1529 results about "Rocket" patented technology

The invention provides a reusable rocket landing stage starting height prediction method, a guidance system and a processor, and aims to dynamically and accurately determine the ignition time of an engine by constructing a real-time online starting height prediction model based on an energy conservation principle and combining multiple safety thresholds (including an allowable starting height upper limit and a standby starting height lower limit) so as to predict the starting height of the rocket in a landing stage. And the safety and fuel optimality in the whole flight envelope are ensured, so that the reliability, accuracy and economical efficiency of the reusable rocket landing are remarkably improved. The problems that in the prior art, engine starting height prediction precision and robustness are insufficient, online adaptation to uncertainty of a complex flight environment cannot be achieved, and safety margin consideration is insufficient are solved, and potential safety risks are thoroughly eliminated.

Ejection seat systems include an ejection seat and a propellant assembly mounted to the ejection seat. A nozzle assembly having an adjustable nozzle is moveably attached to the ejection seat and configured to receive propellant exhaust from the propellant assembly and exhaust through the adjustable nozzle. A first linear electric actuator is operably connected to the adjustable nozzle and arranged to adjust an orientation of the adjustable nozzle in a first direction and a second linear electric actuator is operably connected to the adjustable nozzle and arranged to adjust an orientation of the adjustable nozzle in a second direction. The second direction includes a directional component that is normal to the first direction. A control module is configured to control operation of the first linear electric actuator and the second linear electric actuator.

The invention discloses an offshore rocket erection state anomaly detection method based on sparse representation and adaptive filtering, and the method comprises the following steps: obtaining state data in a rocket erection process through multiple sensors, constructing a time sequence, setting a sliding window with change capability in each time period, and extracting sparse features for state prediction. A feedback adjustment signal is formed by calculating the difference between a predicted value and an observed value, and parameters in the state estimation process and the external disturbance modeling process are adjusted respectively. A double-path structure is adopted to independently estimate the state and disturbance, and a final estimation result is obtained through a fusion strategy. The system continuously updates parameters according to error conditions, continuous self-adaptive adjustment and anomaly recognition are achieved, and the recognition precision and the response capability of the anomaly trend are improved. The method is suitable for the recognition processing of the abnormal state in the erection process of the rocket in the marine environment, and has high dynamic adaptive capacity and time sequence anomaly detection capacity.

The invention provides a high-precision optical pose measurement method based on tower recovery, which adopts a close-range bidirectional optical cooperation mode, measures a mark point of a recovery tower through an optical detector on a rocket, and determines the distance and the pose of the rocket relative to the recovery tower; and the optical measurement system of the recovery tower synchronously detects and identifies the rocket and controls a capture arm and a buffer mechanism of the recovery tower to move, and optical measurement data of the recovery tower are fed back to the rocket and are integrated with optical measurement data on the rocket to control landing of the rocket. According to the invention, optical measurement and navigation are introduced into rocket tower type recovery, and the method has the advantages of high measurement precision, high data frequency, interference resistance and the like. The relative relation between the rocket and the recovery tower can be directly obtained through optical pose measurement, compared with inertial navigation, the coordinate conversion process can be reduced, and the advantage of no accumulated error is achieved. Compared with satellite navigation, optical pose measurement does not depend on external signals, and higher independence and electromagnetic resistance are achieved.

The invention discloses a space-based telemetering transmission system for a carrier rocket, and relates to the technical field of rocket measurement and control communication, and the system comprises a rocket-borne network camera which is used for collecting video image data of a preset monitoring point position of the rocket, compressing the video image data, and outputting the compressed video image data as telemetering data; the telemetering transmission equipment is respectively connected with the rocket-borne network camera and the ground equipment, performs telemetering framing on the telemetering data, outputs first telemetering data and outputs second telemetering data to the ground equipment; the space-based power amplifier unit equipment is connected with the telemetering transmission equipment, receives the first telemetering data, amplifies the first telemetering data and then outputs the first telemetering data through radio frequency; the space-based telemetering antenna is connected with the space-based power amplifier unit equipment, receives the first telemetering data and then transmits the first telemetering data to a relay satellite, so that the satellite sends the first telemetering data to a ground measurement and control station; and the ground equipment is used for receiving the second telemetry data and then analyzing and uploading configuration parameter information. The problem that the coverage blind area of a telemetering system mainly based on ground-based measurement and control is large in the prior art is solved.

