A liquid rocket engine fault simulation experiment device and method

Abstract

The present application belongs to the technical field of rocket engine simulation experiment, and discloses a liquid rocket engine fault simulation experiment device and method, which comprises a fuel supply system; a center of the injector is provided with a metal injection pipeline, and a plurality of coaxial shear nozzles are arranged around the metal injection pipeline; the center of the coaxial shear nozzle is a fuel nozzle, and the periphery of the fuel nozzle is an oxygen ring gap nozzle; the outlet of the fuel supply system is in communication with the inlet end of the fuel nozzle, and the outlet of the oxygen supply system is in communication with the inlet end of the oxygen ring gap nozzle; the outlet of the metal supply system is in communication with the inlet end of the metal injection pipeline, and is used for injecting the mixture of metal powder and engine fuel into the metal injection pipeline; the present application adds the metal impurities in the plume into the combustion chamber in the form of metal powder, realizes the simulation detection of the combustion condition in the running process of the liquid rocket engine, reduces the detection cost, shortens the detection period, and improves the safety of the detection process.

Description

Technical Field

[0001] This invention belongs to the field of rocket engine simulation experiment technology, and specifically relates to a liquid rocket engine fault simulation experiment device and method. Background Technology

[0002] As a key power unit in the aerospace field, the stability and reliability of liquid rocket engines directly affect the success of space missions. During the operation of liquid rocket engines, due to prolonged high temperature, high pressure, and high-speed airflow, the internal components of the engine undergo wear, aging, or ablation, resulting in some metallic impurities entering the engine plume and affecting the combustion status. In order to promote the reusability of spacecraft, real-time and accurate detection of the combustion status of liquid rocket engines during operation is particularly important.

[0003] Currently, the detection of combustion status during the operation of liquid rocket engines often relies on actual engine test experiments. Although this can directly reflect the actual working state of the engine, it has problems such as high cost, long cycle and difficulty in guaranteeing safety. Especially in the context of spacecraft reusability, frequent test experiments obviously do not meet the original intention of reducing costs. In addition, if the engine fails during actual test, it often causes serious consequences and may even lead to the destruction of the entire test facility. Summary of the Invention

[0004] To address the technical problems existing in the prior art, this invention provides a liquid rocket engine fault simulation experimental device and method to solve the technical problems of high cost, long cycle and difficulty in guaranteeing safety in existing methods for detecting combustion status during the operation of liquid rocket engines.

[0005] To achieve the above objectives, the technical solution adopted by the present invention is as follows: This invention provides a liquid rocket engine failure simulation experimental device, including a fuel supply system, an oxygen supply system, a metal supply system, an injector, a combustion chamber, and a tail nozzle; The injector, the combustion chamber, and the tail nozzle are connected coaxially in sequence. The injector has a metal injection pipe at its center and a plurality of coaxial shear nozzles around its perimeter; wherein, the center of the coaxial shear nozzle is a fuel nozzle and the circumferential slit around the fuel nozzle is an oxygen circumferential slit nozzle. The outlet of the fuel supply system is connected to the inlet end of the fuel nozzle, and the outlet of the oxygen supply system is connected to the inlet end of the oxygen annular nozzle. The outlet of the metal supply system is connected to the inlet of the metal injection pipeline; wherein the metal supply system is used to inject a mixture of metal powder and engine fuel into the metal injection pipeline.

[0006] Furthermore, the fuel supply system includes an air compressor, a first switching valve, a fuel tank, and a first solenoid valve; The outlet of the air compressor is connected to the inlet of the first switching valve, the inlet of the first switching valve is connected to the pressure inlet of the fuel tank, the fuel outlet of the fuel tank is connected to the inlet of the first solenoid valve, and the outlet of the first solenoid valve is connected to the inlet end of the fuel nozzle.

[0007] Furthermore, the oxygen supply system includes an oxygen storage tank, a second switching valve, and a second solenoid valve; The outlet of the oxygen storage tank is connected to the inlet of the second switching valve, the outlet of the second switching valve is connected to the inlet of the second solenoid valve, and the outlet of the second solenoid valve is connected to the inlet end of the oxygen annular nozzle.

[0008] Furthermore, the oxygen supply system also includes a nitrogen storage tank, a third switching valve, and a third solenoid valve; The outlet of the nitrogen storage tank is connected to the inlet of the third switching valve, the outlet of the third switching valve is connected to the inlet of the third solenoid valve, and the outlet of the third solenoid valve is connected to the inlet end of the oxygen annular nozzle.

[0009] Furthermore, the metal supply system includes an injection pump and a static mixing tube; The outlet of the injection pump is connected to the inlet of the static mixing tube, and the outlet of the static mixing tube is connected to the inlet end of the metal injection pipeline.

[0010] Furthermore, the injector has a fuel inlet and an oxygen inlet on its side wall, and the injector has a fuel buffer chamber and a gas buffer chamber inside, which are isolated from each other; One end of the fuel inlet is connected to the outlet of the fuel supply system, and the other end of the fuel inlet communicates with the fuel buffer chamber; the fuel buffer chamber is located at the inlet end of the fuel nozzle and communicates with the inlet end of the fuel nozzle. One end of the oxygen inlet is connected to the outlet of the oxygen supply system, and the other end of the oxygen inlet is connected to the gas buffer chamber; the gas buffer chamber is located at the inlet end of the oxygen annular nozzle and is connected to the inlet end of the oxygen annular nozzle.

