• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.324-327
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• 2008

Two stage startup of high thrust liquid rocket engine can reduce the abrupt impulse to the vehicle and engine by changing oxidizer flow rate to the combustion chamber. Also it ensures stable ignition of combustion chamber against hard start and to prevent pump stall by the sudden supply of large mass flow rate. However high discharge pressure of oxidizer pump or temperature rise in gas generator may be a problem in applying the preliminary stage. To solve this problem, we analyzed the effect of the slope of oxidizer pump's head curve and the oxidizer mass flow rate to combustion chamber during preliminary stage using the rocket engine startup analysis code. A moderate slope(${\circleddash}{\sim}$-3) of head curve and 80% mass flow rate during preliminary stage can reduce the oxidizer pump discharge pressure by 15 to 20% comparing with the condition of ${\circleddash}$=-4.37 head curve and 70% mass flow rate. Also it can maintain the turbine inlet temperature rise within 50K from the nominal value.

• Transactions of the Korean Society of Automotive Engineers
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• v.8 no.5
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• pp.36-46
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• 2000

The performance improvement and emission reduction in a turbocharged D.I. diesel engine was studied experimentally in this paper. The system of intake port, fuel injection and turbochager are very important factors which have influence on the engine performance and exhaust emission because the properties in the injected fuel depend on the combustion characteristics. Through these experiments it can be expected to meet performance and emission by optimizing the main parameters; the swirl ratio of intake port, fuel injection system and turbocharger. The swirl ratio of intake port was modified by hand-working and measured by impulse swirl meter. Through this steady flow test, we knew that the increase of swirl ratio is decreasing the mean flow coefficient, whereas the gulf factor is increasing. And the optimum results of engine performance and emission are as follows; the swirl ratio is 2.43, injection timing is BTDC 13。 CA, compression ratio is 16, combustion bowl is re-entrant 5$^{\circ}$, nozzle hole diameter is $\Phi$0.28＊6, turbocharger is GT40 model which are compressor A/R 0.58 AND turbine A/R 1.19.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.5
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• pp.109-115
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• 2004

A 1-D system design program has been developed for the preliminary design of the turbopump system in liquid rocket engines, which use LOx and kerosene as propellants. Gasgenerator cycle and staged combustion cycle were considered as turbopump type liquid rocket engine systems. In the system analysis, mass flow balance, thrust, specific impulse, mixture ratios, turbopump power, and turbine expansion ratio of engine system were analyzed. Results show that most of the parameters agree well with real engine parameters except gasgenerator. Therefore, the l-D system design program developed in this study can be used to derive the preliminary design parameters of a turbopump with any thrust level liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.109-114
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• 2008

This paper presents the development status of a subscale precooled turbojet engine "S-engine" for the hypersonic cruiser and space place. S-engine employs the precooled-cycle using liquid hydrogen as fuel and coolant. It has $23cm{\times}23cm$ of rectangular cross section, 2.6 m of the overall length and about 100 kg of the target weight employing composite materials for a variable-geometry rectangular air-intake and nozzle. The design thrust and specific impulse at sea-level-static(SLS) are 1.2 kN and 2,000 sec respectively. After the system design and component tests, a prototype engine made of metal was manufactured and provided for the system firing test using gaseous hydrogen in March 2007. The core engine performance could be verified in this test. The second firing test using liquid hydrogen was conducted in October 2007. The engine, fuel supplying system and control system for the next flight test were used in this test. We verified the engine start-up sequence, compressor-turbine matching and performance of system and components. A flight test of S-engine is to be conducted by the Balloon-based Operation Vehicle(BOV) at Taiki town in Hokkaido in October 2008. The vehicle is about 5 m in length, 0.55 m in diameter and 500 kg in weight. The vehicle is dropped from an altitude of 40 km by a high-altitude observation balloon. After 40 second free-fall, the vehicle pulls up and S-engine operates for 60 seconds up to Mach 2. High altitude tests of the engine components corresponding to the BOV flight condition are also conducted.