1826 results about "Rocket engine" patented technology

The invention relates to the technical field of rocket engine thrust chamber additive manufacturing, and discloses a metal powder recovery system for rocket engine thrust chamber additive manufacturing. Comprising a powder collection module, a screening classification module, a purification treatment module, a particle size adjustment module and an intelligent control module. The powder collecting module captures residual metal powder and collects data to generate a data set; the screening and grading module performs particle size separation by using a joint algorithm; the purification treatment module is used for constructing a multi-stage purification model to remove an oxide layer and decompose organic matters; the particle size adjustment module designs a composite regulation strategy to adjust particle size distribution; and the intelligent control module establishes an optimization model to generate a parameter adjustment instruction. In addition, the system is further provided with a powder characteristic analysis module, a circulation monitoring module and an exception handling module. The system can efficiently recover the metal powder, improve the quality and reutilization value of the metal powder, reduce the cost and guarantee the quality of additive manufacturing products, and is suitable for metal powder recovery treatment in the field of additive manufacturing of rocket engine thrust chambers.

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

The invention discloses a fatigue crack propagation analysis method of a rocket engine turbine blade. The method comprises the following steps: establishing a fatigue crack propagation model of the rocket engine turbine blade; determining high-low cycle fatigue load and input through the working condition load; the fatigue crack propagation calculation of the turbine blade is carried out in combination with ABAQUS and FRANC3D; fatigue fracture parameters such as stress intensity factors and J integral are output, and the critical crack size of the turbine blade is determined; and according to a fatigue crack propagation model, performing high-low cycle fatigue life prediction on the rocket engine turbine blade. Compared with existing fatigue crack propagation analysis, the method has the advantages that working condition loads in the whole cycle process of starting, stopping, stable operation and overload of the rocket engine turbine blade are considered, and calculation of the high-low cycle composite fatigue problem of the rocket engine turbine blade is achieved; and the fatigue crack propagation life and the critical crack length of the rocket engine turbine blade under the real working condition can be more accurately predicted.

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

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 provides a spacecraft quartz powder heat insulation tile which is prepared by adopting high-purity quartz powder (SiO2) as a main raw material and combining a high-temperature-resistant ceramic fiber reinforcement and a nano aerogel heat insulation layer through a high-temperature sintering process. The heat insulation tile has the following advantages: 1, the heat insulation tile has excellent high temperature resistance, and can tolerate 1600 DEG C high temperature for a long time and tolerate 1800 DEG C instantaneous thermal shock for a short time; 2, repeated use: the structure is stable, and high strength is still kept after multiple thermal cycles; light weight: the density is low (less than or equal to 0.5 g / cm < 3 >), and the load of a spacecraft is reduced; and 4, the thermal conductivity is low (less than or equal to 0.05 W / m.K), and heat transfer is effectively isolated. The method is suitable for thermal protection systems (TPS) of spacecrafts, such as thermal insulation of reentry capsules and rocket engine nozzles.

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

The invention discloses a rocket-boosted warplane ultra-short-distance launching take-off system and method. The system comprises a chassis, a high-precision track, a hydraulic lifting mechanism and a rocket sled. The rear end of the high-precision track is connected with the top of the chassis in a vertically swinging mode, the front end of the high-precision track is driven by a hydraulic lifting mechanism to ascend and descend, and the rear portion of the high-precision track is provided with a check block. The bottom of the rocket sled is provided with a sliding block used for being in sliding fit with the high-precision track. The mounting interface is used for being connected with a warplane undercarriage through an electronic lock, and the electronic lock is used for timing when a rocket engine is started, unlocking when the rocket engine reaches the working time and sending a separation success signal after being separated from a warplane. Small-overload ultra-short-distance take-off of the fighter is achieved, the volume of ground take-off facilities of the fighter is effectively reduced, the requirement for ground conditions is lowered, the flexibility and convenience of deployment are effectively improved, and the structure or a power system of the fighter is not greatly changed.

The invention discloses a pilot-operated type self-driving fluid release valve, and belongs to the field of control valves. Comprising a main valve assembly, a pilot valve assembly and a one-way valve assembly. Valve bodies of the pilot valve assembly and the one-way valve assembly are integrated on a main valve body of the main valve assembly; the main valve assembly comprises a main valve body, a main valve piston, a main valve element and a main valve spring. The pilot valve assembly comprises a pilot valve air control opening, a pilot valve inlet, a pilot valve outlet channel and a pilot valve pressure relief opening. The one-way valve assembly is arranged at an outlet of the main valve and used for blocking medium backflow. Through mechanical configuration reconstruction and system integration innovation, the technical contradiction of the liquid rocket engine high-pressure gas control valve in weight, response speed and reliability is thoroughly solved.

