212 results about "Rocket launch" patented technology

The invention discloses a diversion cooling device and a rocket engine ignition test system, and belongs to the technical field of rocket launching tests. The flow guide cooling device comprises a spraying piece, a flow guide mechanism and a water supply mechanism. The spraying piece is arranged at the initial section of the tail flame of the rocket engine, a water curtain barrier can be arranged through spraying from top to bottom, cooling and noise reduction are achieved, fuel gas of the tail flame is prevented from burning other devices, and the temperature of the cooled tail flame cannot be too high when the cooled tail flame makes contact with the flow guide mechanism, and the flow guide mechanism cannot be damaged; in the flow guide mechanism, a flow guide groove water spraying panel which is in an arc shape and protrudes downwards in an inclined mode and flow guide groove bodies located on the two sides of the flow guide groove water spraying panel are used for conducting comprehensive spraying and flow guide on the high-impact-force and high-temperature tail flame, cooling and noise reduction are achieved, and water spraying areas with different spraying densities can conduct targeted spraying according to the temperature of the inner flame and the outer flame of the tail flame. And the arc plate and the second nozzle are detachable, so that the arc plate and the second nozzle are convenient to replace, assemble and disassemble after a certain part is damaged by high temperature, the cost is reduced, and the use economy is improved.

The utility model provides a supporting and quick separating mechanism of a rocket stabilizing device, the supporting and quick separating mechanism is connected to the bottom of the rocket stabilizing device, and the supporting and quick separating mechanism is connected to an anti-overturning device chassis of the rocket stabilizing device and is supported at the bottom of an anti-overturning arm of the rocket stabilizing device; after the rocket is ignited, the supporting and quick separating mechanism is unlocked, and the anti-overturning arm is released; before the rocket is not ignited, the supporting and rapid separating mechanism is connected to the bottom of the anti-overturning arm in a supporting mode. The supporting and rapid separating mechanism comprises a two-stage deloading supporting mechanism and an aluminum strip; the two-stage load shedding supporting mechanism is connected to the anti-overturning device underframe and is connected to the bottom of the anti-overturning arm in a supporting manner; the aluminum strip is arranged at the bottom of the rocket, and the two ends of the aluminum strip are connected with the two-stage load shedding supporting mechanisms correspondingly. According to the supporting and rapid separating mechanism, stable supporting and unlocking releasing of the anti-overturning arm of the rocket stabilizing device are achieved, so that the anti-overturning arm is rapidly opened, and rocket launching is not hindered.

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 utility model relates to the technical field of artificial precipitation enhancement, and provides a universal figure rocket launcher which comprises a launch canister, four hollow plates are fixed on the outer wall of the launch canister, the hollow plates are communicated with the inner side of the launch canister, the inner sides of the hollow plates are movably connected with fixing assemblies in an inserted mode, the number of the fixing assemblies is four, and the number of the fixing assemblies is four. And the fixing assembly comprises a moving plate, an arc-shaped fixing plate and a connecting rod, the moving plate is movably connected to the inner side of the hollow plate in an inserted mode, and the arc-shaped fixing plate is fixed to one side wall of the moving plate. The fixing assembly is arranged on the inner side of the hollow plate of the outer wall of the launch canister, the rocket projectile is inserted into the inner side of the launch canister, the empennage is inserted into the inner side of the empennage groove, the four arc-shaped fixing plates are pushed by the spring to clamp and fix the outer wall of the rocket projectile, and the size of a circular space defined by the four arc-shaped fixing plates can be changed according to elastic stretching of the spring. And therefore, rocket projectiles with different specifications and sizes can be fixed, and the universality during use is enhanced.

The utility model discloses a flow and pressure controllable water spraying noise reduction system and rocket launching auxiliary equipment relates to rocket launching auxiliary equipment technical field, and it is including: water storage tank, flow pump group, flow pump group connects water storage tank and drainage pipeline, water spraying pipe, water spraying pipe's first end connects drainage pipeline, and the second end of water spraying pipe sets up first control valve, circulating water pipe, circulating water pipe's first end connects drainage pipeline, and the second end of circulating water pipe connects water storage tank, and sets up second control valve on circulating water pipe. The above-mentioned flow and pressure controllable water spraying noise reduction system has solved the technical problem that the current water spraying noise reduction system cannot adjust water spraying flow, and water spraying timeliness is poor, and cannot provide large flow water flow in time, and influences the noise reduction effect, and before the rocket launches, the cooling water in the drainage pipeline keeps certain flow rate circulation, guarantees that the cooling water in the water spraying pipe can be in time, and the sufficient amount of water spraying pipe is sprayed from water spraying pipe, reaches the best noise reduction effect.