The invention provides multichannel spinning and turning integrated equipment for a cylindrical part of a carrier rocket and a machining method. The multichannel spinning and turning integrated equipment comprises a lathe bed, a main spindle box, a closing-in spindle box, a spinning roller frame, a tail top, a heating system, a control system and a hydraulic system. The main spindle box, the closing-in spindle box, the spinning roller frame and the tail top are installed on the lathe bed. The main spindle box is connected with the closing-in spindle box and drives the closing-in spindle box to work, a cylindrical part blank is clamped between the closing-in spindle box and the tail top, the spinning roller frames are located on the two sides of the cylindrical part blank, and the heating system is installed on the spinning roller frames; the control system is connected with the main spindle box, the closing-in spindle box, the spinning roller frame, the heating system and the hydraulic system. The hydraulic system is connected with the main spindle box, the closing-in spindle box, the spinning roller frame and the tail top. Through cooperation of the main spindle box, the closing-in spindle box, the spinning roller frame, the tail top and the like, the in-situ machining function combining forced spinning machining, closing-in machining and turning machining of cylindrical parts, conical parts and curved-bus-shaped parts of the carrier rocket is achieved.

The system and method for decelerating space objects moving at high velocities are provided. The system comprises at least one closed volume containing a combination of gas, liquid, solid particles, or a mixture thereof. The volume can be transported to the targeted location using a range of means, including a chemical gun, light gas gun, electromagnetic coil gun, superconducting quench gun, a rocket, or a combination thereof. The volume is strategically positioned on the trajectory of the moving object. Upon penetrating the walls of the volume, the object passes through it, experiencing deceleration.

The invention relates to an industrial CT composite material defect reconstruction method based on an unsupervised neural network. The method comprises the following steps: carrying out industrial CT shooting to obtain a cross section two-dimensional image of a composite material component; an unsupervised model for converting an industrial CT composite material image into a defect model for reconstruction and discrimination is constructed, sampling is performed through a multi-scale sampling model, a channel attention mechanism is introduced for weight adjustment, and a multi-scale residual module is introduced for reconstruction; constructing a joint loss function for calculating the unsupervised model; training the constructed unsupervised model by using a data set, wherein a loss function is used for updating parameters; and inputting a to-be-detected industrial CT composite material image into the model, so that three-dimensional reconstruction of the defects of the composite material can be completed. According to the method, the defects of the composite material are reconstructed with high precision, nondestructive testing of the composite material is achieved, and the method can be used for structural health monitoring and defect detection of aerospace vehicles such as airplanes and rockets.

The invention discloses a satellite angular rate identification method and related equipment, and relates to the technical field of rocket launching, and the method comprises the steps: obtaining a rocket telemetry video frame sequence after satellite-rocket separation; for each frame of image of the rocket telemetry video frame sequence, determining a dynamic image area of the target satellite based on the static area mask; based on a feature detection algorithm, binary descriptors are extracted from dynamic image areas of two adjacent frames of images; based on the binary descriptor, matching corresponding feature point pairs between two adjacent frames of images; estimating a rotation matrix between two adjacent frames of images through an iterative optimization algorithm based on the corresponding feature point pairs; and determining the angular rate of the target satellite according to the rotation matrix and the time interval between the two adjacent frames of images. According to the method, rapid and high-precision identification of the satellite angular rate is realized, the problem that the traditional technology depends on indirect data and is high in delay is solved, task success or failure can be judged in real time, and the launching task evaluation efficiency is improved.