A fuel grain assembly for a hybrid rocket engine includes a hybrid fuel grain including fuel grain material, wherein an outermost portion of the fuel grain material defines an outer surface of the fuel grain, and wherein the fuel grain material includes a polymer based rocket fuel material; and a solid propellant material disposed in contact with the fuel grain material of the hybrid fuel grain, wherein the solid propellant material includes an oxidizer and a binder material, wherein a hollow combustion port extends through the fuel grain assembly, wherein at least a portion of a wall of the hollow combustion port is defined by the solid propellant material.

The invention discloses a water-air bimodal self-adaptive rocket engine system and a control method, and belongs to the technical field of rocket control. The system comprises an environment sensing recognition module, a thrust output regulation and control module, a fuel combination switching module, a heat exchange regulation module, an attitude propulsion control module and a modal coordination control module, and precise recognition and rapid classification of a rocket operation environment are realized through combined judgment of a multi-point distributed sensing unit, density and flow characteristics; by constructing a multi-level thrust adjusting system, the dynamic change of the geometrical shape of the spray pipe is achieved, combustion supply ratio adjustment and multi-parameter linkage control are fused, the self-adaptability and response sensitivity of thrust output are remarkably improved, rapid conversion between high-pressure short-range propulsion and efficient long-range flight is achieved, and the high-pressure short-range propulsion and high-efficiency long-range flight are achieved. Regulation and control commands are preloaded and distributed in parallel according to the current environment, so that all functional modules can respond cooperatively, the operation mode in the new environment can be quickly adapted, and the system can capture the execution state of the regulation and control commands in real time.

The invention relates to a multifunctional pump rear swing liquid rocket engine single-machine rack which is characterized in that two main force bearing beams and two auxiliary force bearing beams are welded to form a lower force transmission frame; the four upper connectors and the four upper connecting beams are welded to form an upper force transmission frame. The four main force bearing rods are welded to the upper force transmission frame and the lower force transmission frame respectively. The two servo supports are arranged to serve as upper fulcrums of the servo mechanism, and the bidirectional swing requirement of the liquid rocket engine can be met. And a pair of turbine pump supporting seats are symmetrically arranged with one of the servo supports. And a pair of turbine pump auxiliary support lugs is arranged on the same side of the turbine pump support. The turbine pump support and the turbine pump auxiliary supporting lug are jointly responsible for being connected with the turbine pump, and stable fixed supporting of the turbine pump is achieved. The main force bearing beam is connected with the lower end force transmission assembly, the upper connector is connected with a rocket body or a multi-machine rack, and the thrust transmission function of the engine is achieved.

The invention discloses a solid rocket engine composite material skirt connecting structure and a forming method thereof.The connecting structure comprises a composite material skirt, a skirt hole reinforcing ring, a connected cabin section and a connecting piece, the composite material skirt and the skirt hole reinforcing ring are integrally formed, and the rear end of the composite material skirt and the rear end of the skirt hole reinforcing ring are connected with the connected cabin section through the connecting piece; the connected cabin section is of an L-shaped nested inner lap joint structure and has the butt joint limiting and bearing functions, and the thickness transition section at the front end of the composite skirt is connected with a fiber winding structure layer of a combustion chamber shell. Compared with a traditional metal skirt connecting structure, the weight reduction effect is obvious, compared with a traditional composite skirt hinge block connecting structure, the connecting mode is simpler, and the fiber continuity of a composite skirt connecting area is guaranteed; the problems that in-plane shear rigidity and strength of the composite material are insufficient, and orifice positions are prone to damage are solved. The structure and the forming process are simple, the maturity is high, and the cost is low.

The invention provides an online fault detection method for a reusable liquid rocket engine, and aims to realize real-time online adjustment of a fault threshold according to the change of the working condition of the engine under the condition of large-range variable thrust of the reusable engine so as to form a red line shutdown strategy based on an online dynamic threshold. Faults occurring in the working process of the engine are pre-warned and judged, so that effective measures are taken in time, and the safety of the carrier rocket and the load is protected.

