1383results 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.

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 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 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 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 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 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.

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 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.

The invention provides a non-contact distance measuring system and method. The non-contact distance measuring system comprises a first sensor, a second sensor and a converter, the converter comprises a sensor excitation unit, a sensor signal acquisition unit and a control unit; according to the environment temperature, the temperature compensation coefficient corresponding to the environment temperature is determined; compensating a voltage signal acquired by a sensor signal acquisition unit according to the temperature compensation coefficient; and obtaining the distance between the first sensor coil and the surface of the measured target object according to the compensated voltage signal. According to the invention, the environment temperature is measured through the second sensor, and the influence of the temperature on the measurement precision is reduced according to the voltage signal acquired by the temperature compensation sensor signal acquisition unit. According to the invention, the conversion of the eddy current signal can be realized, and the separation state of the detected target object is judged.

The invention discloses a visual trajectory tracking system and method for a rain enhancement and hail suppression rocket projectile. The system comprises a trajectory tracking system, a data server and a GPS module carried by the rain enhancement and hail suppression rocket projectile. The GPS module transmits flight path data of the rain enhancement and hail suppression rocket projectile to the data server in real time; the trajectory tracking system calls the data in the database server in real time, retrieves and counts the data, and displays the flight state and the real-time trajectory curve of the rocket projectile on a system interface in a visual form by using the retrieved data; according to the invention, based on a mode of combining a theoretical model and actual data, the problem that no signal data cannot be transmitted back in time is solved, and an experiment verifies that a curve combined by actual theoretical data is matched with a curve of complete real-time data; a catalyst sowing path, a rocket projectile flight path and a rocket projectile wreckage drop point can be accurately positioned.

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 present invention relates to a projectile, in particular deformation and / or partial fragmentation bullet, comprising an essentially cylindrical projectile tail, an adjoining bow sided projectile head with an essentially central opening which opens into a cavity that extends axially from the projectile head in the direction of the projectile tail, preferably into the projectile tail, which has a cavity bottom and is bounded by a wall, and an tear-off groove introduced into the wall and, at least partially surroundings the cavity, which is arranged at a distance of at least 10% of the longitudinal extension of the cavity from the cavity bottom and has a radial depth of at least 10% of a caliber diameter and / or of at least 30% of a radial wall thickness of the wall surrounding the cavity.

The invention relates to the technical field of cosmic sailing, and particularly discloses an offshore rocket recovery system and method, and the system comprises a floating body which floats on the sea surface; the inclination measuring unit is arranged on the floating body and is used for measuring the inclination angle between the floating body and the sea surface and generating an inclination signal; the recovery unit comprises a recovery deck, a rocket fixing device and a plurality of lifting devices, the rocket fixing device is installed on the recovery deck, the lifting devices are arranged on the periphery of the bottom of the recovery deck and installed on the floating body, and the lifting devices are used for driving the recovery deck to move up and down; and the control unit is arranged on the floating body, the control unit is in signal connection with the inclination measuring unit and the lifting device, and the control unit is configured to receive an inclination signal of the inclination measuring unit and control the lifting device to drive the recovery deck to move up and down. According to the offshore rocket recovery system and method, the rocket can be stably recovered, the recovery efficiency is improved, and the cost is reduced.

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 invention provides an inverse orbit cooperative guidance method based on a virtual target point. The method comprises the following steps: estimating a target speed direction according to target observation information; assuming that the target moves in the reverse direction of the missile with the maximum maneuverability, calculating the attack time required by each missile to hit the target according to the geometrical relationship and the motion relationship between the missile and the target, and taking the maximum value as the ideal attack time of the missile group; calculating a predicted hit point according to the target observation information, the target speed direction and the ideal attack time; according to the position relationship among the missile, the target and the predicted hit point, selecting a reverse orbit strike mode or a lateral strike mode, and determining a strike angle; virtual target points of all guided missiles are calculated according to the strike angles; and each missile is controlled to fly to the virtual target point to hit the target. By using the method, the target striking probability and the striking effect can be improved.

The invention relates to a further HDRM defeat device, there is a high density reactive material (HDRM) munition warhead comprising a main body with a payload cavity capable of receiving a high explosive payload, said cavity comprising, a high explosive, an HDRM material located between the high explosive and the main body, wherein the main body has a concave outer surface, and comprises a plurality of preformed fragments located on the outer surface of the main body, said fragments formed from a further HDRM material, said further HDRM material further comprising tungsten metal in the range of 10 to 50 wt%, such that in use the HDRM material and further HDRM preformed fragments are explosively driven together. such as to allow an explosive output from the high explosive to be more effectively transferred to initiate the HDRM material reaction.