246 results about "Thrust chamber" 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 an integrated closed circulation system for an RBCC engine, and solves the problems that a constant-pressure extrusion mode of an existing RBCC engine propellant supply system is not suitable for variable working condition adjustment of an RBCC rocket thrust chamber, and a high-pressure storage tank is huge in structure, large in weight and not beneficial to application and expansion of the system on the day. According to the system, high-temperature oxygen is generated through catalytic decomposition of hydrogen peroxide and acts on the turbine, then the hydrogen peroxide pump and the coal oil pump are driven to work, integrated pressurization, supply and adjustment of propellants can be carried out on the multiple rocket thrust chambers and the ramjet combustion chamber at the same time, and therefore wide-range fuel oil adjustment is met, and a widely-applied electric pump system can be replaced; and the structural envelope and weight are greatly reduced, so that the power requirement of a longer voyage is met.

The invention relates to the technical field of rocket thrust chamber manufacturing, and discloses an additive forming part linear cutting and taking auxiliary system for a rocket thrust chamber. The system comprises a multi-source data acquisition module for acquiring geometrical morphology data of the additive forming part through various sensors; the cutting path planning module is used for generating a global cutting path sequence based on an improved genetic algorithm; the dynamic deformation compensation module is used for compensating thermal deformation and vibration interference in real time by using a fuzzy adaptive observer; and the multi-axis cooperative control module is used for constructing a distributed constraint optimization model to realize accurate control of multi-axis motion. The system can accurately collect data, reasonably plan a cutting path, effectively compensate deformation, ensure multi-axis cooperative work, remarkably improve the quality and efficiency of linear cutting and taking of the rocket thrust chamber additive forming part, reduce the production cost and provide powerful technical support for spaceflight manufacturing.

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.

The invention discloses a metal powder spreading device for additive manufacturing of a rocket engine thrust chamber, and relates to the technical field of additive manufacturing. Comprising a machine body, a mounting base is slidably arranged on the inner wall of a cavity of the machine body through a rail, and the bottom of the mounting base is connected with a powder spreading head through an air cylinder; the powder spreading device further comprises a powder feeding pipe, the powder feeding pipe is fixed to the top of the powder spreading head in a penetrating mode, the powder feeding pipe is connected with a powder feeding mechanism, the top area of the powder spreading cavity is connected with a guide net plate through a first electric push rod, and a sealing plate is fixed to the bottom of the guide net plate through a connecting rod and arranged at an opening in the bottom of the powder spreading cavity. A paving assisting strip is vertically arranged in the mounting sleeve in a sliding mode, and a pressing plate is fixed to the bottom of the paving assisting strip and located on the sealing plate. According to the metal powder spreading device for additive manufacturing of the rocket engine thrust chamber, the powder spreading uniformity is guaranteed in the longitudinal direction, meanwhile, powder caking is avoided, and the powder spreading efficiency is improved in cooperation with vibration.

The invention discloses a pressure-adjustable fuel gas pressurization system and an RBCC engine. The technical problem that an existing pressurization system cannot meet the variable working condition adjustment requirement of a thrust chamber of the RBCC engine is solved. The system comprises a pressurizing storage tank, a gunpowder starter, a generator, an oxidizing agent storage tank, a fuel storage tank and a controller, the inlet end of the pressurizing storage tank is connected with the gunpowder starter, and the outlet end is connected with the inlet end of the generator through the first electric explosion valve, the electric pump and the first one-way valve; a pressure sensor is installed at the outlet end of the generator and connected with the controller. The outlet end of the generator is connected with the inlet end of the pressurizing storage box and is connected with the inlet ends of the oxidant storage box and the fuel storage box; the outlet ends of the oxidant storage tank and the fuel storage tank are respectively connected with the thrust chamber; a pressure release valve is mounted between the outlet end of the electric pump and the outlet end of the oxidant valve or the fuel valve; the controller is connected with the gunpowder starter, the electric pump and all the valves and used for sending out corresponding control instructions.