[0011] Furthermore, the diameter of the fuel nozzle is 0.6 mm; the diameter of the oxygen annular nozzle is 3.6 mm, with an equivalent diameter of 2.0 mm.

[0012] Furthermore, the combustion chamber is designed according to predetermined combustion chamber modeling criteria; wherein, the predetermined combustion chamber modeling criteria include Reynolds number similarity criteria and Damcole number similarity criteria.

[0013] Furthermore, the tail nozzle is a Laval nozzle.

[0014] The present invention also provides a liquid rocket engine fault simulation experiment method, utilizing the aforementioned liquid rocket engine fault simulation experiment device; Experimental methods include: Turn on the oxygen supply system and continuously supply oxygen into the oxygen annular nozzle of the injector. Ignite in the combustion chamber and turn on the fuel supply system to continuously supply engine fuel into the fuel nozzle of the injector. Turn on the metal supply system and introduce a mixture of metal powder and engine fuel into the metal injection line of the injector; The excitation state of the metal powder was tested at the tail nozzle; After the test is completed, the metal supply system, fuel supply system and oxygen supply system are shut off in sequence, and the flame is extinguished.

[0015] Compared with the prior art, the beneficial effects of the present invention are as follows: The liquid rocket engine failure simulation experimental device provided by this invention provides simulated propellant for operation by setting up a fuel supply system and an oxygen supply system. A metal supply system injects a mixture of metal powder and engine fuel into the metal injection pipe of the injection tube, adding metal impurities from the plume into the combustion chamber in the form of metal powder. This simulates the metals that may be entrained during the combustion process of a liquid rocket engine, thereby simulating the combustion status of the liquid rocket engine during operation. This effectively reduces detection costs, shortens the detection cycle, and improves the safety of the detection process, strongly promoting the development of reusable spacecraft. Furthermore, using engine fuel as a carrier for the metal powder protects the metal powder from oxidation and prevents the introduction of other working fluids, improving the accuracy and realism of the simulation results.

[0016] The liquid rocket engine fault simulation experiment method provided by this invention has all the advantages of the aforementioned liquid rocket engine fault simulation experiment device. Attached Figure Description

[0017] Figure 1 A schematic diagram of the overall structure of the liquid rocket engine fault simulation experimental device provided in the embodiment; Figure 2 This is a three-dimensional structural diagram of the injector in the embodiment; Figure 3 This is a partial structural diagram of the injector in the embodiment; Figure 4 This is a cross-sectional view of the injector in the embodiment; Figure 5 This is a three-dimensional structural diagram of the combustion chamber in the embodiment; Figure 6 This is a cross-sectional view of the combustion chamber in the embodiment; Figure 7 This is a three-dimensional structural diagram of the tail nozzle in the embodiment; Figure 8 This is a cross-sectional view of the tail nozzle in the embodiment; Figure 9 This is a graph showing the experimental temperature range of the injector in the embodiment; Figure 10 The graph shows the relationship between the mixing ratio of oxygen and kerosene and the temperature of the combustion gas in the example. Figure 11 This is a schematic diagram of the experimental timing of the liquid rocket engine fault simulation experimental device in the embodiment; Figure 12 This is a photograph of high-temperature, high-speed metal powder at the tail nozzle in the embodiment. Figure 13 This is a comparison of the emission spectra of pure kerosene combustion and combustion with added copper powder.

[0018] Among them, 1 is an air compressor, 2 is a first switching valve, 3 is a pressure relief valve, 4 is a fuel tank, 5 is a first solenoid valve, 6 is a first flow meter; 7 is an oxygen tank, 8 is a second switching valve, 9 is a second flow meter, 10 is a second solenoid valve; 11 is a nitrogen tank, 12 is a third switching valve, 13 is a third solenoid valve; 14 is an injection pump, 15 is a static mixing pipe, 16 is an injector, 17 is a combustion chamber, 18 is a tail nozzle; 19 is a control subsystem; 161 is a fuel inlet, 162 is a fuel buffer chamber, 163 is an oxygen inlet, 164 is a gas buffer chamber, 165 is a metal injection pipeline, 166 is a fuel nozzle, and 167 is an oxygen annular nozzle. Detailed Implementation

[0019] To make the technical problems solved by the present invention, the technical solutions, and the beneficial effects clearer, the following specific embodiments provide a further detailed description of the present invention. It should be understood that the specific embodiments described herein are merely illustrative and are not intended to limit the scope of the invention.

[0020] Before describing the specific embodiments of this application, some of the technical terms involved in the embodiments of this application are explained as follows: Plume: The high-temperature, high-speed gas stream ejected from a rocket engine, whose state directly reflects the engine's combustion efficiency and operating status. The main components of a liquid rocket engine's plume include high-temperature combustion gases and incompletely burned carbon particles and metal oxide particles. The metal content and combustion gas composition in the plume are key information for judging the operating status of a liquid rocket engine.

