• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.413-418
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• 2008

Flow instability in the rocket turbo pump system can be caused by various reasons such as valve, orifice and venturi, etc. The inception of cavitation, especially in the propellant feeding system, is the primary cause of the mass flow and pressure oscillation due to cyclic formation and depletion of cavitation. Meanwhile, the main propellant in liquid rocket engine is the cryogenic one, which is very sensitive to temperature variation, and the variation of propellant properties caused by thermodynamic effect should be accounted for in the flow analysis. The present study focuses on the formation of cryogenic cavitations by adopting IDM model suggested by Shyy and coworkers. Also, the flow instability was investigated in the downstream of orifice by using a developed numerical code. Calculation results show that cryogenic cavitations can lead to flow instability resulting in mass flow fluctuations due to pressure oscillations. And the prediction of cavitations in cryogenic fluid is of vital importance in designing feeding system of LRE.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.3
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• pp.109-118
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• 2018

Korea Aerospace Research Institute has developed a staged combustion cycle rocket (SCCR) engine with high specific impulse to send a 3-ton class satellite into geostationary orbit while conducting a Korean Space Launch Vehicle (KSLV) II project. The SCCR engine is different from the KSLV-II engine, which is an open cycle engine using a gas-generator. The SCCR engine with a closed cycle engine is composed of a pre-burner, a turbo pump, and a main combustor. The technology demonstration model (TDM0) was assembled and tested in the 7ton-class engine combustion test facility of Naro Space Center, and the combustion test was successfully conducted.

• Proceedings of the Korean Vacuum Society Conference
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• 2012.08a
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• pp.265-266
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• 2012

의료용 X-ray의 발전에 따라, 영상의 Digital화가 필요하게 되었다. Digital 영상 구현을 위해 다양한 형태의 영상 검출기가 개발되었다. 진단 영상의 조건으로는 구현 시간이 빠르고 해상도가 높아야 한다. 조건에 부합하는 Flat panel 형태의 직접방식과 간접방식 검출기의 개발이 주로 이루어졌으며, X-ray 검출 효율이 높고 공간 분해능이 높은 직접 방식의 검출기에 대한 연구가 활발히 진행되고 있다. 기존 직접방식의 X-ray 검출물질로는 A-Se이 이용되었다. 하지만 A-Se의 경우 낮은 원자번호로 인해 X-ray에 대한효율이 낮으며, 제조 공정과 수율의 문제로 인해 대체 물질의 개발과 공정의 개선이 필요하다. 선행 연구를 통해 X-ray 검출물질의 전기적 특성을 파악을 통해 대체 물질로서 가능성을 알아보았다. 본 연구에서는 기존에 제작된 X-ray 검출물질의 상부전극 증착 물질과 증착법 선정에 대한 연구이다. 선행 연구를 통해 선정된 X-ray 검출물질은 HgI2이다. 상, 하부 전극 선택에 있어 HgI2의 일함수 값(4.15eV)을 고려하여 그와 비슷한 일함수 값을 가진 물질로 전기적 장벽을 제거하여야 한다. 따라서, ITO (일함수 4.45eV)와 Au (일함수 5.1eV)을 선택하였다. ITO의 증착으로 이용된 방법으로는 on-axis 형태의 magnetron plasma sputtering을 이용하였으며, Au의 증착으로 이용된 방법은 Thermal evaporation deposition을 이용하였다. plasma sputtering에 이용된 타겟은 In2O3;SnO2 (조성비:90:10wt%)를 사용하였으며, Chamber의 크기는 넓이 456 ${\phi}cm^2$ 높이 25 cm이며, 로 target과 기판과의 거리는 15cm이다. plasma발생에 필요한 가스로는 Ar과 O2를 이용하였다. 고 진공 환경 조성에 이용된 장비로는 Rotary pump와 Turbo molecular pump이다. plasma 발생 전 진공도는 $3.2{\times}10^{-5}$ Torr, 발생 후 진공도는 $5.1{\times}10^{-5}$ Torr이다. plasma 환경이 조성된 후 증착 시간은 1분 30초이다. Au는 순도 99.999%를 이용하였으며, 이용된 금은 1회 증착에 0.3 g을 이용하였다. Chamber의 넓이 1,444 ${\phi}cm^2$이며, 높이 40 cm, boat와 기판과의 거리는 25 cm이다. 고 진공 환경 조성에 이용된 장비로는 Rotary pump와 diffusion pump를 이용하였다. Au의 승화 전 진공도는 $2.4{\times}10^{-5}$ Torr 증착 시 진공도는 $4.2{\times}10^{-5}$ Torr이며, Boat에 가해준 전압, 전류는 0.97 V, 47 A이며, 증착 시간은 1분 30초이다. 광도전체 층에 각각 증착된 전극의 저항을 통해 증착상태를 판단하였다. DMM (Digital Multimeter)로 1 cm 간격으로 측정된 표면의 저항은 ITO 약 $8{\Omega}$, Au 약 $3{\Omega}$으로 전극으로서 이용이 가능한 상태이다. Au와 ITO가 증착된 HgI2 시편의 전기적 특성은 기존에 이용된 X-ray 변환물질의 성능보다 우수하였다. 하지만 Au와 ITO가 각각 증착된 시편의 전기적 특성은 큰 차이를 보이지 않았다. ITO의 경우 진공 상태에서 이용되는 Gas가 이용되며, Plasma 환경 조성 유지가 어려운 점이 있다. Au전극은 증착 환경 조성이 쉽지만, 전극 물질 이용효율이 떨어지는 단점이 있다. 본 연구를 통해 X-ray 변환물질인 HgI2의 전극물질로 Au와 ITO의 이용가능성을 알아보았다. 두 전극으로 제작된 검출기의 성능은 큰 차이 없이 우수하였고, 전기적 장벽 상태가 낮아 높은 검출 효율을 보였다. 상대적으로 Au 전극의 공정이 간단하고 수율이 높다. 하지만 Au Source의 이용 효율이 떨어지는 단점이 있다. 본 연구의 결과를 통해 공정상의 유리함과 Source의 이용효율을 고려한 분석에 대한 연구가 필요할 것으로 사료된다.