The utility model discloses a docking device based on a hoisting type small envelope rocket multi-engine single-machine parallel structure. The docking device is characterized by comprising a crane (1), an inclined steel cable (2), a cross beam (3), a length adjusting device (4), a profiling hoop (6) and a combined hoop (7), the rocket engine (5) is composed of a spray pipe (8), an engine body (9) and a rack (10). According to the butt joint device, a combined type holding ring and profiling hoop structure is designed by means of only two force bearing points on the engine, the structure is simplified, the envelope of the butt joint device of the multi-engine single-machine parallel connection structure is smaller than that of a single engine body, single engine butt joint of the multi-engine single-machine parallel connection structure in a narrow space under complex working conditions is achieved, and the butt joint efficiency is improved. The problem of interference in the butt joint process of all the engines in the multi-engine single-machine parallel structure is solved, meanwhile, safety risks are eliminated, and the butt joint reliability and working efficiency of the multi-engine single-machine parallel structure in a narrow space under complex working conditions are improved.

The invention discloses a liquid rocket engine airtight clamp and a using method thereof. The clamp comprises a handle, a large nut, a jacking hand wheel, a supporting disc, a long pull rod, a short pull rod, a spring, an air inlet pipe, a sealing disc and a turning plate. When the airtight clamp enters the engine, the turning plate is in an inclined state, after the turning plate completely enters the throat part of the engine, the handle is pulled to enable the turning plate to be in a state perpendicular to the axis of the engine, and the turning plate is clamped above the throat part of the engine; then the large nut is rotated, the sealing disc moves towards the throat of the engine until the large nut cannot rotate, at the moment, the turning plate and the throat sealing ring on the sealing disc clamp the engine, meanwhile, the supporting disc is attached to the large end of the engine under the action of the spring, and the airtight clamp is fixed to the engine. An external air source is connected with the air inlet pipe to inflate the inner cavity of the engine. The airtight test clamp can be widely applied to the field of sealing of small rocket engines, and has the advantages of being novel in structure, reliable in sealing and easy to operate.

The invention discloses a liquid rocket engine fault simulation system, method, medium and equipment, the system comprises a fault simulation module, an environment interference simulation module and an image acquisition and fault identification module, the fault simulation module is used for simulating various fault states of a liquid rocket engine in a test run process; the environment infection simulation module is used for simulating various environment interferences occurring in the test run process of the liquid rocket engine; and the image acquisition and fault identification module is used for acquiring video image data of the liquid rocket engine in the fault simulation and environment interference simulation process so as to carry out real-time identification and diagnosis on the fault of the liquid rocket engine. According to the invention, the fault of the liquid rocket engine in the test run process can be accurately identified.

The invention discloses a rocket engine performance prediction method based on physical parameter identification and residual learning. The method comprises the following steps: constructing a data set based on historical test or simulation data; determining key physical parameters influencing engine performance based on the engine performance simulation model, and constructing an agent model for describing a mapping relationship between the physical parameters and performance parameters; constructing a hybrid deep neural network, taking the sensor time sequence as input, and outputting a prediction value of a key physical parameter and a residual sequence used for correcting an error between a reference prediction result and a real performance parameter; and inputting the predicted key physical parameters into the agent model to obtain a reference prediction result, and superposing the reference prediction result with the residual sequence to obtain a final engine performance prediction result. Training a hybrid deep neural network to minimize an error between a prediction result and a real label; and deploying the trained hybrid deep neural network model to an online system to realize real-time prediction of rocket engine performance parameters.

A cooling device for rocket engine nozzle comprising a first member including a cylindrical or tapered inner circumstance shape, channels provided over the part or all of the inner circumstance surface, and a helical groove part on each channel, and a second member including an outer circumstance surface of being the substantially same shape as the inner circumferential surface of the first member, and a nozzle-shaped cavity part therein to form a nozzle part, and then by assembling both of the members, forming tube parts corresponding to the shape of the helical groove parts as helical cooling channels placed on positions spaced predetermined distances from a throat part of the nozzle part, by which preventing coolant running in the helical cooling channels from excessively heated and helping the coolant perform its designed cooling ability to suppress nozzle erosion, upon flue gas of the rocket engine passing through the throat part.

The invention discloses a solid rocket engine bearing structure lightweight determination method based on a carbon fiber composite material, and relates to the technical field of composite material structure design, and the method comprises the steps: obtaining the structure size and geometric shape of a fixed body; replacing a titanium alloy material with a carbon fiber composite material, and determining a carbon fiber composite material structure response model based on a finite element modeling method; determining an optimal layering scheme based on structural response sensitivity analysis and a multi-objective optimization method; then determining a thickness optimization scheme based on the structural size and geometric shape of the fixed body and load distribution; determining a reinforcement optimization scheme based on a parametric modeling method; determining a metal reinforcement configuration by adopting a structural rigidity-quality collaborative evaluation method; and finally, determining a lightweight structure of the fixed body based on a composite material layering process design method. The mass of the fixing body structure can be reduced, the mechanical property of the fixing body structure is improved, and the high-performance design requirement of a solid rocket engine is met.

