In order to increase the thrust of the air turbine ramjet, it is effective to increase the compressor pressure ratio. In particular, an air turbine ramjet engine having a large working range requires the compressor to fully function when the engine inlet temperature is increased to increase the flight speed. In this study, a compressor using a single stage to obtain a high pressure ratio was employed. This also controls the increase in quality and the like caused by the multistage of the compressor. In addition, in addition to the intake port of the ram combustion chamber, the most important of the structural elements is the length, since the overall length of the engine is given. Therefore, shortening the length of the punching combustion chamber is most effective for miniaturization of the engine. Therefore, this paper studies the structural elements of increasing thrust and reducing the size, and implements factor tests to clarify the performance of each structural element. In particular, as an engine integration technology which is a main technical subject of an air turbine ramjet, an engine test combining these elements was carried out, and matching between the respective elements was confirmed. In addition, the liquid gas generator used in this study uses a 2-liquid type with a wide range of flow control.

The engine has an outer diameter of 350mm and a total length of 1690mm. The two intake ports are mounted at the symmetrical position of the engine and are a two-dimensional shape of a variable throat type. The turbine is a front-mounted axial flow stage 1 uncooled turbine. The turbine is coupled to the compressor via a rotating shaft. A gas generator that drives the turbine is placed in front of the turbine. The gas generated by the gas generator passes through the turbine blades and is introduced into the stamping combustion chamber through the struts on the inside of the rotor blades of the compressor. The tail nozzle has a variable adjustment mechanism to expand the performance of the engine.

The gas generator gas generator generates a gas that drives the turbine, and its temperature is set to be less than 1000e that the cooling turbine can withstand. Depending on how the gas is produced, it can be divided into a solid mode and a liquid mode. In the liquid mode, there is an expander method in which heat is exchanged by using ultra-low temperature liquid hydrogen to generate gas. The manner in which the gas is produced is generally divided by the manner in which the propulsion device is used. In this study, a two-liquid gas generator that generates gas for a wide range of flow control is used. The 2-liquid type uses a liquid fuel such as a liquid rocket engine and a liquid oxidant, that is, a combination of kerosene and nitric acid, liquid hydrogen and liquid oxygen.

In the present study, the fuel of the gas generator is a jet fuel JP4 which can be shared with the stamping fuel and has a high heat release amount. The oxidant is selected to be a hydrogen peroxide water having a concentration of 60%. The hydrogen peroxide itself decomposes into water and oxygen. At this time, the water becomes water vapor to be gasified because of the heat release. When 60% of hydrogen peroxide is decomposed, 82% of water vapor and 18% of oxygen are produced by volume ratio. If the concentration is above 65%, the hydrogen peroxide will become superheated vapor. In this regard, only saturated water vapor humidity can be achieved with 60% hydrogen peroxide water, which maintains a certain humidity. However, in this study, it was considered that the treatment was convenient and the liquid layer range was wide, and the freezing point was as low as that of the jet fuel JP4 to determine the oxidant. Figure 2 shows the combustion gas temperature characteristics relative to the mixing ratio of oxidant to fuel. The driving turbine gas allows an equivalent ratio of 1000 e to be 0.51 in terms of fuel leanness and 2.4 in terms of fuel over-concentration. The reaction in terms of fuel over-concentration has the advantage of reducing the amount of oxidizing agent used. However, the generated gas is flammable and may be combusted in reaction with the gas for sealing inside the engine. Therefore, this study used a lean fuel.

JP4 does not self-ignite with hydrogen peroxide. The hydrogen peroxide having a concentration of 60% has a low reaction performance and is highly safe in handling. Therefore, the combustion performance, the ignition performance, and the outlet temperature distribution were evaluated, and an experiment for determining the state of the reaction chamber was carried out. The experiment was carried out using a 135b fan model with two lengths of reaction chambers, using a centrifugal nozzle and an evaporating tube. Both the centrifugal nozzle and the evaporation tube are placed in three circumferential positions.

For outlet gas flow and gas temperature. It is difficult to directly measure the gas flow rate, so it is determined based on the oxidant flow rate, fuel flow rate, and gasification efficiency. In the gas generator monomer working test, the gas flow rate was confirmed to be 1 kg/s, and the outlet temperature was about 950 e.

