500results about "Self-propelled projectiles" patented technology

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 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 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 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 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.

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 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.

The application discloses a recovery control device for sinking mine training recovery, which comprises an ignition driving module, a depth sensor and a control module; the output port of the ignition driving module comprises a first ignition head, a second ignition head, a third ignition head and a fourth ignition head; the first ignition head and the second ignition head are connected in parallel, and the third ignition head and the fourth ignition head are connected in parallel; when the device adopts the mode of using a pyrotechnic product to open an air inlet valve to realize airbag inflation, the first ignition head and the third ignition head are connected to the same group of controls, and the second ignition head and the fourth ignition head are connected to the same group of controls; when the device controls airbag inflation by controlling the opening and closing of a solenoid valve, the first ignition head and the second ignition head are connected to the same group of controls, and the third ignition head and the fourth ignition head are connected to the same group of controls; the application can realize precise control of sinking mine floating recovery in sinking mine training, and improve the reliability of the control structure.

The invention discloses a synthetic attack angle control method during aircraft-rocket separation, and relates to the technical field of spaceflight, and the method comprises the following steps: controlling a deviation value between a pitching program angle and a trajectory inclination angle to be smaller than a preset deviation threshold value, and adjusting the pitching program angle in real time according to the trajectory inclination angle, thereby reducing a trajectory attack angle; when the ballistic attack angle is within a preset ballistic attack angle threshold value, an overload difference value between the actual normal overload at the current moment and the normal overload at the corresponding current height is calculated, and the correction amount of the pitching program angle is calculated based on the overload difference value; the correction amount of the pitching program angle is compensated into the pitching program angle, and the synthetic attack angle of the rocket is converged, so that the wind attack angle is reduced; before device-rocket separation, transverse normal acceleration data are continuously collected, and if the transverse normal acceleration data exceed a change threshold value, device-rocket separation is carried out after the acceleration data are recovered to be normal. According to the method, potential safety hazards caused by consideration of a single factor can be effectively avoided, and the safety and reliability of the separation process are remarkably improved.

The application discloses a rocket attitude self-adaptive control method and system based on reinforcement learning, and belongs to the technical field of rocket attitude control. The method constructs an adaptive controller for controlling the rocket attitude through a backstepping method, and the adaptive controller is used for realizing the self-adaptation of the rocket to unknown disturbances based on an Actor-Critic (AC) of reinforcement learning, so that the system can rapidly converge to a desired state, unknown nonlinear observation is realized without depending on an internal model, and the tracking control of the rocket attitude is realized.

This application proposes a small-scale universal modal testing device for rockets, comprising a suspension mechanism for suspending the rocket body and a testing mechanism with an impact end, wherein the impact end can impact the rocket body around a first axis. It also includes a drive mechanism for driving the impact end to impact the rocket body, and a testing platform for mounting the suspension mechanism and the testing mechanism. In use, the impact end is rotated around the rocket body with the first axis as the center and stopped at a suitable position. At this point, the drive mechanism drives the impact end to impact the rocket body. Compared to existing technologies that only include four impact points (left, right, top, and bottom), the impact end that can surround the rocket body has more impact positions, allowing for more flexible adjustment and facilitating the simulation of multi-position impact tests.

The invention relates to a hybrid engine (10) for projectiles in which a rocket engine with an annular nozzle (7) is arranged along with a base bleed unit with a circular nozzle (8) where the circular nozzle (8) of the base bleed unit is arranged centered in the annular nozzle (7) on the rocket engine and where the hybrid engine (10) is arranged with a propellant charge for the rocket engine (2) and a propellant charge for the base bleed unit (5). The invention further relates to a projectile (100) outfitted with a hybrid engine.

A roll decoupling joint is provided for connecting a rocket motor to a payload. The joint includes a payload adapter, a motor sleeve, a rotational separator and a fastener. The adapter includes an annular socket and a cylindrical core. The socket is bounded by an opening lip and a bulkhead. The payload inserts into the socket through the lip. The core extends axially from the bulkhead opposite the lip. The sleeve is annularly axi-symmetric and inserts into the motor. This sleeve is disposed adjacent to the adapter and around the core supported by the separator. The fastener secures the core to the separator. In various embodiments, the separator constitutes a pair of bearings, and the fastener is a threaded bolt.

A system comprising an unmanned aerial vehicle (UAV) (100) configured to transition (520) from a terminal homing mode (510) to a target search mode (530), responsive to an uplink signal (451) and / or an autonomous determination of scene change.

An apparatus includes at least one memory configured to store map data. The apparatus also includes at least one processor configured to segment one or more objects from one or more environment surfaces in the map data. The at least one processor is also configured to determine an offset based on a projectile drift. The at least one processor is further configured to generate a safety bounding box around each of the one or more objects using the offset.