The embodiment of the utility model provides a storage tank bottom structure, a storage tank and a carrier rocket, the storage tank bottom structure comprises a shell and a top cover, the shell is of a hollow arc-shaped revolving body structure which is through in the front-back direction, the two ends of the shell are the small-diameter end and the large-diameter end respectively in the axis direction of the shell, and the small-diameter end of the shell is fixedly connected with the top cover; a propellant is arranged on the top cover; a plurality of protruding structures are evenly distributed on the outer side face of the shell in the circumferential direction in the axis direction of the shell. According to the technical scheme, the bottom structure of the thrust bearing type storage tank is simple in process, easy to manufacture in batches and beneficial to achieving light weight of a carrier rocket.

The invention relates to the technical field of verification and evaluation, in particular to a reusable rocket attitude control system redundancy verification and fault tolerance evaluation method, which comprises the following steps: performing multi-dimensional combination on disturbance units according to disturbance types, action time periods and intensity grades to form a disturbance combination set, constructing a group of induction task profiles covering different attitude control abnormal situations; recording a switching state of an internal control path of the attitude control system, a redundant component intervention time sequence and a function takeover sequence in real time, and constructing a dynamic response path map of a redundant link; and generating a response integrity index and a fault tolerance conflict degree criterion of redundancy configuration as a fault tolerance capability evaluation result of the attitude control system. Compared with an existing method which only verifies individual static disturbance scenes, the method has the advantages that systematic design and composite excitation of disturbance conditions can be realized, scene complexity and coverage of redundancy verification are enhanced, and authenticity and representativeness of fault-tolerant testing are improved.

The invention relates to a self-adaptive wind load passive slowing-down design method based on threshold triggering. The method comprises the steps that aerodynamic parameters during non-load-shedding are calculated according to a nominal program angle; calculating an initial load shedding program angle ideal value; when the flight Mach number is within the allowable load shedding range and the qalpha value exceeds the current amplitude limiting value, load shedding is carried out; only the part exceeding the qalpha amplitude limiting value is subjected to load shedding; and carrying out amplitude limiting processing on the angle correction of the load shedding program in combination with the dynamic characteristics and maneuverability of the rocket body. According to the method, the load shedding opportunity is triggered through the qalpha amplitude limiting value, amplitude limiting processing is carried out on the load shedding program angle and the angular velocity correction amount, the structural strength of a rocket body can be fully utilized, the dynamic lag characteristic and the actual control capacity of the rocket are considered, the rocket body can track the corrected program angle with high precision, and the method has high application value in engineering.

Combined aerospace system is a multi-stage technology to deliver payloads to orbit and to midair. The system comprises a carrier aircraft, a tow aircraft, a towed spaceplane or a rocket plane or a glider. The carrier aircraft, coupled with the spaceplane and / or the rocket plane or several gliders, takeoff from a runway and ascends to a staging altitude and speed under the carrier aircraft power. After staging, the tow aircraft will pick up the spaceplane or the rocket plane or the glider from the carrier aircraft in flight. The towed spaceplane or the rocket plane or the glider continues ascending and accelerating by the tow aircraft. The spaceplane or the rocket plane will disconnect from the tow aircraft and continue to ascend and accelerate in an atmosphere by using the ramjet engine. The spaceplane or the rocket plane equipped with a propulsion unit is used to place the spacecraft into the earth orbit. The glider will disconnect from the tow aircraft and lands in designated place. Gliders can be used for instantly delivering of correspondence, parcels, and civil goods from a supplier to wide variety of individual consumers or corporate customers.