The invention discloses a multi-scale bidirectional coupling sweating cooling simulation method which comprises the following steps: firstly, on a mesoscale, constructing a mesoscopic flow model, simulating flow of a coolant, and obtaining parameters such as porosity, permeability and inertia coefficient of a porous medium; introducing the obtained parameters into a porous medium panel area of a thrust chamber geometric model constructed under a macroscale to realize cross-scale parameter coupling; the flow and heat transfer process in the thrust chamber geometric model after cross-scale parameter coupling is simulated, and the speed, pressure, temperature and local heat flux density distribution of a combustion chamber and a porous medium panel area under the macroscale are obtained; and taking the local heat flow density as a new wall surface boundary condition, capturing the flow and heat transfer characteristics of a coolant in pores under a thermal state working condition based on the mesoscopic flow model, and further evaluating the overall cooling performance and the thermal protection effect of the porous medium panel. Through coupling calculation of the mesoscopic pores and the macroscopic combustion chamber, the accuracy and reliability of prediction of the cooling performance of the porous medium are improved.

The invention provides a diagnosis method and device for unstable combustion oscillation characteristics of a rocket engine, and relates to the field of rocket engine testing, and the method comprises the steps: obtaining part parameters of all parts of a thrust chamber of a to-be-analyzed rocket engine; based on the nozzle parameters and the channel parameters, nozzle outlet sound admittance is determined, and mass flow pulsation of each nozzle channel outlet is determined; engine combustion chamber design parameters and configured flame parameters are processed, and heat release pulsation of a single flame is determined; and based on the configured modal cutoff value, mass flow pulsation and heat release pulsation, solving the configured acoustic characteristic network to obtain a plurality of actual main frequencies and modal numbers corresponding to the actual main frequencies, and determining modal categories of the actual main frequencies. According to the dynamic response characteristics of the injection pulsation of a single nozzle or multiple nozzles and the heat release pulsation of the combustion flame, the actual main frequency of the low-temperature engine during oscillation combustion under different working conditions can be predicted, and relevant parameter analysis is carried out.

The invention discloses a low-temperature upper-stage power system, and relates to the field of spaceflight. Comprising a gas cylinder, a pressure supplementing electromagnetic valve, a pressure supplementing pore plate, a self-generation pressurizing electromagnetic valve, a self-generation pressurizing pore plate, a storage tank, an overflow valve, a storage tank pressure sensor, a pressure stabilizing valve, a circulating pump, a circulating control valve, a pump front valve, a Joule-Thomson throttling expansion valve, a heat exchanger and a branch supply control valve; the main engine subsystem comprises a pump, a turbine, an auxiliary valve, a fuel gas generator, a main valve, a precooling backflow valve, a main thrust chamber, an evaporator, a starting gas cylinder, a starting gas cylinder control valve, a main thrust chamber igniter and a fuel gas generator igniter. The attitude control engine subsystem comprises a plurality of attitude control engines; each attitude control engine comprises an attitude control thrust chamber, and an oxidant control valve, a fuel control valve and an attitude control thrust chamber igniter are arranged on each attitude control thrust chamber. Propellant supply of the main engine and the attitude control engine is achieved through the unified pressurization conveying system.

The present invention belongs to the technical field of rocket engines. It relates to a monopropellant liquid propellant rocket engine, which includes a combustion chamber and a thrust chamber provided at one end of the combustion chamber. A catalytic zone is provided in the combustion chamber, and a propellant distributor connected to the propellant supply system is provided at the end of the combustion chamber away from the thrust chamber. A fragmentation and evaporation zone, a saturated steam zone, and an adsorption zone are successively provided between the propellant distributor and the catalytic zone; the saturated steam zone is used for pressure stabilization to balance the change in propellant flow rate caused by the pressure fluctuation of the propellant supply system; an adsorbent and a catalyst are provided in the adsorption zone. By increasing the resistance of the propellant from atomization / evaporation to entering the catalytic region, the fluctuation of the propellant concentration entering the surface of the catalyst is reduced. In addition, the specific surface area of the catalyst is larger than that of the adsorbent. The adsorption zone is used to increase the internal diffusion resistance of the propellant on the surface of the adsorbent during the catalytic combustion process, reduce its external diffusion resistance, improve the adsorption rate, and achieve the rapid release and adsorption of the propellant.