The utility model discloses an unmanned aerial vehicle booster rocket launching separation mechanism, and relates to the technical field of unmanned aerial vehicle design, the unmanned aerial vehicle booster rocket launching separation mechanism comprises a rocket connecting piece, one end of the rocket connecting piece is used for installing a rocket body, and the other end of the rocket connecting piece is provided with a push rod; the push rod is of a hollow structure, a pressure spring is installed in the push rod, and the end, away from the rocket connecting piece, of the push rod is movably sleeved with a driving guide head. The rocket mounting base is used for being connected with the unmanned aerial vehicle, a second hook is mounted above one end of the rocket mounting base, and the lower portion of one end of the rocket mounting base is connected with the driving guide head through a rotating disengaging mechanism. The push rod is placed in the driving guide head after the compression spring is placed in the push rod, thrust is generated after a rocket is ignited and launched to compress the compression spring, the first hook and the second hook are separated, the spring pin arranged on the driving guide head is bounced into the guide groove of the push rod, the first hook and the second hook are hooked again, and therefore it is guaranteed that the launching disengaging mechanism is smoothly disengaged from a machine body.

The invention discloses a figure rocket launcher pitching transmission shaft torque dynamic detection method and system, and the method comprises the steps: enabling a stepping motor to drive a driving gear through a speed reducer to drive a pitching transmission shaft to rotate, and collecting a torque signal in real time through a magnetoelectric induction torque sensor; and the phase difference measuring unit converts the dual-channel sinusoidal signal into a square wave and calculates the phase difference, and then outputs a digital torque value through the analog-to-digital converter. And the control module adopts Kalman filtering and a calibration curve to realize dynamic compensation. The recording sub-module is provided with a high-capacity solid-state memory and supports continuous recording of detection data in a circulating storage mode. The differential sensing structure is adopted, the neodymium iron boron magnetic steel and the multi-layer winding signal coil design are combined, the detection precision of the micro torque fluctuation under the low-speed and heavy-load working condition is effectively improved, the technical problems of detection lag, transient impact missing detection, data tracing difficulty and the like existing in a traditional method are solved, and the detection precision of the micro torque fluctuation under the low-speed and heavy-load working condition is improved. And the safety monitoring capability of the rocket launcher transmission system is obviously enhanced.

According to the rocket flight result rapid evaluation method, rocket flight data can be processed, an interpretation report is automatically generated, and working state rapid interpretation of pre-cooling before launching of a liquid rocket engine, fault diagnosis before ignition and in the flight process is achieved. According to the rapid judgment method for the rocket flight result, the analysis efficiency can be remarkably improved, and the time from the end of rocket flight to the obtaining of the engine analysis result is greatly shortened through automatic data processing, analysis process and rapid report generation. According to the evaluation method, the standard interpretation report is automatically generated, the problems of low manual analysis efficiency, incomplete analysis and the like are solved, and powerful support is provided for evaluation and improvement of rocket launching tasks.

The invention provides a small solid engine thrust performance test bench, and belongs to the field of rocket launching ground equipment. The fixed base is fixed to the ground, the linear sliding rail assembly comprises a sliding supporting frame and a sliding rail, and the sliding rail is fixed to the fixed base; the platform is fixed to the top of the sliding supporting frame, and a clamping supporting mechanism is arranged on the platform. The force transmission contact part is fixed at the front end of the sliding support frame; the vertical support frame is fixed at the front end of the base, forms an inverted L-shaped structure with the base and serves as a mounting surface of the thrust sensor; when an engine is ignited, thrust enables the platform and the sliding supporting frame to integrally move towards the thrust sensor along the sliding rail, and when the force transmission contact part collides with the thrust sensor, the thrust sensor outputs a thrust signal. The universality, the reliability and the test precision of the solid rocket engine ground static test system are remarkably improved, and the method is particularly suitable for research and development verification and performance evaluation stages of small and medium-sized engines.

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.

This invention discloses a method for extending the range of guided rockets using a single-chamber, dual-thrust design. By rationally designing the guided rocket's acceleration process after launch, this method effectively reduces energy loss during passage through dense atmosphere, improves the guided rocket's energy conversion efficiency, and ultimately, with a comparable charge, enables the guided rocket to achieve greater mechanical energy at the end of its active phase, thereby effectively increasing its range. A notable feature of this method is that the end of the guided rocket's active phase is located at a higher altitude, thereby achieving a better thermal environment in the active phase while simultaneously increasing its range.

