1807results about "Self-propelled projectiles" patented technology

Various embodiments include an actuated projectile launcher, including: a base structure; a projectile support, and a holding area from which a projectile launches; propulsion wheels selectively driven to rotate; and support arms, each supporting at a first end thereof a separate one of the propulsion wheels and being supported at an opposed second end thereof by the base structure. The support arms are configured to move the propulsion wheels between an expanded position and a contracted position, in which the propulsion wheels are closer to one another in the contracted position as compared to the expanded position. The movement of the propulsion wheels toward the contracted position brings the propulsion wheels into engagement with the projectile and drives the projectile from the holding area.

The invention discloses an online efficient load shedding control method based on a forecast wind field. The method comprises the following steps: firstly, acquiring a load attack angle and a sideslip angle which are calculated in real time on a rocket; amplitude limiting processing is carried out on the load attack angle and the sideslip angle, an attack angle control instruction is generated after gain control network calculation is carried out, and an attitude control system carries out direct attack angle control in a load sensitive area in combination with the load shedding requirement and the guidance deviation correction capacity; and pitching and yaw feedback control instructions are generated, active load shedding control is implemented, and errors generated by attack angle control are compensated. The on-line active and passive load shedding mode covers the real deviation or interference influence during flight, the generated load shedding instruction is better matched with the actual flight, the universality is good, the implementation is easy, and the load shedding effect is better.

The present disclosure discloses a method for designing a terminal guidance law based on deep reinforcement learning, and relates to the field of missile and rocket guidance. The method includes: establishing a relative kinematics equation between a missile and a target in a longitudinal plane of a target interception terminal guidance section of the missile; to adapt to a research paradigm of reinforcement learning, abstracting a research problem and modeling as a Markov decision process; building an algorithm network, and setting algorithm parameters, where a selected deep reinforcement learning algorithm is a deep Q-network (DQN); in a terminal guidance process of each round, obtaining a sufficient number of training samples through Q-learning, training a neural network and updating a target network at fixed frequencies respectively, and continuously repeating the above process before set learning rounds are reached.

The invention relates to the field of intelligent control and software management, and discloses a carrier rocket software control method and decision making device based on SAC, and the method comprises the following steps: obtaining the state information of a carrier rocket software system, and constructing a unified expression; an optimal action is generated based on an SAC algorithm; executing the action and collecting feedback information after the action is executed; utilizing feedback information to optimize an actor network and a commentator network in the SAC model, and continuously updating a decision strategy; based on the optimized strategy, the task priority, the running state of the software module and the abnormal recovery operation are dynamically adjusted, and the system state is collected again to optimize the model in a closed-loop mode. The system operation state can be adjusted in real time, the task scheduling efficiency, the resource utilization rate and the exception handling capacity are improved, intelligent management of a carrier rocket software system is achieved, and the method is suitable for complex task environments and high-dynamic operation scenes.

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.

A separation device assembly with efficiency that allows for multi-fault-tolerance. The separation device assembly comprising an inflation device that applies force to a shear plate assembly which in-turn applies focused force to a frangible portion. The assembly further comprising a set of upper and lower compressive load bearing elements which are configured as a stopping means. The limitation of excess movement after fracture allows residual energy from the inflation device to be applied to non-fractured portions of the assembly.

The invention provides a carrier rocket box bottom and a forming method thereof. The forming method comprises the following steps: S1, calculating box bottom stress distribution under the action of internal pressure by adopting a film stress method and a moment-free theory; s2, determining a latitudinal tension-compression stress boundary according to box bottom stress distribution under the action of internal pressure; s3, the top is subjected to forming machining, and the bottom is subjected to forming machining; and S4, the machined top and the machined bottom are subjected to tailor welding according to the latitudinal tension and compression stress boundary, and therefore the finally-formed box bottom is obtained. According to the carrier rocket box bottom forming method, the problems that under the action of internal pressure, the bottom of the box bottom is prone to wrinkling and elastic buckling, and the top of the box bottom is prone to cracking and instability are solved.

A vacuum insulation layer is wrapped around the length of a warhead to thermally insulate the warhead from fire or aerodynamic heating. The vacuum insulation layer may be integrally formed into the warhead casing or provided as a sleeve that may be permanently or removably positioned about the warhead casing. The vacuum insulation layer is held under vacuum with a pressure of less than 25 Torr and a thermal conductivity Tcond_vac of less than one-third of the thermal conductivity of air Tcond_air.

