1441 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.

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 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 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 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.

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 relates to the technical field of rocket engines, in particular to a ground engine tail flame infrared radiation end-to-end prediction method based on a neural network, and the method comprises the following steps: 1, calculating a rocket engine tail flame infrared radiation data set; 2, constructing an asymmetric convolution encoder and a parallel deconvolution decoding module; 3, taking the profile spectrum parameter data set as an input data set; processing the tail flame infrared radiation intensity distribution cloud picture data, and taking the processed tail flame infrared radiation intensity distribution cloud picture data and tail flame infrared radiation integral intensity data as an output data set; 4, dividing the input data set and the output data set into a training set and a test set; training and testing the neural network; and 5, inputting the profile spectrum parameters into the neural network, and calculating the infrared radiation data of the tail flame of the rocket engine to be detected. The method solves the problems that an existing rocket engine tail flame infrared radiation calculation method is multiple in calculation links, long in calculation time consumption and difficult to meet mass calculation requirements, and is suitable for rocket engines.

The invention discloses a rocket engine health management system and method for full-life data value mining, and the system comprises a big data storage and calculation center which receives and stores the information from all modules of the system, and achieves the data storage; the test run data management module is used for creating model, station number, measuring point and working condition labels, and importing, labeling and classifying the test run data; the fault diagnosis strategy module provides functions of strategy construction, training, testing, evaluation and publishing; a health assessment and prediction strategy is formed, and the health assessment and prediction strategy is sent to the intelligent analysis module; the intelligent analysis module is used for developing, managing and evaluating a fault diagnosis strategy and integrating a diagnosis result into the fault knowledge base module; and the fault knowledge base module is used for establishing a fault tree and associated fault modes and case libraries, realizing accumulation and application of fault knowledge and storing related information in a big data storage computing center.

The invention provides a carrier rocket engine stage-switching motor redundancy control system, which is combined with a multi-input instruction redundancy voting method, a voting path software and hardware arbitration method and the like to complete high-reliability carrier rocket engine stage-switching motor control. According to the system, two-out-of-three voting is carried out by using a triple-redundancy input control instruction and interactive heartbeat, software and hardware joint arbitration is carried out on voting results of three modules, and an instruction control channel is determined to complete phase sequence control of the stage conversion winding. The method has strong fault capability, has no single-point failure link, and can ensure the normal operation of the engine conversion motor in the flight process of the carrier rocket.

The present application relates to a kind of rocket engine injector thin-wall milling self-positioning fixture and assembly method, it mainly includes upper positioning tire body component, conical surface expansion ring, tire unloading bottom plate, pressing cover plate, tire unloading lifting ring, pressing bolt, expansion screw, lower positioning bottom plate component, locking bolt, long fixed bolt, lower positioning bottom plate component is fixed in machine tool workbench by locking bolt long-term, always not disassembled.Product is positioned and clamped and supported by conical surface expansion ring, tire unloading bottom plate, pressing cover plate, pressing bolt, expansion screw before milling injector thin-wall and positioning tire body component, limit the radial and axial movement of injector thin-wall;Injector thin-wall is positioned and clamped and supported by conical surface expansion ring, tire unloading bottom plate, pressing cover plate, pressing bolt, expansion screw before milling injector thin-wall and positioning tire body component, limit the radial and axial movement of injector thin-wall;Injector thin-wall is positioned and clamped and supported by conical surface expansion ring, tire unloading bottom plate, pressing cover plate, pressing bolt, expansion sleeve before milling injector thin-wall and positioning tire body component, limit the radial and axial movement of injector thinn

The invention discloses a liquid oxygen methane rocket engine fault prediction method and system based on an ARMA model, and the method comprises the steps: constructing a modularized simulation model of a liquid oxygen methane rocket engine, injecting fault factors to simulate typical faults (such as methane pump cavitation, turbine blade ablation and the like), and obtaining normal and fault data containing Gaussian noise; carrying out time sequence modeling on the six key parameters by adopting an ARMA (Autoregressive Moving Average Model), and optimizing model parameters through ACF / PACF order determination and AIC criteria; establishing a self-adaptive residual threshold formula (threshold = residual mean + / -2 times of standard deviation), and realizing early warning of the fault by combining a judgment criterion that the threshold is continuously exceeded for three times and three parameters cooperatively give an alarm; experiments show that the alarm time is 0.1-0.13 seconds earlier than that of a traditional red line threshold method, the false alarm rate is reduced from 18.7% to 0.3%, and the spaceflight-level real-time performance and reliability requirements are met; the method does not need a large number of fault samples, is high in anti-interference capability, and is suitable for online health monitoring of reusable rocket engines.

