376 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 invention discloses a remote supportable rocket test launch control system based on a cloud platform. The remote supportable rocket test launch control system comprises a rear-end remote test module, a remote test network and a front-end remote test module, wherein communication and data interaction between the rear-end remote test module and the front-end remote test module are realized through a remote test network; the front-end remote test module transmits the collected test data of the launching site and the workshop to the rear-end remote test module through a remote test network; and the rear-end remote test module realizes communication and data interaction with carrier rocket test launch control equipment and post personnel in various regions through a remote test network, and realizes remote auxiliary test, comprehensive management and remote command of the whole rocket launch process. The problems that a traditional rocket test launching mode is low in efficiency, and front-rear cooperation cannot be achieved are solved.

The invention discloses a universal liquid carrier rocket launching pad technology. A satellite general assembly test subsystem realizes recovery and detection of a satellite and forms a star cover assembly; the rocket horizontal final assembly test subsystem realizes subsystem test, matching test and total inspection test of a first-stage rocket body and a second-stage rocket body in a horizontal state, and the first-stage rocket body, the second-stage rocket body and a star cover combination body form a whole rocket in the horizontal state; the rocket horizontal transfer subsystem transfers the whole rocket to the fixed launching platform in a horizontal state; the rocket launching subsystem erects the whole rocket and achieves testing, filling and launching of the whole rocket. According to the universal liquid carrier rocket launching pad technology, by adopting a three-plane test launching mode of horizontal final assembly testing, horizontal transferring and vertical launching, the construction cost of a rocket launching site is effectively reduced, the operation process is simplified, the rocket launching efficiency and frequency are improved, and the requirement for rapid and batch launching is met.

The invention belongs to the technical field of rocket launching, and particularly relates to a carrier rocket launching perpendicularity control method and device and a storage medium, the perpendicularity control method comprises the following steps: constructing an attitude dynamic model in a carrier rocket launching process, and the attitude dynamic model is used for describing change rules of rocket attitudes under different conditions; setting a prediction window; according to the real-time attitude state information of the rocket, predicting attitude changes of a plurality of time steps in the future in the current prediction window; designing an objective function, and optimizing the objective function through a predictive control algorithm in each time step to obtain an optimal control signal sequence; in each prediction period, executing a first signal in the obtained optimal control signal sequence, and transmitting the first signal to a rocket attitude control device; and repeating the steps. According to the invention, the launch attitude control precision of the carrier rocket can be improved, and the response speed of the system is optimized.

A projectile and method for stopping aerial vehicles. The projectile comprises a shell—and a bunch of filaments. The bunch is wound in a coil and placed inside the shell. The projectile according to the solution may be brought to a target by different methods and utilising different means. For example, the projectile may be fired with compressed air or a propulsion charge, by using a compressed air weapon or rocket launcher. In addition, the projectile may be taken to the target by using a rocket or aerial vehicle.

The invention provides a method, a device and equipment for determining parameters of a flow baffle of a launching ship, and relates to the technical field of rocket launching analysis. The method for determining the parameters of the flow baffle of the launching ship comprises the following steps: acquiring a three-dimensional model of the flow baffle when a rocket is launched at a preset launching height; performing grid division on the three-dimensional model of the flow baffle to obtain a plurality of unit grids; performing numerical simulation calculation on the plurality of unit grids to obtain simulation calculation results; the simulation calculation result comprises the temperature and the pressure of each unit grid in the three-dimensional model of the flow baffle; according to the simulation calculation result, parameters of the flow baffle are optimized, target parameters are obtained, and the target parameters comprise the target height, the target width and the target thickness of the flow baffle. According to the scheme provided by the invention, the temperature and pressure of the flow baffle impacted by heat flow can be accurately simulated and analyzed at the preset emission height, and theoretical and data support is provided for parameters during design of the flow baffle.

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

The invention discloses an offshore rocket launching heat flow field test method and system, and the method comprises the following steps: S1, collecting multi-source heat flow data on an offshore rocket launching platform, and carrying out the synchronous processing, and generating a heat flow test data set; s2, constructing a sparse auto-encoder model; s3, optimizing model structure parameters by utilizing a king butterfly optimization algorithm to obtain an optimal parameter combination; s4, processing the multi-source heat flow data by applying the optimal structure parameters, and outputting a two-dimensional heat flow distribution diagram; and S5, identifying the heat flux area and the potential thermal anomaly, and outputting thermal shock early warning information. According to the method, the sparse auto-encoder and the empire butterfly optimization algorithm are fused, so that high-precision modeling and thermal shock early warning of the maritime rocket launching heat flow field are realized.

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 discloses a method and a system for calibrating a jerk error term coefficient of a PIGA and a medium, belongs to the technical field of rocket launching, and aims to calibrate # imgabs0 # by exciting jerk error term output in a uniform acceleration stage of a centrifugal machine. PIGA output is used as observed quantity, and the calibration precision of the jerk error coefficient is simulated and analyzed. Analysis shows that in the specific application of rocket launching, the circular probability error CEP of a rocket drop point when the # imgabs 1 # is calibrated is two orders of magnitude smaller than the circular probability error CEP of a rocket drop point when the # imgabs 2 # is not calibrated, and the precision and practicability are improved.