The invention discloses a liquid engine ignition device. The liquid engine ignition device comprises a storage tank used for storing an ignition agent required by preset ignition times and an ignition tube used for storing the ignition agent required by single ignition; the ignition tube is communicated with the storage tank through a communicating tube; a one-way valve for preventing an ignition agent in the ignition pipe from flowing back into the storage tank is arranged on the communicating pipe; the ignition tube is configured to be capable of extruding an ignition agent which is stored in the ignition tube and is used for single ignition into an engine combustion chamber under the extrusion effect; the storage tank is configured to be capable of automatically filling the ignition tube with the discharged ignition agent with the ignition agent stored in the storage tank under the extrusion effect. According to the method for sharing the storage tank and the ignition tube, the reliability of the ignition tube and the multi-time ignition characteristic of the storage tank are combined, automatic quantitative ignition agent filling is achieved, and an efficient ignition scheme is provided for multi-time ignition of a rocket engine.

The invention discloses a method for detecting position information of a butt joint hole of a skirt end face of a solid rocket engine, and the method comprises the steps: 1, installing a tool on the butt joint hole of the skirt end face, and completing the laser scanning of the skirt end face and a cylinder segment which is about 200mm around the skirt; the tool is conical, the tail end of the tool is provided with threads, the tool can be directly installed on a skirt end face threaded hole, and the axis of the conical face is coaxial with the axis of the threaded hole. 2, performing preliminary optimization on the three-dimensional point cloud model obtained by scanning, then eliminating errors caused by noisy points by using a filtering model combining a system filtering algorithm and a moving least square method, and finally completing overall optimization of the point cloud data; 3, packaging the optimized model by using a triangle principle, carrying out reverse modeling on the packaged three-dimensional model, and extracting to-be-detected data by using the reverse regularized model and the three-dimensional point cloud model; according to the invention, detection of position information of butt joint holes of skirt end surfaces of various models of solid rocket engines including large diameters can be realized.

The present application discloses a control method and system for a booster system of a launch vehicle engine, wherein the control method includes the following steps: collecting pressure data; making a rational judgment on the collected pressure data; if the rational judgment is passed, making a validity judgment on the pressure data; making a comprehensive calculation on the pressure data according to the validity judgment result and the rational judgment result; judging whether the result after the comprehensive calculation is valid; if the result after the comprehensive calculation is valid, performing a closed boost control; if the result after the comprehensive calculation is invalid, performing a boost valve opening and closing control according to a preset valve action sequence. The present application also makes a reasonable threshold judgment and data fusion on the data of redundant sensors, and when one sensor fails, it can ensure the correctness of the boost calculation and the correct opening and closing of the boost valve. Under extreme conditions, the boost system can be guaranteed to work within the normal range by fixing the boost timing. Thereby improving the reliability of the boost control system of the launch vehicle.

The invention relates to a vacuum chamber for a high-altitude simulation test of a liquid rocket engine, belongs to the technical field of aerospace liquid rocket engine tests, and solves the technical problems of single function, difficulty in operation and poor safety of a traditional vacuum chamber. The cabin comprises a cabin body, a cabin door, a heat sink system, a fire fighting system, a gas replacement system, an anti-arch blasting device, a blowdown system, a mounting platform, an operation platform, a propellant supply pipeline and measurement and control equipment. The cabin body is cubic, and the working pressure in the cabin body is 0-0.1 MPa. The cabin door is used for entering and exiting of personnel, small ground equipment, an engine and a thrust measuring device. The heat sink system, the fire fighting system, the gas replacement system, the sewage discharge system, the mounting platform and the propellant supply pipeline are mounted on the cabin body, and the measurement and control equipment is a transmission channel for engine control and measurement signals. The vacuum chamber is used for a liquid rocket engine high-altitude simulation test.

The invention discloses a rocket-based rotary detonation engine combustion chamber structure and application thereof, and belongs to the technical field of rocket engine research. The combustion chamber structure comprises an isolation section, an equal-direction oil injection section and an expansion-concave cavity section which are connected in sequence. A central rocket is arranged in the isolation section and the equal direct oil injection section; and four groups of first ring-jet injection points are arranged on the equal straight oil injection section, two groups of second ring-jet injection points are arranged on the central rocket, and two groups of third ring-jet injection points are arranged on the expansion-concave cavity section. Oil is injected into the expansion-concave cavity section firstly, a pre-detonation pipe is adopted for igniting an oil-gas mixture, rotating detonation waves are formed, the built-in center rocket can provide thrust for the aircraft at the ground zero speed and form an injection effect for air suction, extra thrust is provided when the cruising thrust of the aircraft is insufficient, and when the rotating detonation combustion flameout boundary exists, the detonation effect is achieved. A flame stabilizing mechanism is provided for rotary detonation combustion, and stable work of wide-speed-range and wide-height-range rotary detonation combustion is achieved.

