1184 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 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 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 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 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 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 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 discloses an artificial influence weather rocket projectile three-explosion self-destruction body assembly detection system, and relates to the technical field of multivariable comprehensive measurement, and the system comprises a multi-dimensional positioning measurement module which is used for measuring the coaxiality and fitting pressure of a self-destruction body and an engine housing; the multi-parameter measuring module is filled with the medicament to measure the weight, density and internal defects of the medicament; the connection strength dynamic measurement module is used for measuring torque, tensile strength and vibration displacement; the igniter performance parameter measuring module is used for measuring resistance, time delay and safety reliability; the measurement data integration analysis module is used for integrating a data generation report and an encryption log; and the system control and exception handling module is used for coordinating module time sequences, monitoring data and handling exceptions. According to the system, multi-parameter integrated measurement of the three-explosion self-destruction body is realized, the measurement precision of positioning, medicament filling, connection strength and igniter performance is improved, the remain after self-destruction is ensured to be compliant, and the data is traceable.

The invention discloses a reusable carrier rocket marine landing recovery device which comprises a tower body, a driving module, a first mechanical capturing arm, a second mechanical capturing arm, a first clamping arm and a second clamping arm. The tower body comprises first to fourth stand columns and first to eighth cross beams; the driving module comprises first to twelfth driving units; each of the first to twelfth driving units comprises a driving motor, a working track, a lead screw nut, a first bearing, a second bearing, a first bearing seat and a second bearing seat. On the premise that a simple electric control system is adopted, marine recovery of the carrier can be achieved, the structure is simple, response is rapid, good reliability and maintainability are achieved, the device is suitable for recovery tasks of carrier rockets of various models, and the launching cost is effectively reduced.

The invention discloses a mobile medium-large liquid carrier rocket capturing and recycling system, and relates to the technical field of rocket recycling. According to the mobile medium-large liquid carrier rocket capturing and recycling system, through the mobile platform design that a crane is adopted to carry a connecting truss, a switching frame and a capturing arm, rocket body information is obtained in real time in combination with a camera module, a rocket position sensor and the like, and interaction with a ground center is achieved through a data transmission module; the control module accurately controls the transmission assembly to adjust the capturing arm, the buffering assembly is matched to absorb impact, and a parachute landing mode is replaced to improve the control precision. According to the movable characteristic, the dependence on a fixed land or seaborne recovery site is reduced, the site construction and maintenance cost is reduced, the influence of environmental factors on recovery is reduced, and the problems that in an existing medium-large liquid carrier rocket recovery technology, parachute landing control is difficult, the land / seaborne recovery cost is high, and the influence of the environment is large are solved.

The invention discloses an offshore rocket precise docking path planning method based on a hierarchical Markov decision process, and the method comprises the following steps: S1, collecting the real-time state data of an offshore rocket and a recovery platform, and constructing an environment state space; s2, discretizing and coding the spatial region, initializing a parrot optimization algorithm, and generating an optimal path; s3, dividing paths in stages, mapping the paths into option sets, and constructing a hierarchical Markov model; s4, performing option scheduling based on the model, selecting options by a high-level controller, and outputting a control action by a low-level controller; s5, state change is monitored in real time, path deviation is evaluated, and path re-planning is triggered; s6, disturbance change drives low-layer control strategy adjustment, and model strategy parameters are updated; and S7, recording feedback data, and incrementally updating the state transition probability and the reward function. The invention also discloses a method. According to the invention, self-adaptive high-precision path planning and docking control of the marine rocket in a complex disturbance environment are realized.

The invention discloses a reusable carrier rocket ground active recovery system and recovery method.The system comprises a landing buffer mechanism, an omnidirectional self-driven carrying platform, a rocket body posture auxiliary stabilizing mechanism and a landing structure, the landing structure comprises bearing feet, the bearing feet are located on a rocket, and a positioning device is arranged on the omnidirectional self-driven carrying platform; the rocket body posture auxiliary stabilizing mechanism comprises two pairs of clamping rods which are arranged in parallel in a staggered mode, the two ends of each clamping rod are located on a sliding table, the clamping rods are driven by the sliding table to move, and the sliding table is located on a sliding rail; the landing buffering mechanism is located on the omni-directional self-driven carrying platform and comprises a force bearing buffering base and a flexible contact pad located on the force bearing buffering base. The rocket recovery control difficulty is reduced while the carrying capacity of the rocket is guaranteed, and the rocket recovery control device is reasonable and stable in structure, capable of being applied to recovery of rockets of multiple specifications, wide in applicability and wide in application prospect.

