109 results about "Rocket launch" patented technology

The invention provides a simulation method and system for the launch process of a carrier rocket and electronic equipment, and relates to the technical field of rocket launch simulation. The method comprises the steps that a launching scheme of the carrier rocket is determined, and the launching scheme comprises launching parameters of the carrier rocket; modeling and verifying the launching scheme; and performing simulation test on the launch process of the carrier rocket based on the verified launch scheme, and responding and correcting faults in the simulation test process. According to the invention, the problem that the safety of the launch task of the carrier rocket is threatened due to the fact that key design defects are not found in time due to multiple technical bottlenecks existing in the distributed simulation technology of the reusable carrier rocket is solved.

The invention relates to the technical field of rocket launching, in particular to a steam protection rocket launching system which comprises a steam energy storage system, a steam conveying system, a steam releasing system and a launching rack structure. The steam energy storage system comprises a steam boiler, a steam pressure accumulator and a steam energy storage control unit and is used for preparing and storing high-pressure steam before rocket launching. The steam conveying system comprises a steam blending table and a steam supply pipeline and is used for conveying high-pressure steam in the steam pressure accumulator to the steam release system; the steam release system comprises a steam injection unit and a steam injection control system and is used for releasing steam when the rocket is ignited, launched or recovered, forming a steam protection air curtain and reducing thermal shock of high-temperature airflow on the structure. By establishing the high-pressure steam protection air curtain, active protection of structural parts, pipeline parts, hydraulic parts and other parts of the rocket launching system influenced by the gas tail flame area of the rocket engine is achieved, and the durability and reliability of the launching system are improved.

The utility model provides a launcher enabling a rocket to rotate forwards, which comprises a rocket body, a rocket airfoil mounted at the tail of the rocket body, a fixing frame arranged on the outer side of the rocket body, a fixing plate mounted on the inner side wall of the fixing frame, an adjusting ring arranged in the fixing plate, clamping pressure plates arranged on the upper surfaces of the fixing plate and the adjusting ring, and bolts arranged in the clamping pressure plates. Bolts penetrate through the clamping pressing plates to be spirally connected with nuts, and spiral guide rails are movably installed at the tail ends of the inner sides of the adjusting rings. According to the launching device capable of enabling the rocket to rotate forwards, the safety, stability and reliability of the rocket during launching can be improved, and the anti-jamming capability of the rocket is effectively improved.

A space traffic management device (100) mounted in a mega-constellation business device (41) specifies one to a plurality of representative satellites from a mega-constellation satellite group flying on orbits having the same nominal orbital altitude, and has quasi-real-time high-accuracy orbital information of the representative satellite and orbital information relative values of constituent satellites other than the representative satellite, relative to the representative satellite. The space traffic management device (100) shares the quasi-real-time high-accuracy orbital information of the representative satellite and the orbital information relative values of the constituent satellites relative to the representative satellite, with the space traffic management devices (100) mounted in a plurality of business devices.

The invention belongs to the technical field of marine rocket launching, and particularly relates to a marine rocket launching platform with a rocket body stabilizing structure. A base for mounting a rocket is arranged on the ship body; a group of fixing rings are arranged on the ship body, and a group of first hydraulic rods which are annularly and uniformly distributed are fixedly connected to the inner walls of the fixing rings; the output end of the first hydraulic rod is fixedly connected with a supporting and stabilizing block which is used for supporting and stabilizing a rocket. The rocket is installed on the base, at the moment, the rocket is located in the center of the fixing ring, when the ship body operates on the sea, the first hydraulic rod is started, the output end of the first hydraulic rod pushes the supporting and stabilizing block, the supporting and stabilizing block moves towards the side close to the rocket body, and at the moment, the supporting and stabilizing block makes contact with the rocket body; and therefore, the rocket body is supported stably, and the stability of the rocket during transportation is improved.