The invention provides a universal assembling and testing method for a bipropellant attitude and orbit control liquid rocket engine. The assembling and testing method provided by the invention comprises 27 process steps, the method specifically comprises the steps of part collection, laser marking, cleaning, assembly assembling, strength testing, engine outer leakage rate detection, electrical performance detection, control valve inner leakage rate detection, mechanical testing, liquid flow testing, pore plate debugging, alcohol replacement, nitrogen purging, vacuum drying, purging, dew point testing, metering, assembling before warehousing, multimedia recording, engine final inspection and warehousing. Wherein the leakage rate detection and the electrical performance test are carried out for multiple times according to specific installation and test requirements. The attitude and orbit control engine assembling and testing method has the advantages that the assembling and testing effect of the attitude and orbit control engine is tested from multiple dimensions, generalization and standardization are achieved, and reliable work of the attitude and orbit control engine can be guaranteed.

The invention provides an electric pump pressure type liquid rocket engine and a carrier rocket. The electric pump pressure type liquid rocket engine comprises a gas-driven generator system, an electric pump pressure type liquid rocket engine body and an energy management system, and the gas-driven generator system is electrically connected with the energy management system; and the electric pump pressure type liquid rocket engine is electrically connected with the energy management system. According to the technical scheme, the rocket is prevented from carrying a large-mass battery through the generator system, and the highest-efficiency operation of the generator under all working conditions is realized through the energy management system. According to the technical scheme, the energy loss is small, the structural mass is light, the development and application cost of the carrier rocket is reduced, and the reliability of the carrier rocket is improved.

Omnivorous solar thermal thrusters and adjustable cooling structures are disclosed. In one aspect, a solar thermal rocket engine includes a solar thermal thruster configured to receive solar energy and one or more propellants, and heat the one or more propellants using the solar energy to generate thrust. The solar thermal thruster is further configured to use a plurality of different propellant types, either singly or in combination simultaneously. The solar thermal thruster is further configured to use the one or more propellants in both liquid and gaseous states. Related structures can include valves and variable-geometry cooling channels in thermal contact with a thruster wall.

The invention relates to the technical field of rocket engine simulation tests, in particular to a vertical high-altitude simulation test bed compatible with a ground test and a test method. Comprising a vacuum chamber which is provided with a ventilation assembly capable of being opened and closed; and the diffuser is communicated with the vacuum cabin and is vertically arranged at the tail end of the engine, and engine gas enters the diffuser. When a high-altitude simulation test is carried out, the long-nozzle engine is arranged in the vacuum cabin, and the openable and closable ventilation assembly is closed, so that the inner side of the vacuum cabin is isolated from the outside. Engine fuel gas enters the diffuser and is expanded in the diffuser to form supersonic airflow to suck the vacuum chamber, the fuel gas is injected to the injection expansion section under injection of the injection working medium, the fuel gas is mixed with the injection working medium airflow to continue to be pressurized in the supersonic speed state, and finally the fuel gas is exhausted from the injection expansion section when the pressure is higher than the atmospheric pressure. The air pressure in the vacuum chamber is lower than the atmospheric pressure, and high-altitude simulation test conditions are met.

The invention discloses a core mold for forming a complex heat insulation coating sleeve and a forming method, and belongs to the technical field of solid rocket engines. The core mold comprises an end socket core mold, a barrel section core mold assembly and a connecting buckle. The barrel section core mold assembly is of a split type, a plurality of fan-shaped plate-shaped barrel section core molds are sequentially connected along the same circumferential face to form a columnar barrel structure, and every two adjacent barrel section core molds are detachably and fixedly connected through a connecting buckle. The end socket core mold is buckled to one end of the barrel section core mold assembly and detachably and fixedly connected with the barrel section core mold assembly through a connecting buckle. According to the core mold with the structure, the butt joint interface structures of all the components are unified, detachable assembly can be achieved only through a plurality of connecting buckles of the same specification, disassembly and assembly are convenient, and the mold preparation time can be effectively shortened. The coating sleeve forming method adopting the core mold is good in manufacturability and high in product quality consistency, and due to the fact that the outer surface structure of the core mold is smooth, vacuumizing tightening can be conducted through a vacuum bag, vulcanization forming is conducted in a vulcanization tank, and the production efficiency is improved.