The invention relates to the technical field of machine manufacturing, and provides a stainless steel rocket storage tank barrel section forming method, device, equipment and medium, the method comprises the steps that an X-ray diffraction test is conducted on a sample, and the martensite volume fraction and dislocation density are obtained based on an obtained diffraction curve; constructing a first function based on the martensite volume fraction, the corresponding environmental parameters and the deformation amount of the deformed sample; obtaining yield strength according to the dislocation density, and constructing a quantitative model according to the yield strength and a first function; and actual operation parameters are obtained and input into the quantitative model, corresponding optimal operation parameters are obtained, and the rocket tank cylinder section is manufactured. According to the method, the base metal and the welding seam are deformed at the same time by adopting the obtained optimal operation parameters such as low-temperature parameters to conduct a low-temperature deformation method, so that the base metal and the welding seam are both subjected to deformation strengthening, the joint strength coefficient of the cylinder section welding seam is improved, the integrity and reliability of the component structure are improved, resource waste is reduced, and the production efficiency is remarkably improved; and the comprehensive performance of the rocket tank is improved.

The invention discloses a grid rudder folding and unfolding mechanism, a rocket interstage section and a rocket, and relates to the technical field of grid rudders.The grid rudder folding and unfolding mechanism comprises a rudder shaft, a grid rudder, a driving cylinder, a piston, a piston rod and a connecting assembly. According to the technical scheme, the piston rod is driven to move through sliding of the piston in the driving cylinder, and then unfolding and folding actions of the grid rudder are achieved through the connecting assembly. Meanwhile, the piston rod is matched with a key groove in the folding state of the driving cylinder, the integrated design of folding and locking can be achieved, and unfolding and locking are achieved through air pressure after unfolding. Through the arrangement mode, mechanism integration is achieved, two sets of locking mechanisms are reduced, meanwhile, use of traditional servo mechanisms, torsional springs and other elastic elements is reduced, cost is reduced, reliability is improved, meanwhile, enough driving force can be provided, and it is ensured that the grid rudder can be unfolded and folded rapidly and reliably.

The invention discloses an omni-directional transfer device for vertical recovery of a rocket body, and relates to the technical field of rocket recovery, the omni-directional transfer device for vertical recovery of the rocket body comprises a plurality of automatic positioning and locking mechanisms, and the automatic positioning and locking mechanisms are arranged on the circumferential outer side of the rocket body; the automatic positioning and locking mechanism comprises a vehicle body, wherein a sliding block is slidably arranged on the top surface of the vehicle body; the vertical electric cylinder is rotatably arranged on the top surface of the vehicle body, and the upper end part of the vertical electric cylinder is connected with the locking hook through a first piston rod; and the first end of the overturning electric cylinder is rotatably connected with the sliding block, and the second end of the overturning electric cylinder is rotatably connected with the locking hook through a second piston rod. The omni-directional transferring device for vertical recovery of the rocket body solves the technical problems that an existing rocket body locking process is tedious in operation, needs manual intervention, has high requirements for operation space, is high in potential safety hazard, is long in consumed time, cannot be flexibly adjusted according to the diameter of the rocket body, and is small in application range.

The invention discloses a rocket sublevel active rotor pneumatic deceleration recovery system and a control method, and aims to solve the problems of low recovery precision, large fuel consumption, large landing impact and the like. The recovery process is divided into five stages of reentry, rotor wing unfolding, main pneumatic deceleration, accurate approaching and soft landing. Fine adjustment of the descending speed and the horizontal position is achieved through cooperative control of aerodynamic resistance of the body and the rotor wing system. And a multi-source sensor and a trajectory tracking algorithm are fused, so that stable soft landing at the sub-level zero horizontal speed and the minimum vertical speed is finally ensured. According to the invention, the recovery precision is obviously improved, the fuel consumption is reduced and the landing impact is reduced.