The invention discloses a glass breaking and fire extinguishing unmanned aerial vehicle system and an operation method, the glass breaking and fire extinguishing unmanned aerial vehicle system comprises a rotor unmanned aerial vehicle and an observing and aiming launching system, and the observing and aiming launching system comprises an automatic holder, a photoelectric pod, a fire control system, a launch canister and a ground control station system. During working, the unmanned aerial vehicle transmits a fire scene image in real time by means of the photoelectric pod, and a target launching position is identified and determined by a ground operator; after the unmanned aerial vehicle arrives at the target area, distance measurement and trajectory calculation are carried out on the target through the photoelectric pod; the automatic holder adjusts the aiming posture according to the resolving result; after the ground control station completes rocket launching time sequence setting, the fire control system controls the launching device to sequentially launch the glass breaking rocket and the fire extinguishing rocket, the glass breaking rocket and the fire extinguishing rocket accurately hit the core area of the fire scene, and fire suppression is achieved. The unmanned rotorcraft is used as a carrying platform, an observing, aiming and launching system is integrated, and the unmanned rotorcraft has multiple functions of target distance measurement, trajectory calculation, impact point superposition and rocket detection and launching and is suitable for complex fire scene environments such as high-rise buildings.

The application discloses a kind of based on liquid rocket's interstage separation gap calculation method and system, including establishing the inertial coordinate system of liquid rocket launch, rocket body coordinate system and the coordinate system of storage tank;Obtain the dynamics equation of liquid rocket mass center motion in inertial coordinate system and the dynamics equation of liquid rocket rotation around mass center in rocket body coordinate system;Obtain the displacement H of liquid rocket core first stage and liquid rocket core second stage and attitude angle;Obtain the displacement L1 of liquid rocket core second stage mass center deviating from nozzle center line, the displacement L2 of liquid rocket core second stage mass center deviating from liquid rocket core first stage axis and the distance L3 of nozzle outer edge to nozzle center line;Obtain interstage separation gap as L4, L4=R-L3-(L1-L2), solve the problem that core second stage engine nozzle and interstage section sleeve may collide in the process of large-diameter liquid rocket separation, to improve the precision of separation scheme design, attitude control system design and storage tank anti-slosh structure design.

The application provides a modular airborne detection and attack micro-rocket launching device, which comprises a weapon station module, a shell cabin module and a female head, a male head and a transmission module are respectively electrically connected to two ends of an electric control module and are controlled by the electric control module; a rotating shaft rotates along an axis; the shell cabin module is detachably connected to the rotating shaft and rotates with the rotating shaft; a shell cabin body is used for accommodating spare bullets; the female head is detachably connected to the male head, at this time, the weapon station module, the female head and the shell cabin module are in a non-integral structure, and the weapon station module, the female head and the shell cabin module can be separately disassembled, so that the modules of the modular airborne detection and attack micro-rocket launching device cannot be separately maintained, the maintenance convenience of the modules of the modular airborne detection and attack micro-rocket launching device is improved, in addition, the shell cabin module is detachably connected to the rotating shaft and rotates with the rotating shaft, so that the angle of the shell cabin module relative to the transmission module can be adjusted.

A wireless model rocket launcher and rocket which provide for safety checking and advanced functionality to be performed wirelessly using a mobile communication device such as a cell phone. A set of safety checks can be performed before the rocket is armed and triggered for launch. In addition, a location of the rocket can be tracked wirelessly on the mobile communication device.

The present application belongs to the field of rocket launch measurement and control technology, and particularly relates to a method and system for timing the separation of a rocket connector. The method comprises: collecting and preprocessing a separation trigger signal generated when the rocket connector separates from the rocket body; identifying the first generation time of the separation trigger signal based on the preprocessed separation trigger signal; determining the separation start time using the first generation time as a reference; simultaneously receiving a positioning signal of the rocket connector and using the disappearance time of the positioning signal as the separation end time; obtaining the separation duration based on the difference between the separation end time and the separation start time; storing the separation start time, the separation end time, and the separation duration; comparing the separation duration with a preset standard duration; generating a timing report and sending a timing completion instruction; and completing the timing of the separation of the rocket connector. The present application achieves timing of the separation of the rocket connector without the need for additional on-rocket equipment, has strong anti-interference capability, and has high timing accuracy.