A guidance system for a guided munition has an inertial measurement unit (IMU) or another type of first sensor on the guided munition, electro-optical / infrared (EO / IR) sensor on the guided munition, and a guidance computer assembly (GCA) having a Programmable Real-Time Unit Industrial Communication SubSystem (PRU-ICSS), wherein the PRU-ICSS is in operative communication with the IMU or another type of first sensor and the EO / IR. The PRU-ICSS has a first Programmable Real-Time Unit (PRU), wherein the first PRU is programmed to receive and process input data from the IMU or another type of first sensor on the guided munition, and the PRU-ICSS has a second PRU, wherein the second PRU is programmed to receive and process input data from the EO / IR sensor on the guided munition.

A vacuum insulated blast tube includes an insulating layer of a burn resistant material such as phenolic resin formed on an interior surface of the blast tube to provide the necessary erosion and thermal insulation properties to protect the blast tube and a vacuum insulated layer formed in the walls of the blast tube itself to provide the additional thermal insulation properties to protect any non-propulsive sub-systems positioned in the void space around the blast tube. A void space in the walls of the blast tube is held under vacuum with a pressure of less than 25 Torr and a thermal conductivity Tcond_vac of less than one-third of the thermal conductivity of air Tcond_air to form the vacuum insulation layer.

The invention discloses a structural bearing associated rocket time-phased Q alpha release value design implementation method. The method comprises the following steps: carrying out statistical analysis on actually measured high-altitude wind data aiming at a rocket launching site position to form a high-altitude wind statistical result; carrying out flight standard trajectory design of the rocket to form a standard trajectory data result; developing an attack tilt angle calculation result under a rocket standard trajectory to obtain a Q alpha preliminary design result changing along with flight time; rocket flight load calculation and rocket structure design and optimization are carried out, so that the structure of the rocket meets the carrying capacity; the time-phased Q alpha release value of the rocket launching site is formulated, multiple incremental Q alpha values are obtained on the basis of the preliminary design result of Q alpha, and the corresponding maximum Q alpha release value meeting the maximum bearing bending moment requirement serves as the time-phased Q alpha release value to be output according to the transverse bending moment load corresponding to the Q alpha values; and before emission, emission release evaluation is performed according to the release value of the time-segmented Qalpha. Therefore, the problem of low release probability in a traditional method is solved.

The invention discloses a solid rocket engine composite material skirt connecting structure and a forming method thereof.The connecting structure comprises a composite material skirt, a skirt hole reinforcing ring, a connected cabin section and a connecting piece, the composite material skirt and the skirt hole reinforcing ring are integrally formed, and the rear end of the composite material skirt and the rear end of the skirt hole reinforcing ring are connected with the connected cabin section through the connecting piece; the connected cabin section is of an L-shaped nested inner lap joint structure and has the butt joint limiting and bearing functions, and the thickness transition section at the front end of the composite skirt is connected with a fiber winding structure layer of a combustion chamber shell. Compared with a traditional metal skirt connecting structure, the weight reduction effect is obvious, compared with a traditional composite skirt hinge block connecting structure, the connecting mode is simpler, and the fiber continuity of a composite skirt connecting area is guaranteed; the problems that in-plane shear rigidity and strength of the composite material are insufficient, and orifice positions are prone to damage are solved. The structure and the forming process are simple, the maturity is high, and the cost is low.

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

The invention belongs to the field of rocket recovery, particularly relates to an offshore recovery rocket sealing device and a using method thereof, and provides the following scheme that the offshore recovery rocket sealing device comprises a rocket sealing conveying mechanism movably mounted on an offshore landing platform, and the top end of the outer side wall of the offshore landing platform is rotationally sleeved with a rotating platform with a rotating driving mechanism; a rocket sealing transportation mechanism is movably mounted on one side of the rotating platform through two fixing blocks and a locking bolt, and a weight box is mounted on the other side of the rotating platform; the rocket sealing transportation mechanism comprises a loading cylinder with a sealing gate plate at an opening in one side, and an auxiliary fixing mechanism I with a plurality of clamping mechanisms is mounted in the loading cylinder. According to the device, rockets landing on any landing point on the offshore landing platform can be rapidly clamped and fixed in an auxiliary mode, then the rockets are loaded into the loading cylinder to be transported, the problem that the rockets topple over in the harbor returning process of a rocket recovery platform is solved, and the transportation safety is high.