[0021] It should be noted that the material ablation and wear that occur inside a liquid rocket engine under high temperature and pressure can trigger characteristic spectral radiation. By identifying the characteristic metal peaks in the characteristic spectral radiation, the operating status of the engine's internal structure can be determined. In this invention, metal powder is added to simulate the metal in the plume, thereby simulating the metal impurities that may be carried in the rocket engine during combustion.

[0022] The present invention provides a liquid rocket engine failure simulation experimental device, including a fuel supply system, an oxygen supply system, a metal supply system, an injector 16, a combustion chamber 17, and a tail nozzle 18.

[0023] The injector 16, the combustion chamber 17, and the tail nozzle 18 are coaxially connected in sequence; a metal injection pipe 165 is provided at the center of the injector 16, and a plurality of coaxial shear nozzles are provided around the metal injection pipe 165; wherein, the center of the coaxial shear nozzle is a fuel nozzle 166, and the circumferential slits around the fuel nozzle are oxygen circumferential slit nozzles 167.

[0024] The outlet of the fuel supply system is connected to the inlet end of the fuel nozzle 166, and the outlet of the oxygen supply system is connected to the inlet end of the oxygen annular nozzle 167; the outlet of the metal supply system is connected to the inlet end of the metal injection line 165; wherein, the metal supply system is used to inject a mixture of metal powder and engine fuel into the metal injection line 165.

[0025] In the above embodiments, by setting the fuel supply system, oxygen supply system, and metal supply system to be connected to the corresponding fuel nozzle, oxygen annular nozzle, and metal injection pipeline on the injector, respectively, the scenario of metal impurities entering the plume during engine operation can be simulated. This enables the simulation detection of the combustion status during the operation of the liquid rocket engine. Without the need for actual engine test experiments, various fault conditions can be simulated in the experimental device, effectively avoiding the problems of high cost and long cycle of actual test experiments. At the same time, it ensures experimental safety and provides a reliable and economical means for real-time and accurate detection of engine combustion status.

[0026] The following detailed explanation of the liquid rocket engine fault simulation experimental device provided by the present invention uses some specific embodiments: Example 1 As attached Figure 1-8 As shown, this embodiment 1 provides a liquid rocket engine fault simulation experimental device, including a fuel supply system, an oxygen supply system, a metal supply system, an injector 16, a combustion chamber 17, a tail nozzle 18, and a control subsystem 19.

[0027] The fuel supply system is used to continuously supply engine fuel to the injector 16; wherein the outlet of the fuel supply system is connected to the inlet end of the fuel nozzle 166 in the injector 16; specifically, the fuel supply system includes an air compressor 1, a first switching valve 2, a pressure relief valve 3, a fuel tank 4, a first solenoid valve 5, and a first flow meter 6.

[0028] The outlet of the air compressor 1 is connected to the inlet of the first switching valve 2, the inlet of the first switching valve 2 is connected to the pressure inlet of the fuel tank 4, the fuel outlet of the fuel tank 4 is connected to the inlet of the first solenoid valve 5, and the outlet of the first solenoid valve 5 is connected to the inlet of the fuel nozzle 166; the pressure relief valve 3 is disposed between the first switching valve 2 and the fuel tank 4, and the first flow meter 6 is disposed between the outlet of the first solenoid valve 5 and the inlet of the fuel nozzle 166.

[0029] The air compressor 1 is used to provide fuel pressure to the fuel tank 4 so that the engine fuel in the fuel tank 4 flows out at a preset pressure; the first switching valve 2 is used to control the switching of the fuel supply system; the pressure relief valve 3 is used to maintain the operating pressure of the fuel supply system to prevent accidents caused by excessive pressure; the fuel tank 4 is used to store engine fuel; preferably, the engine fuel is kerosene; the first solenoid valve 5 is used to control the flow rate of the engine fuel; the first flow meter 6 is used to collect the flow rate of the engine fuel.

[0030] The oxygen supply system is used to continuously supply oxygen and nitrogen into the injector 16; wherein, the outlet of the oxygen supply system is connected to the inlet end of the oxygen annular nozzle 167 of the injector 16; specifically, the oxygen supply system includes an oxygen branch and a nitrogen branch, the oxygen branch being used to continuously supply oxygen into the injector 16, and the nitrogen branch being used to continuously supply nitrogen into the injector 16.

[0031] The oxygen branch includes an oxygen storage tank 7, a second switching valve 8, a second flow meter 9, and a second solenoid valve 10. The outlet of the oxygen storage tank 7 is connected to the inlet of the second switching valve 8, the outlet of the second switching valve 8 is connected to the inlet of the second solenoid valve 10, and the outlet of the second solenoid valve 10 is connected to the inlet end of the oxygen annular nozzle 167. The second flow meter 9 is disposed between the second switching valve 8 and the second solenoid valve 10. The oxygen storage tank 7 is used to store oxygen. The second switching valve 8 is used to control the switching of the oxygen branch. The second flow meter 9 is used to collect the flow rate of oxygen in the oxygen branch. The second solenoid valve 10 is used to control the flow rate of oxygen in the oxygen branch.