The invention provides a rocket engine gimbal seat which comprises a pair of upper supports used for being connected with a rocket engine rack, a gimbal ring and a pair of lower supports, the bottoms of the pair of upper supports are hinged to the first two opposite side walls of the gimbal ring in a one-to-one correspondence mode, and the bottoms of the pair of upper supports are hinged to the second two opposite side walls of the gimbal ring in a one-to-one correspondence mode. The tops of the pair of lower supports are hinged to the second two opposite side walls of the gimbal ring in a one-to-one correspondence mode, and the middles of the pair of upper supports and the middles of the pair of lower supports make contact with the gimbal ring through cylindrical surfaces.

The invention discloses a double-layer artificial debonding heat insulation structure of a solid engine and a forming method, and relates to the field of solid rocket engines. A heat insulation layer, a bottom layer coating layer and a cover layer coating layer are adopted on the inner surface of a combustion chamber shell to form a double-layer manual debonding heat insulation structure, the bottom layer coating layer and the heat insulation layer form an open type manual debonding structure, and the cover layer coating layer and the bottom layer coating layer form a closed type manual debonding structure. The heat insulation layer, the bottom layer coating layer and the cover layer coating layer are sequentially paved with green sheets and are formed through one-time vulcanization. The enough bonding area of the charged powder can be guaranteed, and the stress of the grain can be fully released.

In a rocket motor in which burn rate enhancement (BRE) wires are used to accelerate the burn rate of the solid propellent, embedded charge assemblies (ECAs) are configured as support structures for the BRE wires. Each ECA includes an energetic material that is configured to burn along with the solid propellent to produce thrust and, upon detonation, to break up the solid propellent to terminate thrust. The detonation may also be initiated as a part of process to prevent a higher-order reaction, such as in reaction to heating from a fire or other cause. By being located inside the casing, the energetic material and ECAs do not adversely affect aerodynamics of the flight vehicle of which the rocket motor is a part, such as a missile.

The invention discloses a solid rocket engine in-situ heat preservation ignition test run frame which comprises a fixing base, a movable fixing plate, a supporting assembly, a force bearing pier, a first force transmission piece, a second force transmission piece, a force measuring piece, a heat preservation bottom plate and a heat preservation cover. A solid rocket engine to be tested is movably arranged on the movable fixing plate through the two supporting assemblies, and the movable fixing plate is connected to the fixing base through the multiple plate springs. The heat preservation bottom plate is arranged on the lower side of the movable fixing plate, the heat preservation cover is buckled on the heat preservation bottom plate, cold air is conveyed into the heat preservation cover through the cold air connector, the solid rocket engine to be tested can be cooled, and when the temperature of the solid rocket engine to be tested reaches the temperature required by testing conditions, the movable door is opened, and the solid rocket engine to be tested is ignited. And the force measuring piece can obtain thrust data of the solid rocket engine. The solid rocket engine in-situ heat preservation ignition test run rack can accurately simulate the thrust test of the solid rocket engine in an ultralow-temperature working environment.

The invention provides a thermal insulation structure of a rocket engine injector and a reliability assessment method thereof. The thermal insulation structure can solve the problems that the temperature of the upper surface of an injector body is too high and the use reliability is poor. The invention provides a thermal insulation structure of a rocket engine injector. The thermal insulation structure comprises an injector body, a thermal insulation structure body, a boss and an arc-shaped groove. The outlet section of the injector main body is improved into a multi-boss matched arc-shaped groove structure from a traditional solid structure, so that the thermal resistance conducted to the upstream of the injector by a high-temperature spray pipe is increased, the temperature of the upper surface of the injector main body can be reduced, the design difficulty and the production cost of upstream parts are further reduced, good mechanical resistance and high-temperature structural strength are ensured, and the service life of the injector is prolonged. The structural weight is reduced while the central flow channel is not affected, and the problems that the temperature of the upper surface of the injector body is too high and the use reliability is poor are solved.