The invention discloses a device and a method for dynamically monitoring a weak part in a rocket simulation launching process, belongs to the technical field of detection, and solves the technical problem that the deformation condition of the weak part on a rocket shell cannot be remotely and dynamically monitored in real time in the scene of the first few seconds of rocket launching by combining a traditional speckle marking method and a DIC technology. According to the solution, a quantum dot paste matrix is pasted on the surface of a weak component on a rocket shell, and after quantum dot pastes are excited by light to emit light within a certain time, position changes of the light-emitting quantum dot pastes are tracked and observed through a visual signal acquisition system arranged at a proper long-distance position. The quantum dot technology and the DIC technology are combined, the surface deformation degree and structural reliability of weak parts on a rocket shell can be monitored and analyzed in time under the conditions that the rocket bears large thrust and has large acceleration in the first few seconds of rocket launching, the rocket research and development period is shortened, and the safety and reliability of the rocket are improved.

The utility model relates to the technical field of aerospace, in particular to a servo actuator zero position adjusting tool which comprises a guide rail, a plurality of sliding blocks, a plurality of sets of measuring pins and a vernier caliper. The upper surface of the guide rail is provided with a downwards-concave sliding groove, and each sliding block is located in the sliding groove and slides in the sliding groove in the extending direction of the sliding groove. Each group of measuring pins is provided with two same pin shafts, the lower end of one pin shaft is fixed to the upper surface of one sliding block, the upper end of each pin shaft vertically extends upwards, and the two pin shafts of each group of measuring pins are used for clamping the two ends of a servo actuator of a rocket; one end of the vernier caliper is fixed to the outer side surface of one pin shaft of one group of measuring pins, and the other end of the vernier caliper is clamped to the outer side surface of the other pin shaft of the group of measuring pins. According to the invention, the accuracy of zero distance measurement of the servo actuator of the rocket can be ensured, so that the success of rocket launching and aerospace tasks is ensured.

A database (102) of a rocket launch assistance device records orbit forecast information of a mega-constellation satellite group (301) that is acquired from a space information recorder included in a mega-constellation business device and space object information that is acquired from a rocket launch business device. The orbit forecast information of the mega-constellation satellite group (301) is composed of a prediction value of an orbit of at least one representative satellite (331) out of the mega-constellation satellite group (301) and a prediction value of an orbit of a constituent satellite (332), which is a value relative to the prediction value of the orbit of the representative satellite (331).

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 present invention discloses a noise reduction and flow guiding device for a launch ship guide trough and a design method, which belongs to the field of rocket launch technology, wherein the noise reduction and flow guiding device for a launch ship guide trough comprises a curved guide plate and a fixed bracket; the bottom surface of the guide plate is fixedly connected to the side edge of the guide trough outlet arranged on the launch ship; the lower end of the guide plate is smoothly transitioned to the edge of the guide trough outlet; the projection of the upper end of the guide plate in the vertical direction at least covers a part of the guide trough; the fixed bracket is fixedly arranged on the launch ship and fixedly connected to the guide plate. The present invention can change the direction of the gas jet at the outlet of the guide trough, thereby reducing the influence of the vibration noise generated by the jet tail flame on the rocket body.

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 demonstration device for simulating rocket launching, which comprises a rocket body, a launching platform, a leveling device, an adaptive device and a guide post, and the side part of the rocket body is provided with a guide pipe extending along the vertical direction; the launching platform is provided with a bearing surface for bearing a rocket body, and the leveling device is connected with the launching platform and used for leveling the bearing surface; one end of the adaptive device is fixedly connected with the launching platform, and the other end of the adaptive device is telescopically arranged in the horizontal direction towards or away from the rocket body; the guide column is installed at the end, close to the rocket body, of the adaptation device, and at least part of the guide column extends into the guide pipe. Through the arrangement, the leveling device can level the bearing surface for bearing the rocket body, so that the rocket body is in a horizontal and vertical state before being launched, the guide column is matched with the guide pipe, the rocket body can be guided vertically upwards before being launched, and the rocket body is further prevented from inclining after being launched.

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 provides a liquid rocket test run launching integrated launching pad, and belongs to the technical field of carrier rocket launching devices.The liquid rocket test run launching integrated launching pad comprises a launching pad body, a one-way fluid director, a switching device and a lifting device.In the test run state, a lifting platform of a lifter is located below the table top of a launching platform; and in a launching state, a lifting platform of the lifter extends out of the table top of the launching platform and is in contact with the L-shaped split flange plate. According to the liquid rocket test run launching integrated launching pad, a traditional test run platform and a launching pad of a launching site are integrally designed, and on the premise that the test function is met and the flight mission risk is reduced, the construction cost of the test run platform can be greatly saved, the transportation and transfer links are reduced, the transportation cost is saved, and the total launching period is shortened. Due to the overall modular design of the integrated launching pad, the launching pad has migration maneuverability and can meet the task requirements of different launching sites, the launching tasks of multiple sites are achieved through one set of equipment, and cost is saved.