[0032] The nitrogen branch includes a nitrogen storage tank 11, a third switching valve 12, and a third solenoid valve 13; the outlet of the nitrogen storage tank 11 is connected to the inlet of the third switching valve 12, the outlet of the third switching valve 12 is connected to the inlet of the third solenoid valve 13, and the outlet of the third solenoid valve 13 is connected to the inlet end of the oxygen annular nozzle 167; the nitrogen storage tank 11 is used to store nitrogen; the third switching valve 12 is used to control the switching of the nitrogen branch; and the third solenoid valve 13 is used to control the flow rate of nitrogen in the nitrogen branch.

[0033] The metal supply system is used to continuously supply metal powder into the injector 16 to simulate metal impurities in the plume of a liquid rocket engine. The outlet of the metal supply system is connected to the inlet of a metal injection line 165 in the injector 16 to inject a mixture of metal powder and engine fuel into the metal injection line 165. Specifically, the metal supply system includes an injection pump 14 and a static mixing pipe 15. The outlet of the injection pump 14 is connected to the inlet of the static mixing pipe 15, and the outlet of the static mixing pipe 15 is connected to the inlet of the metal injection line 165. Preferably, the injection pump 14 is an SPLab01 type injection pump.

[0034] It should be noted that the mixture of metal powder and engine fuel, as a mixed suspension, is continuously and intensely mixed and diffused through the static mixing tube 15 until the mixture is achieved, and then injected into the combustion chamber 17. During the process of injecting the mixture of metal powder and engine fuel into the metal injection line 165 using the metal supply system, the metal powder is first injected into a syringe. After mixing the engine fuel and metal powder in the syringe, a gel-like state is formed. The piston pushes the metal powder through the metal injection line 165 into the combustion chamber 17. The static mixing tube 15 is a transparent... A static mixing tube, using a transparent viewing window, allows observation of the dispersion state of metal powder within the tube. The static mixing tube contains only stationary elements; fluid flowing within it impacts various elements, generating intense turbulent flow that creates strong shear forces that further separate and mix the fluid, forming the desired emulsion. During solid-liquid mixing, the static mixing tube effectively increases turbulence, improves mixing uniformity, and prevents sedimentation or aggregation of solid particles. Preferably, the static mixing tube 15 is an SK static mixing tube, particularly suitable for high-viscosity (≤10...) 6 This method is suitable for fluid or solid-liquid mixing processes (Pa·s), especially for viscous media with low flow rates and impurities. For RP-3 kerosene, the viscosity coefficient is 1.0 × 10⁻⁶ Pa·s. -3 Pa·s meets the requirements for viscosity and solid-liquid mixing, and is therefore applicable.

[0035] The injector 16, the combustion chamber 17, and the tail nozzle 18 are coaxially connected in sequence; wherein, the head of the injector 16 and the tail end of the combustion chamber 17, and the tail end of the combustion chamber 17 and the head of the tail nozzle 18 are fixedly connected by flanges; a metal injection pipe 165 is provided at the center of the injector 16, and a plurality of coaxial shear nozzles are provided around the metal injection pipe 165; wherein, the center of the coaxial shear nozzle is a fuel nozzle 166, and the circumferential seam of the fuel nozzle is an oxygen circumferential seam nozzle 167.

[0036] Specifically, the injector 16 has a cylindrical structure, and the metal injection pipe 165 is located at the center of the injector 16. The inlet end of the metal injection pipe 165 extends to the outer side of the tail end of the injector 16 and is connected to the outlet of the static mixing pipe 15 in the metal supply system. The outlet end of the metal injection pipe 165 is flush with the head of the injector 16. A fuel inlet 161 and an oxygen inlet 163 are provided on the side wall of the injector 16. The fuel inlet 161 serves as a channel for engine fuel to enter the injector 16, and the oxygen inlet 163 serves as a channel for oxygen and nitrogen to enter the injector 16. The fuel inlet 161 is located near the tail end of the injector 16, and the oxygen inlet 163 is located between the fuel inlet and the middle section of the injector 16.

[0037] The injector 16 has a fuel buffer chamber 162 and a gas buffer chamber 164 inside. The position of the fuel buffer chamber 162 is adapted to the position of the fuel inlet 161, and the position of the gas buffer chamber 164 is adapted to the position of the oxygen inlet 163. One end of the fuel inlet 161 is connected to the outlet of the first solenoid valve 5 in the fuel supply system, and the other end of the fuel inlet 161 is connected to the fuel buffer chamber 162. The fuel buffer chamber 162 is located at the inlet end of the fuel nozzle 166 and is connected to the inlet end of the fuel nozzle 166. One end of the oxygen inlet 163 is connected to the outlet of the second solenoid valve 10 and the outlet of the third solenoid valve 13 in the oxygen supply system, and the other end of the oxygen inlet 163 is connected to the gas buffer chamber 164. The gas buffer chamber 164 is located at the inlet end of the oxygen annular nozzle 167 and is connected to the inlet end of the oxygen annular nozzle 167.

[0038] It should be noted that the injector 16 has a plurality of nozzle channels inside, which are evenly distributed in a ring around the metal injection pipe 165; wherein, the nozzle channels are arranged along the axial direction of the injector 16, one end of the nozzle channel is connected to the gas buffer chamber 164, and the other end of the nozzle channel is flush with and connected to the head of the injector 16.