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.

A separation device assembly with efficiency that allows for multi-fault-tolerance. The separation device assembly comprising an inflation device that applies force to a shear plate assembly which in turn applies focused force to a frangible portion. The assembly further comprising a means for limiting excess movement after fracture so that residual energy from the inflation device is applied to non-fractured portions of the assembly. Wherein the stopping means is integrally formed with a frangible portion.

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 utility model discloses a docking device based on a hoisting type small envelope rocket multi-engine single-machine parallel structure. The docking device is characterized by comprising a crane (1), an inclined steel cable (2), a cross beam (3), a length adjusting device (4), a profiling hoop (6) and a combined hoop (7), the rocket engine (5) is composed of a spray pipe (8), an engine body (9) and a rack (10). According to the butt joint device, a combined type holding ring and profiling hoop structure is designed by means of only two force bearing points on the engine, the structure is simplified, the envelope of the butt joint device of the multi-engine single-machine parallel connection structure is smaller than that of a single engine body, single engine butt joint of the multi-engine single-machine parallel connection structure in a narrow space under complex working conditions is achieved, and the butt joint efficiency is improved. The problem of interference in the butt joint process of all the engines in the multi-engine single-machine parallel structure is solved, meanwhile, safety risks are eliminated, and the butt joint reliability and working efficiency of the multi-engine single-machine parallel structure in a narrow space under complex working conditions are improved.

The invention discloses a fairing pulsating pressure scaling wind tunnel test method capable of reducing space-ground difference. The method comprises the following steps: determining initial parameters; based on the determined initial parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test, arranging pulsating pressure measuring points, and carrying out transonic speed state fairing pulsating pressure environment measurement to obtain a measurement result; performing parameter optimization according to a comparison result of a measurement result and a standard parameter under a real flight test condition to obtain an optimal parameter of manual transition; and based on the optimal parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test to obtain test results of the configuration in different free incoming flow Mach number, attack angle and sideslip angle states. According to the method, the free incoming flow local shock wave-boundary layer interference state and the boundary layer separation state of the ground wind tunnel test and the flight test are ensured to be closer through a manual transition mode, and the ground test prediction precision of the carrier rocket fairing pulsating pressure environment is improved.

The invention relates to a gliding guided projectile trajectory optimization method based on a PSO-hpRPM algorithm, and the method comprises the steps: taking the longitudinal plane motion of a gliding guided projectile as a research target in a launching coordinate system, and building a longitudinal motion equation of the gliding guided projectile; determining a ballistic objective function and constraint conditions, and establishing a ballistic optimization mathematical model; establishing a PSO-hpRPM optimization algorithm: the algorithm adopts a double-ring optimization strategy, an outer ring adopts a PSO algorithm to optimize a firing angle and ignition time, and an optimal initial value is provided for an inner ring; an inner ring adopts an hpRPM algorithm to solve an optimal control law u and a fitness value J, and finally a gliding trajectory with the farthest range is obtained under all constraint conditions; initial conditions are set, iterative calculation is carried out, and when the fitness of an objective function is converged, the optimal firing angle, ignition time and the optimal trajectory are obtained. A new solution is provided for the ballistic optimization problem of the gliding guided projectile, and a theoretical reference is also provided for other ballistic optimization problems of the same type.

A system for asset inventory and resupply for supplying individuals in remote locations is disclosed. The individuals may be people and / or machines. The system and method involve providing multiple individuals in the field with sensors that automatically collect data on the expenditure of their supplies. In some cases, wearable sensors are provided to people in the field. The data is sent to a central depository which tracks their inventory of supplies and anticipates needed supplies, compiles and packages additional supplies so that the additional supplies can be delivered to the individuals in the field. A precision supply delivery system is provided for delivering the supplies.