The application discloses a manual adjusting device for the caliber of a human shadow rocket launcher director, which comprises a track assembly and a supporting assembly. The supporting assembly comprises four bearing rail frames. The track assembly comprises two groups of U-shaped track assemblies arranged in the left and right openings of the supporting assembly. Each U-shaped track assembly comprises an upper U-shaped large track, an upper U-shaped small track, a lower U-shaped large track and a lower U-shaped small track. The upper U-shaped small tracks are connected through a first linkage rod. The lower U-shaped small tracks are connected through a second linkage rod. A linkage plate is arranged between the two groups of U-shaped track assemblies. First and second inclined grooves are formed in the side walls of the linkage plate. The first and second linkage rods pass through the first and second inclined grooves respectively. The diameter of the envelope circle of the U-shaped small tracks can be changed through manual adjustment, so that one director can adapt to the launching requirements of rocket projectiles with various calibers.

The invention discloses a sea rocket launching platform wave compensation control method based on deep reinforcement learning. The sea rocket launching platform wave compensation control method comprises the following steps: S1, collecting real-time motion state data; s2, inputting the motion state data into an improved Shenchang differential equation model; s3, constructing a strategy network according to a SoftActor-Critic algorithm, and generating a corresponding continuous control action signal; s4, the control action signal is input into a propeller control system and a ballast adjusting system, and combined compensation control over the posture and the position of the platform is achieved; s5, in a simulation environment, carrying out end-to-end joint training on the ShentActor-Critic strategy network and the ShentActor-Critic strategy network; and S6, realizing wave compensation of the offshore rocket launching platform in the wave disturbance environment in the platform control system. According to the invention, accurate prediction and adaptive compensation of platform stability control under complex disturbance are realized, and the control precision and reliability are significantly improved.

The utility model relates to the technical field of toy rockets, in particular to a hand-operated power generation type rocket toy which comprises a base, a feeding pipe fitting is arranged on the base, an inner layer structure and an outer layer structure are arranged in the feeding pipe fitting, and a vertically-arranged rocket flying body is installed on the feeding pipe fitting in a sliding mode. The water storage rocket body is communicated with the outer layer structure; the fuel generation structure is communicated with the outer layer structure, and the fuel generation structure is used for forming combustible gas and conveying the combustible gas to the outer layer structure; and the ignition control structure is connected with the inner layer structure and used for igniting combustible gas in the feeding pipe fitting, so that in the rocket flying body launching process, timely and effective pressure relief is conducted on the feeding pipe fitting, and structural materials of the feeding pipe fitting are prevented from being broken under the action of pressure intensity.

The application discloses a rocket take-off moment measurement method and device based on an event camera, and belongs to the technical field of spaceflight measurement and control. The rocket take-off moment measurement method comprises the following steps: acquiring a rocket launch event set collected by an event camera, and decomposing the event set into two subsets according to event polarity; setting a dynamic time window to perform segmented stacking projection on the two event sequences respectively, generating an event frame with spatial dense characteristics, so as to eliminate sensor noise and environmental interference; extracting center line features by performing line segment detection, and determining a rocket initial motion moment based on center line coincidence determination in a continuous time window; starting an adaptive time backtracking mechanism, nonlinearly adjusting a backtracking step length according to the number change of feature line segments, and accurately determining a rocket take-off moment at a critical point where the feature line segments disappear. The application realizes the detection of the rocket take-off moment with microsecond-level precision, and has the advantages of high detection precision and strong anti-interference capability.

The utility model provides an adjustable weather modification rocket launching guide rail, including mainframe box, rotary support frame, adjusting plate, rocket launching guide rail, main guide rail, first telescopic vice guide rail and second limit vice guide rail, the rotary support frame is installed on mainframe box upper end, be provided with adjusting plate in the rotary support frame, be installed with rocket launching guide rail on adjusting plate upper end, be provided with main guide rail in rocket launching guide rail, be provided with first telescopic vice guide rail in main guide rail, be provided with second limit vice guide rail in main guide rail, this design has solved the problem that needs to change different model rocket frame guide rail back and forth when the launch of two kinds of rocket bomb length is inconsistent, the utility model discloses reasonable in structure, be provided with movable frame through setting up in fixed frame front end, it cooperates with sliding limit seat and first telescopic vice guide rail and second limit vice guide rail, need not to change rocket launching guide rail to launch different model rocket bomb.