The invention belongs to the technical field of airspace engines, and particularly relates to a thrust chamber body structure of a liquid rocket engine, the thrust chamber body structure comprises a combustion chamber cylinder section, a throat part and a spray pipe, and the throat part is connected with an upper-section combustion chamber cylinder and a tail-section spray pipe. The body structure is composed of an inner wall, an outer wall and a cooling channel interlayer located between the inner wall and the outer wall. The thrust chamber body structure of the liquid rocket engine is characterized in that the thrust chamber body structure of the liquid rocket engine is provided with a cooling channel interlayer structure which is light and high in specific stiffness and specific strength; a liquid collecting ring and a liquid collecting cavity are arranged at the tail of the spray pipe, the liquid collecting cavity is connected with the liquid collecting ring, and the structure is used for containing and guiding cooling liquid in time. A fin support is arranged at the joint of the outer wall of the liquid collecting ring, the overall stability of the thrust chamber structure is guaranteed, and the whole system structure is manufactured through the additive manufacturing technology. The thrust chamber body structure is light, the heat transfer and exchange performance is improved, the structural thermal fatigue performance and structural stability are improved, and the service life is prolonged.

The application relates to a 3D printing-based abnormal cooling channel heat transfer evaluation method of a regenerative cooling engine, which comprises the following steps: determining an overheating area of a thrust chamber, obtaining node overheating temperature of an overheating position; carrying out heat transfer calculation of a normal cooling channel of the thrust chamber and obtaining a gas side theoretical normal wall temperature of the overheating position; assuming an initial coolant flow value of the overheating position, carrying out abnormal cooling channel heat transfer calculation to obtain a theoretical node overheating temperature; judging whether the theoretical node overheating temperature and the node overheating temperature meet a condition; if yes, taking the initial coolant flow value as a final coolant flow value; obtaining a coolant flow variation ratio alpha based on the final coolant flow value, obtaining a rated working condition flow of the abnormal cooling channel, carrying out abnormal cooling channel heat transfer calculation and evaluation, and outputting a rated abnormal cooling channel gas side wall temperature; and if the rated abnormal cooling channel gas side wall temperature is lower than a material temperature resistance limit of the thrust chamber, the rated working condition of a hot test is safe, and the test can be carried out.

The invention provides a multi-channel temperature parameter pre-cooling process liquid rocket engine temperature field display method, and solves the problems that in the prior art, as a data acquisition system measures temperature parameters of a plurality of measuring points when a power system carries out preparatory work, the temperature parameters of the key temperature measuring points are measured on the basis of a temperature sensor; and the change condition of a structural system cannot be visually displayed after long-time numerical value display. According to the method, on the basis that simulation calculation is high in time complexity and cannot achieve real-time calculation, the recurrent neural network is adopted for feature analysis, model training is completed on the basis of gradient descent, interpolation calculation of the temperature field in the channel dimension is finally achieved, the resolution of channel dimension data can be effectively improved, then the resolution of a temperature field image is improved, and the method is suitable for large-scale popularization and application. And subsequent thrust chamber temperature field modeling and analysis based on the whole pre-cooling process are facilitated.

The present application relates to aerospace propulsion technology field, specifically to a kind of open variable thrust rocket engine, rocket and control method, open variable thrust rocket engine includes: pressurizing device;Thrust chamber, with first fuel inlet, first oxidant inlet, first fuel inlet is communicated with fuel outlet fuel main road, fuel main road is equipped with fuel main valve, first oxidant inlet is communicated with oxidant outlet through oxygen main road;Gas generator, with second fuel inlet, second oxidant inlet, gas outlet, second fuel inlet is connected with fuel main road through fuel auxiliary road, second oxidant inlet is connected with oxygen main road through oxygen auxiliary road;Gas outlet is connected with the driving inlet of pressurizing device through gas main road, pressurizing device is also communicated with gas main road by gas shunt road in interior, gas shunt road is equipped with gas regulating valve, gas generated by gas generator is suitable for entering pressurizing device by gas shunt road and discharging.