The ramjet combustion chamber is a large-scale operation corresponding to the air turbine ramjet engine, and the ramjet combustion chamber is required to be combusted under a wide range of conditions from low temperature and low pressure to high temperature and high pressure. The stamped combustion chamber in this study has a flame stabilizer. In addition, a pre-burner is provided to assist the combustion in the low temperature gas stream. Figure 9 is a schematic view of a stamped combustion chamber. The front burner is damped inside the engine below the compressor outlet to allow the incoming air to burn in the annular flame stabilizer. The combusted high temperature gas passes through the mixer to raise the temperature of the compressor outlet air. The warmed air is directed to a flame stabilizer of the stamped combustion chamber. In this way, the stamping combustion in the airflow with a low engine inlet temperature is subsidized. Moreover, in order to shorten the axial length of the punching combustion chamber, the flame stabilizer is divided into two stages in the axial direction. The flame stabilizer consists of 8 radial flare stabilizers and 1 ring stabilizer. In order to achieve a combustion efficiency of more than 80% under ground static conditions with the lowest inlet temperature and pressure, the axial length of the stamped combustion chamber is estimated to be 800 mm (L/D = 2.5). In this study, the axial length was shortened to 500 mm (L/D = 1.6). The combustion efficiency at each axial position when the flame stabilizer is 2 stages. The results of the monomer test of the stamping combustor were carried out using a 120b fan-shaped combustion chamber. The abscissa is the axial distance from the front flame stabilizer. In addition, the combustion efficiency of the single-stage flame stabilizer is the value when only the front flame stabilizer is used. The blocking ratio of each flame stabilizer was set to 36% for the front flame stabilizer and 29% for the rear flame stabilizer. The total pressure loss rate of the experimental results was 0.039 at ground static conditions. In addition, the 2-stage flame stabilizer interval was 1.4 The compressor compressor is a single stage to achieve a pressure ratio of 2.2, using front and rear straight row vane compressors. The concept of straight rows of blades before and after. In the case of a normal blade composed of one moving blade, the front and rear straight rows of blades are composed of two rows of front and rear. The blade is in the shape of a flap for an aircraft. A squeaky speed compressor with a velocity of the moving blade exceeding the speed of sound generates a shock wave from the negative pressure surface of the leading edge blade of the adjacent blade. The shock wave is coherent with the boundary layer of the blade face to cause separation. The front and rear straight rows of blades are provided with long slits between the moving blades to supply the back airflow of the rear blades. With this configuration, the separation of the airflow can be controlled even if the shape of the blade is twisted. As a result, a high pressure ratio can be achieved without increasing the number of stages of the compressor. The pressure ratio per level of the past squeaking blades is about 1.7. When the front and rear straight rows of blades are used, both the efficiency drop and the pressure ratio can be prevented. The compression section of the compressor consists of inlet guide vanes, front and rear straight rows of vanes, stator vanes and outlet guide vanes.

The inlet guide vanes are set to rectify the inlet air flow by distortion caused by a flight attitude or the like. The airflow from the front and rear straight-moving blades is greatly affected by the distortion. To eliminate the turbulence flowing into the punching combustion chamber, an outlet guide vane is disposed behind the stationary vanes. Corrected fan speed 29700r/min pressure ratio inlet total pressure adiabatic efficiency inlet total temperature fluctuation margin corrected air flow rate 2 engine operation test 2.1 Test method An engine operation test was performed to confirm the matching of various structural elements of the engine. Through this test, the basic performance of an air turbine ramjet engine that increases the compressor pressure ratio and shortens the thrust combustion chamber to increase the thrust reduction volume is obtained. The pressure ratio of the front and rear straight row vane compressors in this study is 2.2. When the pressure ratio of the general single stage compressor is 1.7, the increase of the compressor pressure ratio is estimated to increase the thrust by 64% under ground static conditions. At high altitudes (Ma=3), the thrust can be increased by 25%. Especially at low speeds, the thrust is greatly increased, so it is possible to increase the acceleration force at low speeds. The test engine is shown in Fig. 1. The test was carried out by installing a bell mouth at the engine inlet and resting on the ground in the atmosphere.

The test results use electric rotation to confirm the abnormality of the swing system. Then, the number of revolutions of the engine is increased in stages to the rated number of revolutions, and the engine is rotated after the number of revolutions of the engine reaches 90%. It is confirmed that the engine has no vibration problem and the rotation speed can be stably increased. The stamping combustion is performed after the pre-burner and the main fuel are ignited. When the number of revolutions of the engine is 91% of the rated number of revolutions, the outlet temperature of the combustion chamber is about 600e, and the combustion is about 10s. The function is confirmed, and the gas is confirmed. The stable operation of the device finally confirms the feasibility of combining the engines of the various structural elements.

In order to achieve high thrust and miniaturization of the air turbine ramjet, the engine components were studied to confirm the performance of each element. Further, in order to confirm the matching between the respective components, the engine in which the components were combined was tested on the ground stationary state, and the correctness of the design of the propulsion system was confirmed. In the future, it is planned to use the direct-connected and semi-free-jet methods to provide air in the engine corresponding to the pressure and temperature of the flight state, thereby performing an engine test simulating the flight speed and the flying height to obtain engine performance.

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