[0039] Each nozzle channel is coaxially provided with a fuel nozzle 166. One end of the fuel nozzle 166 passes through the gas buffer chamber 164 and extends towards the fuel buffer chamber 162, communicating with the fuel buffer chamber 162. The other end of the fuel nozzle 166 is flush with and communicates with the head of the injector 16. Preferably, the fuel nozzle 166 is a tubular nozzle. The fuel nozzle 166 and the nozzle channel cooperate with each other to form a coaxial shearing nozzle. The annular gas channel formed between the fuel nozzle 166 and the nozzle channel serves as the oxygen annular nozzle 167. One end of the oxygen annular nozzle 167 communicates with the gas buffer chamber 164, and the other end of the oxygen annular nozzle 167 is flush with and communicates with the head of the injector 16.

[0040] It should be noted that at atmospheric pressure (approximately 101 kPa), the auto-ignition temperature range of RP-3 aviation kerosene is approximately 483K-553K, and the coking temperature in its flowing state is 423-435K. The injection temperature of aviation kerosene needs to be determined based on the specific engine design, fuel type, and mission requirements. Normal-temperature kerosene (such as RP-1) is typically maintained at 300K-310K, while low-temperature kerosene may have optimized atomization performance within the range of 238K-258K, but this requires compensation for the changes in physical properties caused by low temperatures through injector design. In practical applications, it is necessary to combine thermal management systems and combustion stability testing to ensure that temperature parameters meet comprehensive performance requirements. For example... For example, the Saturn V F-1 engine, which uses kerosene and liquid oxygen propellants, achieves stable combustion by designing a high-efficiency injector, as the kerosene temperature is close to room temperature. The YF-100 and YF-115 engines of the Long March 7 use RP-1 kerosene and liquid oxygen propellants, with the injection temperature controlled at 298K-308K to balance atomization efficiency and system reliability. When using injectors, strategies to reduce temperature rise are needed by increasing thermal resistance and decreasing heat capacity. For example, using capillary tubes with a large length-to-diameter ratio (30-40) and a small diameter (0.5-5mm) as nozzles or supply pipes can effectively reduce heat transfer to the liquid collection chamber.

[0041] In this embodiment 1, the length-to-diameter ratio of the injector pipe is greater than 80, which can fully develop turbulence and increase thermal resistance, reducing heat transfer to the liquid collection chamber; reducing heat capacity: reducing the volume and mass of the injector; reduced volume and mass mean a reduced total heat capacity, thus the temperature rise rate will be slower under the same heat input; since the injector volume of the scaled-down engine is smaller than that of the actual engine, it naturally has a smaller heat capacity and a slower temperature rise rate; the coaxial shear injector (such as the RL-10 engine) achieves efficient atomization through the shearing action of the oxidant and fuel, while reducing heat accumulation; in this embodiment 1, an injector with a coaxial shear nozzle is used to cool the kerosene using oxygen; the experiment found that the temperature rise of kerosene at room temperature does not exceed 20K, such as Figure 9As shown, it meets the safety requirements for the use of the injector.

[0042] The combustion chamber 17 is equipped with a spark plug and a pressure sensor. The spark plug is used for ignition, and the pressure sensor is used to collect the internal pressure data of the combustion chamber. The combustion chamber is designed according to a predetermined combustion chamber modeling criterion. The predetermined combustion chamber modeling criterion includes the Reynolds number similarity criterion and the Damcole number similarity criterion. It should be noted that a reserved installation port is provided on the side wall of the combustion chamber 17 for installing the spark plug or the pressure sensor.

[0043] Specifically, the design principle of the combustion chamber 17 is explained as follows: The flow field of the combustion chamber 17 is an unsteady, subsonic, highly turbulent, multi-component two-phase reactive flow. Therefore, based on dimensional analysis, similarity criteria in the engine combustion chamber can be obtained using the two-phase reactive flow rate equation and energy equation. Designing the combustion chamber according to these similarity criteria ensures the comparability of the main two-phase reactive flow processes under experimental and real conditions, providing important guidance for the design and implementation of the combustion chamber. The pre-determined combustion chamber modeling criteria involve flow similarity criteria and combustion reaction similarity criteria, as detailed below: 1) Flow similarity criterion The flow similarity criterion adopts the Reynolds number similarity criterion; the Reynolds number similarity criterion guarantees that the Reynolds numbers of the baseline model and the geometrically modeled model are equal; whereby the Reynolds number is defined as:

[0044] in, Reynolds numbers for the baseline and geometrically modeled types; Fluid density, kg / m³ 3 ; The viscosity is dynamic, Pa·s; The fluid velocity is in m / s; The characteristic dimension (or diameter) of the combustion chamber is in meters (m).

[0045] The medium in the injector is preheated by the heat generated from combustion. Thermocouples arranged in the gas and fuel buffer chambers of the injector are used to measure the temperatures of oxygen and kerosene. Oxygen is heated to 670K, and kerosene is heated to a supercritical temperature, such as 670K, to ensure a near-realistic chemical reaction rate. By using the same working fluid as liquid rocket engines, the density and viscosity of the working fluid can be considered constant at the same temperature. Generally, Re < 2300 is considered laminar flow, Re = 2300-4000 is transitional flow, Re > 4000 is turbulent flow, and Re > 10000 is fully turbulent flow. In practice, Re > 3000 is sufficient to determine turbulence. When the length-to-diameter ratio of the injector pipe is greater than 80, it is considered fully developed turbulence. In actual injector operation, the Reynolds number of the oxygen annular nozzle reaches 10. 5 Above, the Reynolds number of the fuel nozzle reaches 10. 4 The above indicates that, under actual operating conditions, the Reynolds number of oxygen at the oxygen annular nozzle outlet of the injector reaches over 45,000, and the Reynolds number of engine fuel at the fuel nozzle reaches 10. 4 The above satisfies the Reynolds number similarity criterion.