The application discloses an oil injection rod and stabilizer integrated structure, which comprises a radial stabilizer, a circumferential stabilizer arranged on the radial stabilizer, a mounting seat arranged at a side end of the circumferential stabilizer, an I-way oil injection rod mounting seat, a II-way oil injection rod and a III-way oil injection rod arranged on the mounting seat, a plug and a plug arranged at a bottom end of the radial stabilizer, wherein the oil injection rod and stabilizer integrated structure is a multi-cavity and multi-hole complex hollow structure, and is provided with evaporation cavities, cooling cavities and fuel flow path cavities, and is provided with a plurality of oil injection holes, cooling holes and evaporation holes. The plug, the plug and the II-way oil injection rod and the III-way oil injection rod are connected through welding, and closed oil path cavities are formed in the inner cavities of the II-way oil injection rod and the III-way oil injection rod; the structure integrates multiple functions such as oil supply, combustion organization, on-duty flame, cooling and anti-spontaneous combustion of fuel oil, has multiple functions, and can work at a total temperature of 1450K in a thrust chamber and at a working temperature lower than 1300K, and is reliable in work.

The present invention provides a reliability assessment method for the thrust chamber of a pump-type liquid rocket engine. The method identifies the main failure modes of the thrust chamber through a fault tree deduction method. For each main failure mode of the thrust chamber, a unit reliability assessment mathematical model is determined respectively, and the probability of occurrence of each failure mode is quantitatively assessed. The reliability of the thrust chamber is quantitatively assessed by making full use of the test data of the thrust chamber and its components. The method identifies the weak links of the thrust chamber through an importance ranking method, and provides support for the development and reliability improvement of the thrust chamber.

The invention relates to a regenerative cooling engine abnormal channel heat transfer evaluation method based on 3D printing, and the method comprises the steps: determining an overheating region of a thrust chamber, and obtaining the node overheating temperature of an overheating position; heat transfer calculation of a normal cooling channel of the thrust chamber is carried out, and the theoretical normal wall temperature of the gas side at the overheating position is obtained; assuming an initial value of coolant flow at a superheat position, and carrying out heat transfer calculation of an abnormal cooling channel to obtain a theoretical node superheat temperature; judging whether the theoretical node overheating temperature and the node overheating temperature meet conditions or not; if yes, the coolant flow initial value serves as the final coolant flow; based on the final coolant flow, a coolant flow change proportion alpha is obtained, rated working condition flow of the abnormal cooling channel is obtained, heat transfer calculation evaluation of the abnormal cooling channel is carried out, and the rated abnormal cooling channel gas side wall temperature is output; and if the gas side wall temperature of the rated abnormal cooling channel is lower than the material temperature resistance limit of the thrust chamber, the rated working condition of the hot vehicle test is safe, and the test run can be carried out.

The invention discloses a method for designing a mode conversion scheme of a combined propulsion device, which comprises the following steps of: acquiring a Mach number interval of the combined propulsion device converted from a first mode to a second mode, and selecting a plurality of representative flight state points; aiming at each representative flight state point, performing discrete scanning on a mass flow range of a primary propulsion unit and a fuel-air ratio range of a secondary fuel to form a plurality of candidate working condition combinations; flow and combustion characteristics in a thrust chamber are calculated for each candidate working condition combination, and a target working condition combination capable of achieving stable combustion and meeting the thrust requirement is screened; and obtaining an optimal target working condition combination under each representative flight state point based on enumeration search, and generating a continuously enforceable mode conversion scheduling path. The method is applied to the field of combined propulsion, and efficient, stable and safe operation of the mode conversion process can be effectively achieved.

The present invention relates to the field of rocket thrust chamber manufacturing technology, and discloses a wire cutting and picking auxiliary system for additively molded parts for rocket thrust chambers. The system includes a multi-source data acquisition module, which collects geometric morphology data of additively molded parts through multiple sensors; a cutting path planning module, which generates a global cutting path sequence based on an improved genetic algorithm; a dynamic deformation compensation module, which uses a fuzzy adaptive observer to compensate for thermal deformation and vibration interference in real time; and a multi-axis collaborative control module, which constructs a distributed constraint optimization model to achieve precise control of multi-axis motion. The system can accurately collect data, reasonably plan cutting paths, effectively compensate for deformation, and ensure multi-axis collaborative work, significantly improving the quality and efficiency of wire cutting and picking of additively molded parts for rocket thrust chambers, reducing production costs, and providing strong technical support for aerospace manufacturing.