The invention discloses a two-way rocket launching window searching method, system and device and a medium, and the method comprises the steps: rapidly calculating two nearest launching windows of an ascending orbit and a descending orbit of a rocket according to the attributes of a central body, target orbit plane parameters, the geographic coordinates of launching points and the flight duration of the rocket; according to the method, the problem is converted into numerical calculation based on space geometry, the crossing moment is directly solved by using the geometrical relationship between the normal vector of the orbital plane and the inertial position vector of the launching point location and adopting efficient numerical methods such as dichotomy and the like, and complex dynamic iteration is avoided. According to the method, the calculated amount is greatly reduced, rapid searching of the launching window is achieved, the method is suitable for an interstellar detection task scene needing rapid response or multiple times of calculation, and the real-time performance of task planning is remarkably improved.

The invention discloses a rapid docking device for rocket air conditioners, and relates to the technical field of rocket launching, the rapid docking device comprises a mounting ring, the bottom of the mounting ring is provided with a plurality of moving roller mechanisms and a plurality of supporting mechanisms, the mounting ring is provided with an annular mounting groove, and the annular mounting groove is provided with two symmetrically arranged moving frames; the movable frames are connected with the mounting ring in a sliding mode, a rotary displacement mechanism is arranged on one movable frame, a folding displacement mechanism is arranged on the movable frame, a rotation limiting clamping block mechanism is arranged on the folding displacement mechanism, a movable displacement mechanism is arranged on the folding displacement mechanism, a deflection ring mechanism is arranged on the movable displacement mechanism, and a butt joint mechanism is arranged on the deflection ring mechanism. According to the rocket air conditioner butt joint device, rapid movement of the device can be achieved through the arrangement of the moving roller mechanisms, positioning and leveling of the device can be achieved through the cooperation of the moving roller mechanisms and the supporting mechanisms, rotary displacement of the butt joint mechanisms can be achieved through the rotary displacement mechanisms, and the efficiency of rocket air conditioner butt joint is improved.

The invention provides a wired and wireless cooperative high-frequency dynamic data testing method and system for a rocket launching tower and a medium, and relates to the field of spaceflight launching testing. Comprising the steps that a far-end display control unit triggers a tower near-end collection and storage unit to be started at the second moment before a rocket is ignited, and triggers a wireless transmission unit to be started at the third moment after the rocket flies away from a launching tower; after the tower near-end acquisition and storage unit is started, rocket launching gas jet impact pressure data and launching tower structure vibration data in the rocket launching period are acquired and stored locally; a wireless link is established after the wireless transmission unit is started; the near-end acquisition and storage unit transmits locally stored data to the far-end display control unit through a wireless link, and the far-end display control unit analyzes and processes the received data to generate a test report. The problem of reliable acquisition and quasi-real-time analysis of multi-channel high-frequency dynamic data of rocket launching near-end high-frequency dynamic data under a limited space condition of a launching tower and a complex electromagnetic environment is solved.

Embodiments of the present application provide a timing control circuit, a safety detection method thereof and a rocket launch system. The timing control circuit comprises a power supply module, a control module and a monitoring module; the power supply module comprises a first power supply bus provided with a first switch and a second power supply bus provided with a second switch; the control module comprises a plurality of timing control relays; the monitoring module comprises a first protection resistor connected in parallel across the first switch, a second protection resistor connected in parallel across the second switch and a monitor, the monitor being configured to monitor an electrical signal across the first protection resistor and / or the second protection resistor during a safety detection process, and determine whether the control module is in a normal working state according to the electrical signal. In the embodiments, both switches are disconnected during the safety detection process, which can improve safety performance, and the two protection resistors can further improve safety performance, because even if misabsorption occurs during the safety detection process, the current supplied to the timing load is less than a safety current.