[0046] 2) Combustion reaction similarity criterion The combustion reaction similarity criterion adopts the Damköhler number criterion; the Damköhler number similarity criterion guarantees the Damköhler first criterion number for both the baseline model and the geometrically modeled model. equal; It represents the ratio of gas residence time to chemical reaction time, the Damköhler first criterion. Defined as:

[0047] in, The characteristic time of a chemical reaction is expressed in seconds (s). The characteristic dimension (or diameter) of the combustion chamber is in meters (m). The fluid velocity is expressed in m / s.

[0048] Since the chemical reaction rate is not affected by geometric modeling, under similar combustion conditions... Approximately a constant; Damköhler's first criterion number Only related to the characteristic length of the combustion chamber L C and fluid velocity Related; generally considered to be the length of the combustion chamber L Equal to the length of the cylindrical section of the combustion chamber L c1 and nozzle convergence length L c2 The sum; once the combustion chamber length is determined, the combustion chamber length... L The longer the length, the greater the volume of the combustion chamber.V The larger the diameter, the longer the combustion gases remain in the combustion chamber; the selected combustion chamber length should ensure sufficient volume so that the residence time of the combustion gases exceeds the time required for complete combustion of the propellant, thereby achieving higher combustion efficiency; tests have shown that for 80%-90% of liquid propellants, the combustion process can be completed within a length of 100-120 mm; therefore, the total length of the combustion chamber... L C The diameter is 170mm, which meets the design requirements.

[0049] Liquid rocket engine exhaust residence time is typically 1-8 ms, with a combustion chamber design length of 170 mm. To meet the required residence time, the exhaust velocity is 22-170 m / s. At a combustion chamber pressure of 100 kPa, the nozzle throat velocity is 300 m / s. Clearly, the nozzle throat velocity is achievable. Therefore, according to the Damköhler first criterion... It is similar to a rocket engine.

[0050] Design performance specifications: In this embodiment 1, the engine fuel is kerosene, meaning the propellant is designed to be oxygen and kerosene; the pressure of the combustion chamber 17 is 0.6 MPa, the specific impulse is 208 s, the thrust is 20 N, and the characteristic velocity is 1715 m / s; the designed length of the combustion chamber 17 is 110 mm, and the designed length of the tail nozzle 18 is 270 mm; the diameter of the combustion chamber 17 is 26 mm; the tail nozzle 18 is a Laval nozzle and is a standard conical nozzle; the throat diameter of the Laval nozzle is 5.8 mm, and the exit diameter is 20 mm; the convergence angle of the tail nozzle 18 is 30°, and the nozzle expansion half-angle is 15°; the injector 16 is designed with six coaxial shear nozzles; metal injection piping. 165 is used to inject a mixture of metal powder and engine fuel. The engine fuel protects the metal from oxidation and prevents the introduction of other working fluids. Inside the combustion chamber 17, the flame envelops the central metal powder for a chemical reaction. The injection of powder can be individually controlled, and the injection time can be controlled by the third solenoid valve 13. The inner diameter of the metal injection pipe 165 is designed to be 1 mm to prevent metal powder from clogging the channel. The metal injection pipe 165 is set through the fuel buffer chamber and the gas buffer chamber, which can effectively preheat the engine fuel and oxygen. The diameter of the fuel nozzle 166 is 0.6 mm. The diameter of the oxygen annular nozzle 167 is 3.6 mm, with an equivalent diameter of 2.0 mm.

[0051] The engine fuel-oxygen mixing ratio is 1.2-12 to ensure adjustable combustion gas temperature, covering a temperature range of 1200-3600K. The design principle is based on the assumption that the engine fuel uses a fuel with the molecular formula C. 12 H 26 The kerosene was burned with oxygen, and the thermodynamic calculations were performed, as shown in the attached figure. Figure 10As shown; since the temperature of oxygen and kerosene at the equivalent mixing ratio is 3340K, the temperature range of 1200-3600K is met under the condition of a mixing ratio of 1.2-12.

[0052] A scaled-down model of the coaxial bicomponent shear nozzle commonly used in liquid rocket engines was adopted as the coaxial shear nozzle. Reducing the diameter of the coaxial shear nozzle significantly decreased fuel consumption and effectively saved experimental costs. Specifically, the fuel nozzle diameter was designed to be 0.6 mm, and the oxygen annular nozzle diameter was designed to be 3.6 mm. Compared to the typical 6-12 mm inner diameter and approximately 10-15 mm inner diameter of the outer annular tube in a liquid rocket engine coaxial bicomponent nozzle, this smaller nozzle diameter makes it more susceptible to Rayleigh-Taylor (RT) instability due to gas-liquid shear. With a smaller diameter, the jet is more affected by aerodynamic forces, the liquid filament breaks up more easily, which is beneficial for atomization and thus improves combustion efficiency.