The invention relates to a ground heating control method and device for an auxiliary power system of a low-temperature carrier rocket. The method comprises the steps that a first microcontroller obtains an initial parameter and a heating initial parameter of an upper computer auxiliary power system, and obtains real-time temperatures, transmitted by a second microcontroller, of an on-rocket thrust chamber, a storage box, an electromagnetic valve and a guide pipe; controlling an on-rocket heating loop according to the initial parameters, the heating initial parameters and the real-time temperatures of an on-rocket thrust chamber, a storage tank, an electromagnetic valve and a guide pipe; and / or, the second microcontroller obtains an upper computer auxiliary power system initial parameter, a heating initial parameter and a temperature measurement loop sensor resistance value, and determines the real-time temperature of the rocket heating loop based on the heating loop current, the temperature measurement loop sensor resistance value, the initial parameter and the heating initial parameter; and the heating current and the temperature measurement temperature on the rocket are sent to a first microcontroller and an upper computer. According to the application, the blank of the ground heating control device of the auxiliary power system can be filled, and unattended operation before the auxiliary power system is launched is realized.

The present invention relates to a rocket-based combined cycle (RBCC) engine with a parallel compressor flow path, which includes a compressor flow path and an RBCC engine flow path. The compressor flow path and the RBCC engine flow path are arranged in parallel, and the two flow paths share an inlet duct. After the inlet duct flow path, a rocket thrust chamber, a primary combustion chamber, a secondary combustion chamber and a nozzle are provided; the airflow in the compressor flow path converges into the secondary combustion chamber and burns in the secondary combustion chamber.

The invention relates to the technical field of rocket engine thrust chamber additive manufacturing, and discloses a forming bin smoke exhaust system for rocket engine thrust chamber additive manufacturing. According to the system, the multi-source data acquisition module acquires smoke multi-dimensional parameters, the smoke feature extraction module decomposes features, the dynamic filtering module generates a filtering strategy and a gas flow optimization matrix, the cooperative control module coordinates equipment operation parameters to achieve energy consumption balance, and the abnormity processing module identifies abnormity and triggers a self-repairing mechanism. The system can accurately collect and analyze flue gas data, dynamically optimize filtering and airflow paths, efficiently cooperate with equipment operation, handle abnormities in time, improve the flue gas removal effect, reduce energy consumption, guarantee safety of operators and stable operation of equipment, and promote development of the rocket engine thrust chamber additive manufacturing technology.

The present invention provides a calculation method for the acoustic damping characteristics of the mixing head of a afterburning engine thrust chamber, which is used to solve the technical problem that the existing calculation method for the acoustic damping of the mixing head has low accuracy and the analysis results are quite different from the actual situation. First, according to the structural parameters of the gas flow channel of the thrust chamber, the geometric blocks of the gas flow channel are extracted; then the geometric blocks are imported into the acoustic solver to obtain the acoustic solution domains of the respective geometric blocks; after that, based on the boundary impedance equation of the honeycomb and the nozzle impedance equation, as well as the source term equation of the background mean flow, the impedance value and the dissipation source term are calculated; finally, the impedance value and the dissipation source term are respectively applied to their respective calculation domains, and the optimal structure of the afterburning engine thrust chamber can be obtained through calculation.

The invention relates to a turbojet engine afterbody comprising:-an exhaust casing (9) carrying a manifold plate (11) surrounding a primary air flow and surrounded by a secondary air flow; -an afterburner (12) comprising a casing (13) and a bushing (16) delimiting together with the casing (13) an annular space (17) extending along the axis (AX), the diameter of the bushing (16) being greater than the diameter of the converging plate (11); -a fixed shroud (22) extending from the bus plate (11) to the inner liner (16), the fixed shroud (22) comprising a window (27); -a moving cover (23) extending along the fixed cover (22) and comprising a window (28), the cover (23) being rotatably movable about an axis (AX) between an open position, in which the window (27) in the fixed cover is open, and a closed position, in which the window (27) in the fixed cover (22) is closed by the moving cover (23).

Additively manufactured thrust chambers and thrust chambers with integral fluid manifolds, and hybrid additive manufacturing methods for their production, are provided. Hybrid additive manufacturing techniques may combine a variety of processes including, WAAM, PBF, cold spray and DED, for example, to produce objects with variant dimensional requirements, i.e., large overall size and small features. Hybrid additive manufacturing may be defined as provide various process layers within any manufactured object. These process layers in turn allow for the introduction of variable feature and size distribution throughout the manufactured object. Hybrid process layers according to aspects may also allow the use of a variety of materials or may use a single material across the various process layers.