The invention discloses a rocket aircraft aerodynamic performance prediction method based on heaven and earth test and simulation data, and relates to the technical field of aerodynamic design. The method mainly comprises the following steps: obtaining a large number of aerodynamic performance parameters by using a non-viscous rapid design theory to construct a low-precision flow field training data set; performing heaven-earth correlation error correction on ground wind tunnel test data through a flight test; performing simulation error correction on a large number of aerodynamic performance parameters obtained by CFD numerical simulation through test data to obtain a high-precision training data set; and performing data fusion on the high-precision data set and the low-precision data set, and constructing a pneumatic prediction model. The invention provides the method for rapidly predicting the aerodynamic performance of the rocket aircraft, and an efficient and feasible solution is provided for rapid prediction of the aerodynamic performance.

The invention relates to the field of rocket recovery pose estimation design, and particularly discloses a multi-mode cooperative perception driven rocket recovery high-precision pose estimation system, which comprises a multi-scale Marker layout module, a multi-scale cooperative perception driven rocket recovery high-precision pose estimation module and a multi-scale cooperative perception driven rocket recovery high-precision pose estimation module, and is characterized in that the multi-scale Marker layout module is used for arranging three-level height detection bands around a recoverer device, and each band is configured with Marker type layout modes of different sizes; the multi-view cooperative detection module is used for collecting images through a plurality of high-resolution cameras installed on the side face of the rocket and outputting visual poses in combination with scale adaptive detection process optimization and a space consistency constraint strategy; the multi-sensor fusion module is used for performing fusion optimization on the visual pose, the IMU and the GPS data through a factor graph to construct a factor graph joint optimization framework; accuracy and stability of a rocket recovery pose estimation system are realized, and real-time performance, accuracy and anti-interference capability of a rocket recovery pose estimation scheme are improved.

An expanding fin aerial rocket in combination with a sleeve for restraining the fins prior to deployment. The rocket has an elongate body with a fore end for a warhead and an opposed aft end for motor and guidance fins. The guidance fins are hingedly attached to the rocket body at a proximal end and extend outwardly to a distal end. Fins may be collapsed during storage to conserve space and prevent damage. The springs are outwardly spring biased for flight upon deployment. The rocket has a durable removable sleeve, which fits over the fins during storage, to retract the sleeves into a collapsed position. The sleeve is removed for deployment, so the fins may radially expend to an extended position. The durable sleeve may then be reused for another rocket.

The invention discloses supporting equipment for rocket transportation and a using method thereof, and belongs to the technical field of rocket transportation, the supporting equipment comprises a supporting base and a rocket supporting mechanism, a nitrogen spring mechanism is arranged between the supporting base and the rocket supporting mechanism, and the two ends of the nitrogen spring mechanism are connected with the supporting base and the rocket supporting mechanism correspondingly; the nitrogen spring mechanism comprises a nitrogen spring body and a nitrogen spring shell, the nitrogen spring body is arranged in the nitrogen spring shell, a locking assembly is arranged on one side of the nitrogen spring shell and connected with the nitrogen spring body, and the nitrogen spring body is locked in position through the locking assembly. According to the rocket transportation support, elastic and rigid switching can be achieved, buffering can be provided in an elastic mode, stable supporting can be provided in a rigid mode, stable fixing of a rocket body is ensured, according to the characteristics of rocket transportation, elastic and rigid supporting switching is actually used, the rocket transportation stability is improved, and meanwhile the rocket transportation safety is improved. And the safety effect in the rocket transportation process is greatly improved.

The invention discloses a shell for a rocket power device and a machining method, and the machining method of the shell comprises the steps that a mandrel is arranged, a foaming core mold is manufactured on the periphery of the mandrel, and the shape of the outer surface of the foaming core mold is the same as that of the inner surface of the shell to be machined; winding a shell material on the outer surface of the foaming core mold to form a shell on the surface of the foaming core mold; and the core shaft is removed, and the foaming core mold in the shell is removed. According to the shell machining method, the core mold blank is formed at a time through foaming injection molding with the core shaft as the standard, the core mold is removed after the shell is formed in the shell material winding machining mode, the machining mode is higher in efficiency and suitable for batch production, and the technical problems that in the related technology, the core mold is large in weight and high in cost are solved.