The invention relates to a design, processing and manufacturing method of a novel silicon-based terahertz frequency selection surface, and belongs to the technical field of electromagnetic fields and microwaves. The specific processing and manufacturing steps are as follows: manufacturing a silicon wafer; manufacturing a base; manufacturing a pressing ring; manufacturing a gasket; the gasket, the silicon wafer and the pressing ring are installed on the base, and manufacturing of the silicon-based terahertz frequency selection surface is completed; the silicon wafer is formed by stacking an upper silicon-based bonding structure, a middle silicon-based bonding structure and a lower silicon-based bonding structure; the thicknesses of the layers are the same; wherein the upper surface and the lower surface of the silicon wafer are respectively etched with periodic array patterns and then are subjected to gold plating treatment; the gap between the vertical side wall of the base and the silicon wafer is adjusted through the gasket; a chamfer is arranged on the side wall of the inner ring of the pressing ring; a supporting rib plate is arranged on the side wall of the base; the structure has the advantages of being high in structural strength, compact, good in universality, high in installation precision and the like, and can resist mechanical conditions in the rocket launching stage while meeting the requirement for filtering electrical performance.

The application discloses an online trajectory planning method for the final stage of a carrier rocket aiming at an elliptical target orbit and belongs to the field of carrier rocket trajectory and guidance, which comprises the following steps: step one, constructing the orbit entry terminal constraint and planning performance index aiming at the elliptical target orbit; step two, establishing the system dynamics constraint of the orbit entry of the final stage of the carrier rocket; step three, convex approximation is conducted on the planning orbit entry terminal constraint and performance index function; step four, state affine approximation and pseudo-spectrum discretization are conducted on the system dynamics constraint of step two; and step five, the optimal trajectory of the final stage entering the elliptical orbit is generated online through sequential convex iteration. The method of the application is suitable for the joint reconstruction of the final stage trajectory and the adaptive guidance of the carrier rocket launching the elliptical target orbit mission, can realize the online generation of the optimal trajectory with high precision and fast convergence, and is favorable for improving the reliability of the carrier rocket and enhancing the launch service fulfillment capability.

This satellite release mechanism (10) that is launched by a rocket (100) and releases satellites (16A, 16B) comprises: a first satellite mounting part (26A) to which a first satellite (16A) is fixed; a base part (18) including a first fixing part (27A) for holding the first satellite (16A) on the first satellite mounting part (26A), and a connection surface (24a) connecting with the rocket (100); and a first push-out mechanism (20U) for pushing out the first satellite (16A) from the base part (18), the first push-out mechanism including a first push-out member (36U) for pushing out the first satellite (16A) by moving in a predetermined movement direction inclined at a first inclination angle (θu) with respect to a line normal to the connection surface (24a) connecting with the rocket (100).

The invention belongs to the technical field of seaborne rocket launching equipment, and particularly relates to a rocket seaborne launching device with a buffer structure, which comprises a launching platform and a spoiler, and further comprises a buffer guide mechanism, and the buffer guide mechanism is used for being matched with the spoiler to realize buffer type flow choking of airflow; by arranging the buffer guide mechanism, on one hand, in the tail gas flow explosion process, impact force given by tail gas to the spoiler is continuously consumed through compression of a buffer spring and directional flowing of seawater, and compared with the spoiler in a fixed state, the impact force can be effectively dispersed; meanwhile, seawater flows under the action of impact force and finally flows to the surface of the spoiler, the heating effect of tail gas flow on the spoiler is delayed through evaporation heat absorption of the seawater, the final temperature peak value of the spoiler is reduced, the damage probability of the spoiler is further reduced, and the service life of the spoiler is prolonged. And the service life of the spoiler is prolonged.

This utility model relates to an RFID-based manipulator rocket launching device, comprising a base plate, a rotating mechanism fixedly connected to the base plate, a rotating seat rotatably connected to the rotating mechanism, a first driving component for driving the rotating seat to rotate within the rotating mechanism, a flipping seat fixedly connected to the rotating seat, a guide hinged to the flipping seat, and a pitch mechanism for driving the guide to flip on the flipping seat. The guide includes a guide rail structure for carrying and guiding the rocket, and an RFID scanning antenna is mounted on the guide. A first limiting component is provided within the rotating mechanism to limit the rotation of the rotating seat, and a second limiting component is provided within the flipping seat to limit the rotation of the guide. This utility model achieves multi-angle launching of manipulator rockets through the pitch and rotating mechanisms, and realizes automatic collection of ammunition information through the integrated RFID scanning antenna, achieving fully automated management of the rocket from loading, identification to launch.