The invention provides a power measurement and control system and method for a liquid carrier rocket, and relates to the field. Comprising a to-be-tested real equipment system used for completing a rocket body test; the simulator is used for installing a simulation equipment system, and simulation equipment in the simulation equipment system is in one-to-one correspondence with real equipment in the real equipment system; the control machine is in communication connection with the real equipment system and / or the simulator and is used for issuing a target instruction to the real equipment system and / or the simulator in the rocket body testing process; the front-end machine is in communication connection with the control machine and the simulator and is used for forwarding first feedback information after the simulator executes the target instruction to generate the first feedback information; the monitoring machine is in communication connection with the front-end machine, the real equipment system and the simulator and is used for receiving first feedback information and / or second feedback information in the rocket body testing process and monitoring the first feedback information and / or the second feedback information; the second feedback information is feedback information generated after the real equipment system executes the target instruction.

The invention relates to a carrier rocket secondary reusable electrical system based on a modular TTE-TSN bus, and aims to solve the problems of reliability, light weight and maintenance efficiency in high-speed reentry and precise recovery of a secondary rocket. The system comprises a modular comprehensive control unit, a redundant TTE-TSN high-speed bus, a distributed thermal protection monitoring network and an intelligent health assessment module. Through the modular design, electrical equipment is greatly reduced, and the mass is obviously reduced; the redundant TTE-TSN high-speed bus realizes high-bandwidth and high-reliability transmission of messages on the rocket; the thermal protection monitoring network and the attitude coordination system dynamically optimize reentry thermal load distribution; the intelligent health assessment model predicts the structural life and reuse feasibility based on machine learning. According to the method, the reuse economy and reliability of the two-stage rocket are remarkably improved, and the method is suitable for commercial spaceflight and hypersonic flight vehicle tasks.

The invention provides a carrier rocket time sequence control full redundancy implementation method, and aims to improve the fault adaptability and the working reliability of a time sequence control system. The full-redundancy design method of three-path information source redundancy, three-path transmission line redundancy, two-out-of-three voting control redundancy and five-tube relay instruction driving redundancy is adopted in the sequential control system, the capacity of adapting to all first-degree faults and most second-degree faults is achieved, a single-point failure link does not exist, and the reliability is high. Normal work of the sequential control system can be ensured in the flight process of the carrier rocket. Based on a three-bus communication electrical architecture of a new generation carrier rocket electrical system, a full-redundancy design implementation method of information source redundancy, transmission line redundancy, voting control redundancy and instruction driving redundancy is adopted, high-reliability carrier rocket time sequence control is completed, the fault adaptive capacity of the time sequence control system is improved, and the time sequence control system has a wide application prospect. And the working reliability of the system in the rocket flight process is ensured.

The application provides an orbit transfer strategy optimization method and system for a large-ellipse frozen orbit satellite orbiting, comprising the following steps: acquiring a satellite-rocket separation orbit and a station parameter, calculating an elevation angle of the satellite and the station, and then determining a range of a true anomaly angle of pulse ignition meeting a measurement and control requirement; in combination with the range of the true anomaly angle of pulse ignition, generating a fuel most-probable pulse orbit change strategy of the satellite from an initial orbit to a target orbit, determining an orbit change speed increment, an ignition position and an ignition frequency; determining ignition-on and ignition-off conditions of each orbit change and a drift circle number between adjacent two orbit changes; according to the pulse orbit change strategy, respectively determining an initial value of an ignition time length, an ignition time and a thrust direction of each orbit change, and determining the final ignition time length, the ignition time and the thrust direction of each orbit change. The application adopts a non-apogee double-pulse orbit change strategy to design a fuel optimal orbit transfer strategy of a large-ellipse frozen orbit satellite orbiting, and solves the design method problem of the orbit transfer strategy.

Flexible hole forming equipment suitable for a rocket tube section end frame comprises an axial hole forming device, a radial hole forming device, a rotary table, a flexible tool and a control system. Wherein the rotary table is fixed on a foundation through an adjusting sizing block and a foundation bolt; the flexible tool is mounted on the rotary table through a positioning pin and a screw; the axial hole forming device and the radial hole forming device are installed on the two sides of the rotary table at the interval of 90-degree azimuth angles. The device is mainly used for machining axial butt joint holes and radial riveting holes of multi-model cylinder section end frames of spaceflight carrier rockets and axial milling, flexible hole forming of the rocket cylinder section end frames with the diameter of phi 3350 mm and phi 3800 mm can be met, and machining of all inner L-shaped and outer L-shaped end frames with the diameter of phi 3350 mm and phi 3800 mm is considered; the problem that a multi-drill jig tool corresponding to multi-type end frames is tedious to replace is solved, the size precision and the position precision of end frame drilling are greatly improved, and the machining period is shortened.