The invention relates to the technical field of rocket engine simulation tests, in particular to a tool and a process for installing an ejection system of a vertical high-altitude simulation test bed. Comprising a cabin bottom platform which is arranged on the ground and located above the plane of a fuel gas diversion trench; the mounting hole is formed in the cabin bottom platform, and a vacuum cabin flange located in the mounting hole is arranged at the bottom end of the vacuum cabin; and the ejection system flange is arranged on the ejection system in a sleeving manner and is suitable for being connected with the vacuum cabin flange. In the application, the plane of the fuel gas diversion trench can be used for placing an injection system, and a vacuum cabin flange can be arranged in the mounting hole and can be connected with an injection system flange. The lifting part is used for lifting the injection system, and lifting force in the vertical direction can be converted into lifting force in the horizontal direction, so that the lifting part can penetrate through the vacuum cabin door to be arranged on the inner side of the vacuum cabin, the requirement for the height of the lifting part is lowered, and interference between the lifting part and the upper space of the vacuum cabin is avoided.

The invention discloses a multichannel thermocouple temperature compensation method for a liquid rocket engine test, and relates to the technical field of liquid rocket engine tests, and the method comprises the steps: determining the type of a target thermocouple temperature sensor according to the characteristics of a tested object in a to-be-tested liquid rocket engine, and determining a corresponding temperature compensation wire; a target thermocouple temperature sensor is connected with a temperature compensation wire to form an acquisition device; a thermocouple voltage signal and a thermistor voltage signal are acquired in real time through an acquisition device, and the signals are transmitted to a processor based on a wiring terminal in the temperature compensation cabinet; and the thermocouple temperature corresponding to the input data is determined through a preset cold end compensation algorithm in the processor. According to the invention, the precision and reliability of temperature measurement can be improved, the system structure is simplified, the system cost is reduced, and the system is suitable for temperature measurement requirements of multi-channel, high-precision and complex environments in a liquid rocket engine test.

The invention discloses a liquid-propellant rocket engine turbine pump inducer design method and an inducer, and the method comprises the steps: determining the working conditions of the inducer according to the overall design requirements of a liquid-propellant rocket engine; according to design requirements and an empirical formula, basic parameters of the inducer are preliminarily determined; different core parameters are combined according to the core parameter range of the inducer, and a plurality of initial three-dimensional models of the inducer are established; respectively carrying out analogue simulation on the plurality of established initial three-dimensional models of the inducer to obtain the cavitation performance of the inducer; according to a flow analysis result, obtaining an inducer of an optimal blade profile in all parameter combinations; and parameter optimization and adjustment are conducted on the inducer with the optimal blade profile till the inducer with the optimal cavitation performance for the turbine pump of the liquid rocket engine is obtained. By means of the method, the liquid rocket engine turbine pump inducer with the good cavitation performance can be designed.

The invention discloses a device and method for forming the inner wall of a copper alloy contraction and expansion section for a liquid rocket engine, and the device for forming the inner wall of the contraction and expansion section comprises the following steps: clamping a blank once, and forming the blank twice by using a hydraulic machine to obtain the inner wall of the contraction and expansion section with machining allowance; and carrying out heat treatment and machining on the inner wall of the convergent-divergent section with the machining allowance to obtain the inner wall of the convergent-divergent section. The required raw materials are light in weight, the material utilization rate is high, the production period is short, the machining efficiency is high, and the manufacturing cost of the inner wall of the copper alloy contraction and expansion section for the liquid rocket engine can be greatly reduced.

The invention discloses a liquid rocket engine digital test run platform, relates to the technical field of liquid rockets, and aims to solve the problem that a liquid rocket engine digital test run system in the prior art cannot perform multi-dimensional simulation. The liquid rocket engine digital test run platform comprises a one-dimensional simulation module, a three-dimensional simulation module, a one-dimensional and three-dimensional joint simulation module and a virtual test run display module. The one-dimensional simulation module is used for performing one-dimensional simulation on the liquid rocket engine under different working conditions; the three-dimensional simulation module is used for performing three-dimensional simulation on a target part in the liquid rocket engine under different working conditions; the target component at least comprises a thrust chamber and a flow regulator; the one-dimensional and three-dimensional joint simulation module is used for performing one-dimensional and three-dimensional joint simulation on the target component; and the virtual test run display module is used for visually displaying the test run process of the liquid rocket engine. Therefore, the multi-dimensional comprehensive simulation of the digital test run of the liquid rocket engine can be realized.