The invention belongs to the technical field of noise elimination and flame suppression equipment, and particularly discloses a noise elimination and flame suppression device which comprises a noise elimination and flame suppression shell, the noise elimination and flame suppression shell is in a hollow ring shape, the noise elimination and flame suppression shell is filled with a noise elimination and flame suppression agent, and the noise elimination and flame suppression shell can be crushed under the action of high-temperature and high-pressure fuel gas. The two-phase mixed combustion noise reduction and flame reduction technology is adopted, so that the emission noise is greatly reduced, when the initial velocity of a bullet is about 200 m / s, the emission noise can be reduced to 170-175 decibels under the open field condition, the noise can be reduced to 180-185 decibels when the bullet is emitted in a limited space, the emission noise can be remarkably reduced, the sound pressure can be reduced to 1 / 3 of the sound pressure which is not additionally provided with the noise elimination and flame reduction device, and the noise reduction and flame reduction effects are good. After emission, a calcium chloride solution at the emission port and the tail part of an emission object is atomized to wrap flames, so that flame at the emission port and emission tail flames can be obviously reduced, and flame visible to naked eyes basically does not exist; the training clothes can be kept complete when being launched in a limited space, and cannot be burnt out, broken down or ignited by tail flames or post-spraying objects.

The invention provides a multi-radar information fusion measurement method for multiplex rocket suspension recovery, which comprises the following steps: acquiring a distance measurement value of the surface of a rocket through a plurality of millimeter-wave radars arranged on a rocket recovery device, and screening out effective measurement points based on a preset effectiveness threshold value; based on the screened effective measurement points, solving an initial estimation value of a circle center position by using circular geometric constraints of the cross section of the rocket; and constructing a state vector containing the center position and the motion state of the rocket, and performing multi-frame data fusion on the initial estimation value by using a Kalman filtering algorithm to obtain the optimal estimation of the center position after smoothing. According to the method, effective data of multiple radars are fused, geometric constraints of the cross section of the rocket are introduced to carry out space domain circle fitting, and dynamic smoothing and tracking are carried out in a time domain in combination with Kalman filtering, so that the purpose of improving the measurement accuracy of the landing pose of the rocket is achieved.

The invention relates to the technical field of rockets, and discloses a storage tank mounting bracket, a liquid attitude and orbit control power system and a rocket, and the storage tank mounting bracket comprises a mounting plate which is of an annular structure, sleeves the periphery of a storage tank, and is detachably connected with an end frame on the periphery of the storage tank in the circumferential direction; the supporting legs are dispersedly arranged on the periphery of the mounting plate, the top surfaces of the supporting legs are detachably connected with the bottom surface of the outer edge of the mounting plate, and the bottom ends of the supporting legs are placed on the same horizontal plane; a mounting space is formed between the plurality of supporting legs and the bottom of the storage tank and is suitable for mounting a pipeline connected with a liquid cavity interface at the bottom of the storage tank. By the adoption of the technical scheme, a pipeline connected with the bottom-layer liquid cavity connector of the storage tank is convenient to install, and the storage tank is convenient to disassemble, simple in structure, easy to machine, low in use cost and long in service life.

The application discloses an energy management method for correcting additional induced drag by online prediction, which comprises the following steps: firstly, online prediction of terminal velocity is realized by using multi-stage variable step length integral parallel calculation; secondly, the additional induced drag is calculated by using the difference between the predicted terminal velocity and the expected terminal velocity; finally, the additional guidance command is determined by using the existing guidance law to realize deceleration. The method has the advantages of strong universality, long action time, high calculation precision and strong engineering practice ability.

The invention discloses a head-tail double-end reverse thrust recoverable rocket based on double-layer spiral interlayer gas storage and a control method of the head-tail double-end reverse thrust recoverable rocket. A rocket body is of a double-layer stainless steel cylinder wall structure, a continuous spiral connecting grating is arranged between the inner wall and the outer wall, the cross section of the continuous spiral connecting grating is petal-shaped, and a closed high-pressure-resistant interlayer storage cavity is formed. An inert gas nozzle array driven by a plurality of independent electromagnetic control valves is arranged at the bottom of the rocket, and the inert gas nozzle array and the main propulsion engine are arranged cooperatively; the head part is provided with centralizing nozzles which are annularly distributed; a double-umbrella system is arranged on the top. Through a unique spiral grating structure and a head-tail double-end cooperative control strategy, the problems that a traditional rocket gas storage system is large in weight, and a long and thin rocket body is difficult to resist wind and stabilize the posture are solved, the landing precision and safety are remarkably improved, and the method is suitable for a reusable carrier which is low in cost and high in reliability.

The application discloses a weather modification rocket bomb three-explosion self-destruction body assembly detection system, and relates to the technical field of multivariate comprehensive measurement. The system comprises a multidimensional positioning measurement module, which is used for measuring the coaxiality and fitting pressure of the self-destruction body and the engine shell; a medicament loading multi-parameter measurement module, which is used for measuring the weight, density and internal defects of the medicament; a connection strength dynamic measurement module, which is used for measuring the torque, tensile strength and vibration displacement; an igniter performance parameter measurement module, which is used for measuring the resistance, delay and safety reliability; a measurement data integration analysis module, which is used for integrating data to generate a report and an encrypted log; and a system control and abnormality processing module, which is used for coordinating the module time sequence, monitoring data and processing abnormalities. The system realizes integrated measurement of multiple parameters of the three-explosion self-destruction body, improves the positioning, medicament loading, connection strength and igniter performance measurement accuracy, guarantees that the wreckage is in compliance after self-destruction, and data is traceable.

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