The invention belongs to the technical field of rocket launching, and particularly relates to a rocket fairing noise reduction method and device and a storage medium. The method comprises the following steps: acquiring a noise signal; performing spectral analysis on the acquired noise signal by adopting fast Fourier transform to obtain a frequency domain signal, the frequency domain signal comprising amplitude and phase information under different frequency components; setting a boundary of each frequency band, extracting feature information of each frequency band, and extracting noise components of each frequency band according to the feature information of each frequency band; the noise reduction strategy of each frequency band is optimized through an artificial bee colony algorithm, and the optimal frequency and amplitude of the anti-phase sound waves of each frequency band are obtained; and according to the output of the artificial bee colony algorithm, anti-phase sound waves are generated, and active noise reduction is carried out on the fairing. According to the method, the overall noise reduction performance of the fairing can be improved, the response speed is high, the requirement for computing resources is low, and the method is suitable for scenes with strict requirements for time and resources in rocket launching scenes.

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 utility model discloses a variable-rigidity high-reliability shock absorber which comprises a gasket, a low-rigidity shock absorption ring, a high-rigidity shock absorption ring, a low-rigidity shock absorption column, a high-rigidity shock absorption column and a sleeve. The section of the sleeve is in an inverted-T shape, the sleeve comprises a hollow cylinder and a circular ring fixed to the lower portion of the hollow cylinder, the hollow cylinder is sleeved with the low-rigidity vibration reduction column and the high-rigidity vibration reduction column, the outer diameter of the low-rigidity vibration reduction column, the outer diameter of the high-rigidity vibration reduction column and the inner diameter of a supporting lug hole of a vibration-reduced object are the same, and the outer diameter of the low-rigidity vibration reduction column is larger than the inner diameter of the supporting lug hole of the vibration-reduced object. The low-rigidity vibration reduction column and the high-rigidity vibration reduction column are installed in a supporting lug hole of an object subjected to vibration reduction in a combined mode, and the gasket and the sleeve are installed in a combined mode. The shock absorber disclosed by the utility model has the characteristic of variable stiffness, can adapt to different working conditions, and can meet the comprehensive shock absorption requirements of small load in orbit of space equipment and large load launched along with a rocket, and the comprehensive shock absorption requirements of small load of airborne weapons in hanging flight along with an airplane and large load in autonomous flight after launching.

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 belongs to the field of rocket seaborne launching, and particularly relates to a rocket seaborne launching ship stabilization auxiliary device which is characterized in that the rocket seaborne launching ship stabilization auxiliary device comprises two telescopic rods installed on the two sides of a ship body, and the two telescopic rods I are arranged on the two sides of the ship body; the output shafts of the two first telescopic rods located on the same side are connected to the same auxiliary stabilizing mechanism through lifting blocks. The auxiliary stabilizing mechanism comprises a U-shaped mounting frame, the bottom end of the interior of the mounting frame is rotationally connected with a buoyancy compartment through a rotating shaft, a second telescopic rod driving the buoyancy compartment to rotate around the rotating shaft is further mounted between the buoyancy compartment and the mounting frame, and storage grooves are formed in the two sides of the buoyancy compartment. The device is unfolded before the rocket launching ship needs to load a rocket and launch the rocket, the stability of the rocket launching ship can be remarkably improved, a more stable bearing platform is provided for follow-up rocket installation and launching, the device can be completely folded in the normal running process of the rocket launching ship, and the ship body flexibility is high.

A shock testing machine including: a carriage for holding a component to be tested, the carriage being configured to be movable in a linear direction along one or more elongated rails; a drum rotatable on a shaft, the drum having a circumferential surface; a cable having one end connected to the carriage and an other end connected to the drum; a motor having an output connected to the shaft to rotate the drum under the motive power of the motor to wind the cable on the circumferential surface of the drum; and a clutch disposed in a power train operatively connecting the motor to the drum, the clutch having a disengaged state and an engaged state. Where the motor is controlled to disengage the clutch while the motor reaches a predetermined rotational speed or predetermined rotational torque and to engage the clutch when the motor reaches the predetermined rotational speed or predetermined rotational torque to accelerate the carriage and component to be tested in the linear direction.

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 invention relates to an air cylinder, in particular to a reliable backward-falling and rebound-free backward-falling buffer cylinder for rocket launching, which comprises an air cylinder body, an air cylinder piston rod, an air cylinder piston, a spring, a buffer cylinder body, a buffer piston, a throttle lever and a connecting piece, the buffer cylinder body, the buffer piston and the throttle lever are all located in a rodless cavity of the air cylinder body, the center line of the air cylinder piston, the center line of the throttle lever and the center line of the buffer piston are located on the same straight line, the rodless cavity of the buffer cylinder body is filled with hydraulic oil, and a pressure measuring opening is formed in the buffer cylinder body. A through hole is formed in the end portion of one end, located in the buffer cylinder body, of the buffer piston, the end portion of the other end of the throttling rod is arranged in the through hole, a gap is reserved between the through hole and the throttling rod, and the gap forms an annular throttling oil channel. The supporting arm can be driven to reliably and rapidly fall backwards after a rocket is ignited, and rebounding is avoided after falling backwards.

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.