[0053] Secondly, due to the complex composition of kerosene, C is commonly used. 12 H 26 C 10 H 22 Single component substitution; under ideal combustion conditions (no high-temperature dissociation, no impurities), kerosene (C 12 H 26 It reacts with oxygen (O2) to produce carbon dioxide (CO2) and water (H2O), as shown in the following reaction equation:

[0054] The above reaction formula is based on the complete oxidation characteristics of hydrocarbons, assuming a combustion efficiency of 100%. Therefore, 1 mol of kerosene requires 18.5 mol of oxygen for complete combustion. Ideally, the area ratio of oxygen to kerosene is 18.5. The designed coaxial shear nozzle has an oxygen to kerosene area ratio of 11. At this point, the oxygen channel cross-section is reduced, which is beneficial for gas acceleration and meets the working conditions of the shear nozzle. When the mixture ratio O / F is 1.5 under fuel-rich conditions, the oxygen to kerosene ratio is 23. The shearing and diffusion effects of oxygen on kerosene enhance the mixing process of kerosene and oxygen.

[0055] Example 2 This embodiment 2 provides a liquid rocket engine fault simulation experiment method, utilizing the liquid rocket engine fault simulation experiment device described in embodiment 1 above; the method includes: turning on the oxygen supply system and continuously introducing oxygen into the oxygen annular nozzle 167 of the injector 16; igniting in the combustion chamber 17; turning on the fuel supply system and continuously introducing engine fuel into the fuel nozzle 166 of the injector 16; turning on the metal supply system and introducing a mixture of metal powder and engine fuel into the metal injection pipe 165 of the injector 16; testing the excitation state of the metal powder at the tail nozzle 18; after the test is completed, sequentially turning off the metal supply system, fuel supply system and oxygen supply system, and blowing out the flame.

[0056] Specifically, the experimental process includes the experimental preparation stage and the experimental process stage.

[0057] The experimental preparation stage includes: first, replenishing the engine fuel in the fuel tank 4, turning on the air compressor 1 to provide fuel pressure, then turning on the oxygen tank 7 to adjust the oxygen pressure, turning on the nitrogen tank 11 to adjust the nitrogen pressure, and injecting the engine fuel and metal powder into the syringe on the injection pump 14.

[0058] The experimental process includes: opening the second solenoid valve 10 to continuously supply oxygen; sending an ignition signal to start the spark plug 171 for ignition; opening the first solenoid valve 5 to inject engine fuel; closing the spark plug 171 and adjusting the flow rate according to the operating conditions to ensure the mixture ratio; continuing combustion and maintaining stable combustion for a preset time period; triggering a synchronization signal to activate the injection pump 14, which delivers engine fuel and metal powder into the combustion chamber 17; operating the spectrometer to test the excitation state of the metal powder at the tail nozzle 18; and after the test is completed, closing the first solenoid valve 5 first, then closing the second solenoid valve 10, and opening the third solenoid valve 13 to purge the flame.

[0059] Specifically, the timeline diagrams for each stage of the experiment are attached. Figure 11 As shown; wherein, at time T, the second solenoid valve 10 is opened; at time T+5s, the spark plug is activated; at time T+10s, the first solenoid valve 5 is opened; at time T+20s, the spark plug is closed; at time T+20s, the flow rate is adjusted according to the operating conditions; at time T+145s, the synchronization signal is triggered, and the injection pump 14 operates; at time T+155s, the first solenoid valve 5 is closed; at time T+160s, the second solenoid valve 10 is closed; at time T+165s, the third solenoid valve 11 is opened; at time T+180s, the third solenoid valve 11 is closed.

[0060] Experimental results: In this Example 2, in order to verify whether the metal powder was successfully added into the combustion chamber and successfully ignited, the following three experiments were conducted.

[0061] (1) Using a high-speed camera at a frame rate of 1000fps, the metal powder injection process was filmed at the tail nozzle. No supplementary lighting was provided; the filming relied on the self-illumination of the flame. (See attached image.) Figure 12 As shown, the metal powder, after being heated in the combustion chamber section, appears as particles scattered at the nozzle along the expansion angle of 30°, which coincides with the 15° expansion half-angle of the tail nozzle; among them, the metal powder particles are incandescent and emit bright white light; the high-speed camera confirmed that the metal powder has been added into the combustion chamber and successfully ejected from the nozzle.

[0062] (2) In order to verify the plume temperature, the temperature was detected by the metal wire melting point calibration method. The detection principle is that when the metal wire comes into contact with the jet, the temperature will exceed the metal melting point and the metal wire will melt. Molybdenum wire and tungsten wire were used respectively. The melting point of molybdenum wire is 2890K and the melting point of tungsten wire is 3686K. The experimental test results showed that the molybdenum wire melted, while the tungsten wire did not melt. Therefore, it can be determined that the combustion temperature is between 2900-3600K.