Rocket launcher control system comprising a storage unit having one or more supercapacitors. Control system (16) for rocket launcher (12), comprising: - a human-machine interface (20) for entering firing parameters and including an integrated electric battery (26), - a control device (22) for receiving firing parameters, and an incoming electric current (34) from the integrated electric battery, the control device comprising: a regulation unit (38) for producing an electric current to be stored (40); a storage unit (42) having one or more supercapacitors (44) for receiving the electric current to be stored in the supercapacitor(s) and for producing a release electric current (46); a conditioning unit (48) for receiving the release electric current and for producing an output electric current (36) intended to be sent to the rocket launcher.Figure for the abbreviation: figure 2.

The liquid rocket engine multi-starting ignition system comprises a support and a base, the support is fixedly connected with the base, a servo motor is fixedly connected to the support, an output shaft of the servo motor is sequentially connected with a first gear and an ignition device, and a percussion device is arranged on the side face of the servo motor. The percussion device is fixedly connected to the side face of the servo motor shell through a guide rail frame. Under the condition that the number of gunpowder starters is not increased, multiple times of ignition on the rocket are achieved; and meanwhile, the reliability of the rocket is enhanced, even if a certain igniter breaks down, other igniters can still work normally, and it is guaranteed that the rocket launching task is completed.

A debris removal satellite includes a capture device, a thruster of a chemical propulsion method, and a propellant tank to store chemical fuel. A solar array wing is operable in an orbit at an orbital altitude higher than a congested orbit region congested with satellites forming a satellite constellation. The debris removal satellite is built in advance for future use as a satellite to be launched, and when a debris intrusion alarm to give a warning about intrusion of debris into the congested orbit region is issued, propellant is loaded into the propellant tank and the debris removal satellite is launched by a rocket built in advance for future use as a launch rocket. The debris removal satellite captures capture-target debris at an orbital altitude higher than the congested orbit region, and operates a propulsion device with the capture-target debris being captured.

The invention relates to a rocket boosting launching device and a launching method thereof, and belongs to the technical field of rocket boosting launching, the rocket boosting launching device comprises a boosting platform, a plurality of racks and a plurality of mechanical clamping arms, the bottom end of the boosting platform is fixedly provided with a plurality of liquid boosting engines A distributed in an array mode at intervals, and a rocket is supported on the table top of the boosting platform; the multiple racks are distributed on the peripheral side of the boosting platform at intervals in the circumferential direction, and the side faces, corresponding to the boosting platform, of the racks are provided with sliding grooves formed in the motion direction of the rocket. And the multiple mechanical clamping arms are oppositely clamped outside the rocket and slide in the multiple sliding grooves correspondingly, and liquid boosting engines B are fixed to the bottom ends of the multiple mechanical clamping arms correspondingly. According to the technical scheme of rocket power vertical boosting launching, the existing rocket vertical take-off launching efficiency can be effectively improved, the propellant mass of a rocket engine is reduced, the effective load of the rocket is effectively improved, meanwhile, the recoverability is improved, and the rocket launching cost is reduced.

The invention of the rocket structure test and launch has made a great contribution to the field of aerospace technology; especially in sounding rockets, it has three basic functions: the rocket's cold flow (cold flow 168; hot-fire test) and launch stages are integrated into a single structure. In addition, it is used to move the rocket from one place to another by attaching the launcher, which makes it different from other rocket launchers.

A debris removal satellite includes a capture device, a thruster of a chemical propulsion method, and a propellant tank to store chemical fuel. A solar array wing is operable in an orbit at an orbital altitude higher than a congested orbit region congested with satellites forming a satellite constellation. The debris removal satellite is built in advance for future use as a satellite to be launched, and when a debris intrusion alarm to give a warning about intrusion of debris into the congested orbit region is issued, propellant is loaded into the propellant tank and the debris removal satellite is launched by a rocket built in advance for future use as a launch rocket. The debris removal satellite captures capture-target debris at an orbital altitude higher than the congested orbit region, and operates a propulsion device with the capture-target debris being captured.