The invention relates to a reusable rocket landing section fuel optimal conversion time sequence online generation method and device and a medium, and the method mainly comprises the steps: simplifying the complex three-dimensional motion of a rocket vertical landing section into a one-dimensional vertical motion model; establishing a nonlinear equation set which takes the working time of the engine before and after stage conversion as an unknown quantity and enables the rocket to meet the terminal speed and terminal position constraints and the fuel consumption to be optimal; constructing and solving an approximate analytical solution as an initial value of Newton iteration; and based on the obtained initial value, carrying out iterative solution on a double-constraint nonlinear equation set formed by rocket speed and displacement terminal constraint. According to the method, a form based on analytic iteration is adopted, the fuel optimal control in the rocket landing process is realized through real-time online calculation of an engine conversion time sequence and giving of a corresponding thrust adjustment control instruction, and safe and accurate landing is ensured.

The invention provides a recovery real-time positioning method based on an inertial navigation-vision method, and the method comprises the steps: setting an on-rocket vision sensing unit, enabling the on-rocket vision sensing unit to obtain a cooperative target image, fusing the measurement results of an inertial navigation measurement system and a vision pose measurement system, carrying out the mutual correction of measurement errors, and outputting a fusion result. The inertial navigation-vision fusion technology is adopted, the problem that inertial navigation measurement errors are accumulated along with time is solved through the vision pose measurement technology, and meanwhile the inertial navigation system is used for making up for the defect in the aspect of vision navigation stability. A recovery site cooperative target is measured through a visual detector on the rocket, pose information of the rocket relative to the recovery site is determined, and inertial navigation measurement error correction is completed by using the result; and meanwhile, high-precision attitude information provided by the inertial navigation system is used for correcting a measurement error caused by jittering of the visual attitude measurement system, so that the precision and robustness of rocket attitude measurement are improved.

The invention relates to the technical field of rockets, in particular to a supporting device for a rocket and a transferring and erecting integrated launching vehicle. The supporting device comprises an erecting arm body and a plurality of auxiliary supporting mechanisms. The auxiliary supporting mechanism comprises a supporting bracket, a disc spring force limiter and at least two linear guide rails. The pre-pressure of the disc spring force limiter is adjustable, so that the rigidity of the supporting bracket is matched with the rigidity of the rocket body; the linear guide rails are arranged on the two sides of the supporting bracket respectively. According to the supporting device for the rocket and the transferring and erecting integrated launching vehicle, the supporting bracket is kept to be in a balanced state through the linear guide rail, the floating function is achieved, and the clamping stagnation phenomenon is prevented; force limiting supporting is achieved by adjusting the pre-pressure of the disc spring force limiter, the rigidity of the supporting bracket is made to be close to the rigidity of a rocket body, and therefore it is guaranteed that the additional bending moment cross section load generated by the erecting arm to the rocket is within the allowable range.

The invention provides composite spinning equipment for forming rocket pipe type structural parts. The composite spinning equipment comprises a structural system, an operation table, an oil cooler, an oil tank, a control cabinet, a water cooler, a heating system, a gas distribution cabinet, a hydraulic station, an oxygen bottle, a circulating filtering device and a propane gas bottle. The structural system comprises a main spindle box, the main spindle box has active power and can achieve controllable rotation, a closing-in spindle box does not have active power, and when closing-in spinning needs to be conducted, the main spindle box and the closing-in spindle box are connected through a coupler, and controllable rotation of the closing-in spindle box is achieved; the water cooling machine is used for cooling the spinning roller in the spinning process; the heating system can heat the workpiece to realize hot spinning; the circulating filtering device sprays cutting fluid to the workpiece, circulating filtering use of the cutting fluid is achieved, and wet-type forced rotation is achieved. The hot spinning and wet spinning device has the integrated hot spinning and wet spinning effects of strong spinning, close-up common spinning and the like.