The invention discloses a rocket projectile launching end point intensity influence factor analysis method, system and device and a medium. The method comprises the following steps: establishing a six-degree-of-freedom rigid in-vitro ballistic equation of the variable mass rocket projectile in a matrix form; the method comprises the following steps: generating a parameter sample based on a Monte Carlo random simulation principle by comprehensively considering three types of influence parameters of rocket projectile processing errors, initial disturbance and meteorological conditions, performing external trajectory equation numerical integration by combining a fourth-order Runge-Kutta method, and calculating a rocket projectile launching end point coordinate; three uncertainty sensitivity analysis methods are adopted, the influence degree of various parameters on the emission end point is quantified, results of the three methods are compared, and the influence degree of each factor is determined. According to the method, the influence degree of each factor on the rocket projectile launching terminal point is compared and analyzed by using multiple uncertainty sensitivity analysis methods, and the accuracy is relatively high.

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 invention provides a rocket launching safety analysis method based on a mobile rocket launching device, and the method comprises the steps: assembling a first finite element model of the rocket launching device and a second finite element model of a rocket body according to each typical working condition demand before rocket launching, and generating an analysis model corresponding to each working condition; corresponding preset conditions are applied to all the analysis models, simulation testing is carried out, and testing data corresponding to all the analysis models are obtained; and comparing each test data with a corresponding preset safety threshold, and outputting a safety evaluation result of rocket launching. In the design stage, the problems of insufficient structural strength, installation interference and the like are found in advance by simulating and covering the full-process working conditions of erecting, docking, to-be-launched and the like, and the subversive risk in a physical test is avoided. According to the scheme, through multi-aspect evaluation, the comprehensiveness of safety analysis is improved, the dependence on a physical test is reduced, and the research and development cost and period are reduced.

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 invention discloses a mechanical clone manufacturing method based on shape memory alloy pre-programmed geometrical shape information, which relates to the technical field of space manufacturing of space structures or components, and is characterized in that the shape memory alloy can be used for pre-programming and storing geometrical shape information and restoring the geometrical shape information under proper conditions; geometric shape information of a structure or a component is programmed and stored in the shape memory alloy on the ground and then emitted to the space, the shape memory alloy is triggered by solar energy to recover the original shape, and therefore the stored geometric shape information is mechanically cloned to a space recovery material. Or mechanically cloning to structures or blanks for rocket launching and transportation, so that large-size components or structures manufactured on the ground are restored in the space, or geometrical shape information programmed and stored on the ground is converted into components in the space, and the problems that space structures or components are limited by rocket space or heavy in weight, and the like are solved. Transportation is difficult, the manufacturing cost is high, and the manufacturing efficiency is low.

The present application relates to a target machine, in particular to a zero-length boost type target machine with foldable cabin door. The target machine comprises a fuselage, a zero-length thrust structure, an automatic folding door and a zero-length boost rocket; a rocket launching port is arranged on the lower side of the fuselage near the tail; the zero-length thrust structure is installed inside the fuselage at a position near the nose at the rocket launching port; the mounting end of the zero-length boost rocket is connected with the zero-length thrust structure through the rocket launching port, the output end is arranged outside the rocket launching port, and an included angle is arranged between the axial direction of the zero-length boost rocket and the length direction of the fuselage; the automatic folding door is installed at the rocket launching port of the fuselage and is matched with the rocket launching port. The present application adopts the zero-length folding door scheme, before boost flight, the automatic folding door is opened, the zero-length boost rocket is filled in the folding door and connected with the zero-length thrust structure, after boost separation, the automatic folding door is closed, the target machine restores the normal aerodynamic shape, and the subsequent flight performance and stealth performance are not affected.

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 present application relates to a kind of air cylinders, in particular to a kind of reliable backward, no rebound backward buffering cylinder for rocket launching, including cylinder body, cylinder piston rod, cylinder piston, spring, buffer cylinder body, buffer piston, throttle rod, connecting piece, the cylinder body one end is provided with air inlet, the buffer cylinder body, the buffer piston, the throttle rod all are located in the rodless cavity of the cylinder body, the center line of the cylinder piston, the throttle rod, the buffer piston is on same straight line, the rodless cavity of buffer cylinder body is filled with hydraulic oil, the buffer cylinder body is provided with pressure tap, the end of the buffer piston in the buffer cylinder body is provided with through hole, the other end of the throttle rod is arranged in the through hole, clearance is left between the through hole and the throttle rod, clearance forms annular throttling oil channel.The present application can drive support arm to reliably and quickly backward after rocket ignition and there is no rebound after backward.

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.

A rocket propellant tank and layout (15) is disclosed. The tank (15) includes a housing operationally coupled to a rocket engine (20). The housing comprises a plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) arranged longitudinally along the housing and is adapted to store a propellant. A plurality of separator walls (24, 41, 42, 32, 33, 34) are disposed adjacent to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f). A set of control valves (45) are coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f). The set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to the rocket engine (20).

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.