The invention discloses a segmented self-destruction rain enhancement and hail suppression rocket controlled by a time schedule controller, and relates to the technical field of artificial rain enhancement and hail suppression rockets. The device comprises a catalysis bin assembly, an engine assembly and an empennage assembly, and the catalysis bin assembly, the engine assembly and the empennage assembly are connected through reducing sleeve assemblies. An explosion self-destruction device is adopted as a safe landing device, a segmented self-destruction device is arranged in a rocket body engine and a tail segment assembly in advance, a time schedule controller outputs an ignition signal according to binding information, an empennage assembly, an engine assembly and a catalytic bin assembly are exploded into floccules not larger than 100 g and fall onto the ground, self-destruction of all the assemblies is reliable, and the safety of landing is guaranteed. The rocket body wreckage preventing device can effectively reduce damage of rocket body wreckage to personnel, buildings, farmlands, livestock and the like, any one of the ignition loops of the time schedule controller does not affect performance of other loops during working, fire spurting is avoided, space is saved, and the function is reliable.

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 a navigation and comprehensive electronic depth integrated system for tactical micromissiles. Comprising a comprehensive information processing module (including launch control and flight control solution, combined navigation solution and rudder control solution), an electrical control module (including a secondary power supply, missile-borne power electrical and rudder release), a satellite navigation module, a rigid-flexible hybrid electrical bottom plate, an interface management module and other functional modules, and sensing and executing devices such as an inertial measurement unit and a steering engine. According to the system, power management is unified on hardware, computing resources are shared, and on software, a real-time operating system is used for operating multi-task software on the same hardware platform. Navigation and comprehensive electronic deep integration are not simple physical combination of traditional on-missile electronic equipment, but efficient integration of needed functions is achieved, resources are utilized to the maximum extent, consumption is reduced, and improvement of the overall performance of tactical missiles is achieved with low cost, small size, light weight and high reliability.

The invention relates to a folding door, in particular to a target drone zero-length booster rocket launching opening folding door mechanism. The mechanism comprises a folding door plate, a torsional spring hinge, a quick-release lock and a quick-release lock locking mechanism, the number of the folding door plates is two, the folding door plates are oppositely arranged at the rocket launching port, and when the two folding door plates are in a closed state, the folding door plates are matched with the rocket launching port; the torsional spring hinges are divided into two groups; one ends of the two groups of torsion spring hinges are respectively mounted on the two folding doors, and the other ends are respectively mounted at rocket launching ports of the target drone; the quick-release locks are divided into two groups; the two groups of quick-release locks are respectively mounted on the two folding doors; the quick-release lock locking mechanisms are divided into two sets, and the two sets of quick-release lock locking mechanisms are installed at rocket launching ports of the target drone respectively and correspond to the two sets of quick-release locks respectively. According to the target drone zero-length booster rocket launching opening composite material folding door mechanism, the function that a folding door is rapidly and reliably closed after work of a zero-length booster rocket in the launching stage is completed and cold separation is conducted can be achieved.

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 discloses a liquid recoverable carrier rocket control system and control method, and belongs to the technical field of rocket recovery control. Comprising a flight control combination, a recovery control combination, an inertia measurement combination, a battery, a safety control combination, an engine controller, an optical target camera, a laser range finder, a comprehensive collecting and editing device, a relay measurement and control terminal and a space-based telemetering phased array antenna. The flight control combination and the recovery control combination are arranged in the second rocket sublevel and the first rocket sublevel respectively, partial hot backup is achieved through the two control combinations, and the reliability of the system is improved. In the first-stage flight section with the worst and most dangerous mechanical environment, the recovery control combination can serve as hot backup execution work of the flight control combination, and normal recovery of the first sub-stage cannot be affected even if flight failure is caused by abnormity of the second-stage rocket. Effective control over launch and recovery of the carrier rocket can be met, and a necessary control system is provided for recovery of the carrier rocket.

The invention provides a carrier rocket control method based on wind field characteristics, and the method comprises the steps: obtaining a wind field model generated through the fitting of actually measured wind field data; according to the speed of the carrier rocket at the current moment under the launching coordinate system, the wind speed at the current moment and the first conversion relation, the relative speed of the carrier rocket relative to the atmosphere at the current moment in the aircraft coordinate system is determined; influence parameters are calculated through the relative speed, and the aerodynamic force at the current moment is calculated according to the influence parameters; integrating the mass center motion differential equation set of the carrier rocket by combining the aerodynamic force, the preset initial value of the preset control parameter and the flight program angle expression of the carrier rocket changing along with time, and circularly calculating the preset state variable of each moment; calculating a target function of the carrier rocket according to the preset state variable at each moment; and on the basis of a sequential quadratic programming algorithm, target functions corresponding to different preset control parameters are calculated through the target functions and preset constraint conditions, so that optimal values of the preset control parameters are determined.