data processing and control signal output; the power mechanism provides power for the rocket.

A projectile capable of coupling or otherwise attaching itself to a target. The projectile includes a coupling mechanism that enables the projectile to be coupled to the target. The projectile can include a variety of electronic components including a processor and a data communications component that enables the projectile to transmit data to a remote device.

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.

An actively-cooled heat shield system includes a heat shield, a tank, a pump, a heat exchanger, and a turbine. The heat shield defines a windward side of a vehicle. The tank stores a coolant. The pump receives the coolant from the tank and outputs a pressurized coolant. The heat exchanger is integrally connected with the heat shield. The heat exchanger receives the pressurized coolant from the pump, transfers heat from the heat shield to the pressurized coolant to generate a heated fluid, and outputs the heated fluid. The turbine includes an inlet, a shaft, and an outlet. The inlet receives the heated fluid output from the heat exchanger. The shaft is coupled to the pump and includes turbine blades. The shaft rotates and powers the pump when the heated fluid received from the heat exchanger acts on the turbine blades. The outlet outputs the heated fluid.

This utility model relates to an explosive bolt damping and shock absorption assembly, which has an internal accommodating space for installing explosive bolts. The explosive bolt damping and shock absorption assembly includes: a main body; and a pad, which is fixed to a first side of the main body by adhesive. The explosive bolt damping and shock absorption assembly further includes: a second baffle, which is fixed to a second side of the main body by adhesive; a first sponge rubber plate located on the side of the second baffle opposite to the main body, and the first sponge rubber plate is fixed to the second baffle by adhesive; and a first baffle located on the side of the first sponge rubber plate opposite to the second baffle, and the first baffle is fixed to the first sponge rubber plate by adhesive. The main body is made of honeycomb panel, and a second sponge rubber plate is provided on the inner wall of the accommodating space. This damping and shock absorption assembly has a simple structure, is easy to process and manufacture, and can effectively solve the problems of difficult installation and low reliability of damping and shock absorption assemblies in actual use.