[0063] (3) Measure the emission spectrum of the metal powder after its addition using a spectrometer, with an integration time of 1000 ms and a sampling interval of 100 ms; see attached. Figure 13 As shown, attached Figure 13 The figure shows a comparison of the emission spectra of pure kerosene combustion and combustion with added copper powder; from the appendix Figure 13 As can be seen, 324.7 and 327.3 are two characteristic peaks of copper, with intensities of 6704 and 6126, respectively. This verifies that copper is already in an excited state, the addition of metal powder is effective, and further verifies the effectiveness of the liquid rocket engine fault simulation experimental device in detecting faults by adding metal powder.

[0064] The above embodiments are merely one of the implementation methods for achieving the technical solution of the present invention. The scope of protection claimed by the present invention is not limited to this embodiment, but also includes any variations, substitutions and other implementation methods that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention.

Claims

1. A liquid rocket engine fault simulation experimental device, characterized in that, It includes a fuel supply system, an oxygen supply system, a metal supply system, an injector (16), a combustion chamber (17), and a tailpipe (18). The injector (16), the combustion chamber (17) and the tail nozzle (18) are connected coaxially in sequence; The injector (16) has a metal injection pipe (165) at its center, and a plurality of coaxial shear nozzles are arranged around the metal injection pipe (165); wherein, the center of the coaxial shear nozzle is a fuel nozzle (166), and the circumferential seam around the fuel nozzle is an oxygen circumferential seam nozzle (167). The outlet of the fuel supply system is connected to the inlet end of the fuel nozzle (166), and the outlet of the oxygen supply system is connected to the inlet end of the oxygen annular nozzle (167). The outlet of the metal supply system is connected to the inlet of the metal injection line (165); wherein the metal supply system is used to inject a mixture of metal powder and engine fuel into the metal injection line (165).

2. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The fuel supply system includes an air compressor (1), a first switching valve (2), a fuel tank (4), and a first solenoid valve (5); The outlet of the air compressor (1) is connected to the inlet of the first switching valve (2), the inlet of the first switching valve (2) is connected to the pressure inlet of the fuel tank (4), the fuel outlet of the fuel tank (4) is connected to the inlet of the first solenoid valve (5), and the outlet of the first solenoid valve (5) is connected to the inlet end of the fuel nozzle (166).

3. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The oxygen supply system includes an oxygen storage tank (7), a second switching valve (8), and a second solenoid valve (10). The outlet of the oxygen storage tank (7) is connected to the inlet of the second switching valve (8), the outlet of the second switching valve (8) is connected to the inlet of the second solenoid valve (10), and the outlet of the second solenoid valve (10) is connected to the inlet end of the oxygen annular nozzle (167).

4. The liquid rocket engine fault simulation experimental device according to claim 3, characterized in that, The oxygen supply system also includes a nitrogen storage tank (11), a third switching valve (12), and a third solenoid valve (13). The outlet of the nitrogen storage tank (11) is connected to the inlet of the third switching valve (12), the outlet of the third switching valve (12) is connected to the inlet of the third solenoid valve (13), and the outlet of the third solenoid valve (13) is connected to the inlet end of the oxygen annular nozzle (167).

5. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The metal supply system includes an injection pump (14) and a static mixing tube (15). The outlet of the injection pump (14) is connected to the inlet of the static mixing tube (15), and the outlet of the static mixing tube (15) is connected to the inlet end of the metal injection line (165).

6. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The injector (16) has a fuel inlet (161) and an oxygen inlet (163) on its side wall. The injector (16) has a fuel buffer chamber (162) and a gas buffer chamber (164) inside it. The fuel buffer chamber (162) and the gas buffer chamber (164) are isolated from each other. One end of the fuel inlet (161) is connected to the outlet of the fuel supply system, and the other end of the fuel inlet (161) is connected to the fuel buffer chamber (162); the fuel buffer chamber (162) is located at the inlet end of the fuel nozzle (166) and is connected to the inlet end of the fuel nozzle (166); One end of the oxygen inlet (163) is connected to the outlet of the oxygen supply system, and the other end of the oxygen inlet (163) is connected to the gas buffer chamber (164); the gas buffer chamber (164) is located at the inlet end of the oxygen annular nozzle (167) and is connected to the inlet end of the oxygen annular nozzle (167).

7. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The diameter of the fuel nozzle (166) is 0.6 mm; the diameter of the oxygen annular nozzle (167) is 3.6 mm, with an equivalent diameter of 2.0 mm.

8. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The combustion chamber (17) is designed according to the predetermined combustion chamber modeling criteria; wherein the predetermined combustion chamber modeling criteria include the Reynolds number similarity criterion and the Damcole number similarity criterion.

9. The liquid rocket engine fault simulation experimental device according to claim 1, characterized in that, The tail nozzle (18) is a Laval nozzle.

10. A method for simulating faults in a liquid rocket engine, characterized in that, Using a liquid rocket engine failure simulation experimental apparatus as described in any one of claims 1-9; Experimental methods include: Turn on the oxygen supply system and continuously supply oxygen into the oxygen annular nozzle (167) of the injector (16); Ignite in the combustion chamber (17) and turn on the fuel supply system to continuously supply engine fuel into the fuel nozzle (166) of the injector (16); Turn on the metal supply system and introduce a mixture of metal powder and engine fuel into the metal injection line (165) of the injector (16); The excitation state of the metal powder was tested at the tail nozzle (18); After the test is completed, the metal supply system, fuel supply system and oxygen supply system are shut off in sequence, and the flame is extinguished.

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