The present application provides an electric pump assisted gas generator circulation system, comprising: an oxidant supply unit having an oxidant storage portion, an oxidant pressurization portion, a first supply portion, a second supply portion, and a first detection portion; an oxidizing agent is stored in the oxidizing agent storage part, and the oxidizing agent pressurizing part can pressurize the oxidizing agent once and supply the pressurized oxidizing agent to the thrust chamber; the two ends of the second supply part communicate with the first supply part and the fuel gas generator correspondingly, and the second supply part can conduct secondary pressurization on an oxidizing agent and supply the oxidizing agent to the fuel gas generator; the other end of the fuel gas generator is connected with the oxidant pressurizing part; the first detection part is in communication connection with the oxidizing agent pressurizing part and the second supply part. According to the gas generator, the consumption of invalid propellants is reduced, the oxidant is secondarily pressurized by introducing the second supply part, and the fuel is secondarily pressurized by introducing the fourth supply part, so that the pressure of the propellants at the inlet of the gas generator is improved, the combustion efficiency of the gas generator is improved, and the gas energy generated by the propellants in unit mass is increased.

The utility model provides a variable-cross-section special-shaped gas exhaust pipe, a liquid rocket engine and a liquid rocket. The variable-cross-section special-shaped gas exhaust pipe comprises an inlet flange arranged at an inlet, an outlet flange arranged at an outlet and a barrel section arranged between the inlet flange and the outlet flange. The inlet and the outlet are eccentrically arranged; the inlet flange is connected with a turbine, and the outlet flange is connected with the thrust chamber gas collecting ring; high-temperature fuel gas enters from the inlet, flows out from the outlet, enters the thrust chamber gas collection ring and then is introduced into the spray pipe extension section. The exhaust pipe is connected between the turbine and the thrust chamber gas collecting ring, the eccentric layout design is adopted, and the problems that in the structural layout of an existing liquid rocket engine, the center distance between a turbine pump and a thrust chamber is large, the radial envelope of the whole machine is large, and the distance between a swing hose and the swing center is large are solved.

The invention relates to a thermodynamic component combined test, and provides a thermodynamic component combined test system and a test method in order to solve the problems that an afterburning cycle engine gas generator is large in propellant mass flow deviation, large in combustion product pollution, high in hot test run cost and the like. The test system comprises a fuel gas generator, a gas venturi tube and a thrust chamber which are coaxially connected; a first component access port is formed in the top of the fuel gas generator, a second component access port is formed in the side face of the fuel gas generator, the first component access port and the second component access port are used for accessing a first component propellant and a second component propellant respectively, and a plurality of fuel gas performance monitoring ports are formed in the side face of the gas venturi tube; a second component regeneration cooling access port is formed in the side surface of the outlet end of the thrust chamber and is used for regenerating and cooling a second component propellant and then accessing the second component propellant into the thrust chamber; the test method based on the test system can effectively improve the efficient, reliable and environment-friendly hot test run test between the fuel gas generator and the thrust chamber.

The invention discloses a light-weight torch ignition system and method. The light-weight torch ignition system comprises an oxidant supply unit, a fuel supply unit, a thrust chamber and an igniter. The oxidant supply unit and the fuel supply unit are both communicated with the thrust chamber so as to supply an oxidant and fuel into the thrust chamber; the oxidant supply unit is sequentially connected with an oxygen pump and an oxygen main valve through a pipeline; the oxygen main valve is connected with the thrust chamber; the fuel supply unit is sequentially connected with a fuel pump and a fuel main valve through a pipeline; the fuel main valve is connected with the thrust chamber; the igniter is mounted on the thrust chamber; an oxygen inlet and a fuel inlet are formed in the igniter; the pipeline between the oxygen pump and the oxygen main valve is communicated through a first auxiliary pipeline; the first auxiliary pipeline is connected with an igniter oxygen valve; the pipeline between the fuel pump and the fuel main valve is communicated through a second auxiliary pipeline; and an igniter oxygen valve is arranged on the first auxiliary pipeline. According to the torch ignition system, infinite ignition can be achieved, and the mass and the size of the torch ignition system are greatly reduced.