The utility model discloses an unmanned aerial vehicle rocket launcher which comprises a mounting plate and a launcher hanging frame arranged on the mounting plate, a connecting rod I is hinged to one end of the top of the launcher hanging frame, and a connecting rod II is hinged to the other end of the top of the launcher hanging frame. The free end of the connecting rod I is fixedly connected with the mounting plate, the free end of the connecting rod II is connected with the mounting plate through a connecting seat, and the connecting rod II is hinged to the connecting seat; a sliding rod of a T-shaped structure is arranged on the connecting base, and a T-shaped sliding groove allowing the sliding rod to stretch into is formed in the mounting plate. A driving motor is arranged on the connecting seat, a groove is formed in the bottom of the mounting plate, a rack is arranged in the groove, and a gear is arranged at the output end of the driving motor; the length of the connecting rod II is larger than that of the connecting rod I. According to the utility model, the angle of the launching cylinder is adjusted through the angle adjusting assembly, so that the launching angle of the fire extinguishing bomb is controlled, the fire extinguishing bomb is launched at different angles, the striking precision is improved, and the fire extinguishing effect is enhanced.

The invention discloses a damage tolerance evaluation method and system for a friction plug repair welding joint of a carrier rocket storage tank, and the method comprises the steps: detecting an interface of the friction plug repair welding joint of the carrier rocket storage tank, and obtaining defect information, constitutive parameters of a material, and material attributes; establishing a finite element model, and accurately setting corresponding distribution interface defects; quantizing in-plane and out-plane restraint effects of the crack tip by adopting restraint parameters; processing and preparing a unilateral notch tensile sample, and measuring a local J-R curve of the tip of the precrack at the interface of the friction plug repair welding joint; correcting a real J-R curve of the friction plug repair welding joint of the rocket tank based on the influence of the distribution interface defect characteristics; then the tensile strength values of the friction plug repair welding joint in all the positions in the circumferential direction are corrected, and a corrected real J-R curve is obtained; and the limit crack size under the action of the corresponding internal pressure is obtained by utilizing the crack tip extension driving force curve, and then the critical J integral of the crack is calculated, so that the maximum crack size under the specific load or the maximum internal pressure under the crack with the specific size is obtained.

The embodiment of the invention provides a rocket vertical recovery method and system and a rocket, and the method comprises the steps: after a first-stage rocket is separated, the thrust of a central engine of the first-stage rocket is adjusted to a preset proportion of rated thrust so as to provide positive overload for a propellant in a storage tank; the propellant can be gathered to an outlet in the bottom of the storage tank through positive overload; the flying direction of the first-stage rocket deviates from the current flying direction by performing bidirectional swinging on the central engine, and the posture of downward diving of the first-stage rocket towards the earth direction is kept; tangential force in two different directions tangent to the current flight direction of the first-stage rocket can be generated at the same time by swinging the center engine in two directions, the posture of the first-stage rocket is adjusted to the upward flight posture from the downward diving posture, and speed reduction is conducted till the first-stage rocket lands on the earth. The thrust of the center engine of the first-stage rocket is adjusted to be small, and attitude control and propellant management of first-stage rocket recovery are achieved.

The invention discloses an offshore rocket self-lifting platform dynamic scheduling method based on cooperative intelligence, and the method comprises the following steps: constructing a plurality of platforms into an intelligent unit capable of autonomously responding to a task, and carrying out the dynamic marshalling through a cooperative mechanism, so as to achieve the self-adaptive matching between the task and the platform. In the method, multiple indexes are introduced as scheduling targets, weight parameters are dynamically adjusted according to real-time tasks, resources and environment information, and the scheduling flexibility is improved. And generating a scheduling strategy through a learning optimization mechanism, and controlling platform task allocation, path planning and execution time sequence. In the execution process, a scheduling strategy and a platform marshalling structure can be updated in time according to environment disturbance and task state changes, and smooth execution of tasks and stable cooperation of the platform are ensured. The method is suitable for a multi-task and high-frequency emission scene.

The invention discloses a fixed-wing unmanned aerial vehicle rocket separation mechanism which comprises a boosting limiting device arranged on a boosting rocket, the boosting limiting device comprises a boosting limiting block fixedly connected with the bottom of an unmanned aerial vehicle and provided with a through hole, and a boosting limiting taper pin fixed in the through hole, and one end of the boosting limiting taper pin forms a conical end face; the boosting limiting device further comprises a conical groove which is formed in the boosting rocket and matched with the conical end face. In the boosting process, the conical end face and the conical groove form surface contact and abut against and are attached, it is guaranteed that thrust is evenly transmitted, the stress concentration risk of the fuselage is reduced, meanwhile, interference of non-directional airflow in the boosting process is counteracted, and the situation that the attitude of the unmanned aerial vehicle is out of control is avoided. And after boosting is completed, unpowered natural separation of the unmanned aerial vehicle and the boosting limiting device is achieved through the resistance difference of air to the unmanned aerial vehicle and the boosting limiting device, the conical groove guides the conical end face, additional driving parts are not needed, the structure is simplified, and the flight safety of the separated unmanned aerial vehicle is improved.

A reusable upper stage rocket or other atmospheric re-entry vehicle includes a nose, a base opposite the nose, and a propulsion engine toward the base. The propulsion engine includes a high pressure chamber and a nozzle configured to exhaust gas generated by the high pressure chamber. The nozzle includes an initial nozzle portion, a secondary nozzle portion downstream of the initial nozzle portion, and a nozzle exit at a downstream end of the secondary nozzle portion. The secondary nozzle portion includes an inner expansion surface, an outer expansion surface outboard of the inner expansion surface, and an expansion cavity defined between the inner expansion surface and the outer expansion surface.

The invention relates to a multi-degree-of-freedom adjustable unmanned aerial vehicle booster rocket mounting trolley, and belongs to the technical field of unmanned aerial vehicle launching. Comprising a base, a lifting mechanism is arranged on the base and comprises a lifting plate, a longitudinal moving mechanism is arranged at the upper end of the lifting plate and comprises a longitudinal moving plate horizontally sliding in the Y direction, and a transverse moving mechanism is arranged at the upper end of the longitudinal moving plate and comprises a transverse moving plate horizontally sliding in the X direction; a rolling mechanism is arranged at the upper end of the transverse moving plate through a pitching mechanism, the rolling mechanism is driven by the pitching mechanism to rotate in the X direction, the rolling mechanism comprises a fixed support and a rotary support, the rotary support is rotationally arranged on the fixed support, a hoop mechanism is arranged on the rotary support, and the hoop mechanism is arranged on the fixed support. The booster rocket is fixedly connected with the rotary bracket through a hoop mechanism; the problem that a booster rocket of an existing large unmanned aerial vehicle cannot be hoisted in the docking process and cannot be accurately docked with the unmanned aerial vehicle is solved.

The invention discloses an artificial influence weather rocket projectile three-explosion self-destruction body assembly detection system, and relates to the technical field of multivariable comprehensive measurement, and the system comprises a multi-dimensional positioning measurement module which is used for measuring the coaxiality and fitting pressure of a self-destruction body and an engine housing; the multi-parameter measuring module is filled with the medicament to measure the weight, density and internal defects of the medicament; the connection strength dynamic measurement module is used for measuring torque, tensile strength and vibration displacement; the igniter performance parameter measuring module is used for measuring resistance, time delay and safety reliability; the measurement data integration analysis module is used for integrating a data generation report and an encryption log; and the system control and exception handling module is used for coordinating module time sequences, monitoring data and handling exceptions. According to the system, multi-parameter integrated measurement of the three-explosion self-destruction body is realized, the measurement precision of positioning, medicament filling, connection strength and igniter performance is improved, the remain after self-destruction is ensured to be compliant, and the data is traceable.