• International Journal of Fluid Machinery and Systems
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• v.4 no.3
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• pp.334-340
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• 2011

Turbo-pumps have weak points, such as the pumping operation is unstable on the positive slope of the head curve and/or the cavitation occurs at the low suction head. To improve simultaneously both weak points, the first author invented the unique pumping unit composed of the tandem impellers and the peculiar motor with the double rotational armatures. The front and the rear impellers are driven by the inner and the outer armatures of the motor, respectively. Both impeller speeds are automatically and smartly adjusted in response to the pumping discharge, while the rotational torques between both impellers/armatures are counter-balanced. Such speeds contribute to suppress successfully not only the unstable operation at the low discharge but also the cavitation at the high discharge, as verified with the axial flow type pumping unit in the previous paper. Continuously, this paper investigates experimentally the effects of the tandem impeller profiles on the pump performances and the rotational speeds against the discharge, using the impellers whose loads are low and/or high at the normal discharge. The worthy remarks are that (a) the unstable operation is suppressed as expected and the shut off power is scarcely large in the smart control, (b) the blade profile contributes to determine the discharge giving the maximum/minimum rotational speed where the reverse flow may incipiently appears at the front impeller inlet, (c) the tandem impeller profiles scarcely affect the rotational speeds, while the loads of the front and the rear impellers are same, but (d) the impeller with the low load must run faster and the impeller with the high load must run slower at the same discharge to take the same rotational torque, and (e) the reverse flow at the inlet and the swirling velocity component at the outlet of the front impeller with the high load require making the rotational speed of the rear impeller with low load fairly faster at the lower discharge.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.892-897
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• 2011

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility(CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The fuel supply system modeling using AMESim was performed based on the results of the detailed design, and the fuel supply characteristics was analyzed in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.6
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• pp.92-97
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• 2012

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The oxidizer supply system modeling using AMESim was performed based on the results of the detailed design, and the oxidizer supply characteristics was analyzed in this paper.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.502-506
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• 2012

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The oxidizer supply system modeling using AMESim was performed based on the results of the detailed design, and the oxidizer supply characteristics was analyzed in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.4
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• pp.87-92
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• 2012

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The fuel supply system modeling using AMESim was performed based on the results of the detailed design, and the fuel supply characteristics was analyzed in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.3
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• pp.82-89
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• 2012

The temperature of cryogenic propellant in the propellant tank increases during flight due to heat input from surroundings. The propellant which temperature rises up over the required condition of turbo-pump remains as unusable propellant at the end of flight. In this paper the estimation method of the heat transfer coefficient at the upper layer of cryogenic propellant was presented. The heat transfer mode at the propellant upper layer was considered as conduction. Temperature distributions near propellant surface obtained from heat transfer coefficient were compared with test data to show the possibility of this method.

• Aerospace Engineering and Technology
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• v.13 no.2
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• pp.160-166
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• 2014

The shutoff valve of a Liquid Rocket Engines (LRE) controls the flow of propellant between turbo-pump and combustion devices of LRE using pilot pressure and spring force. The shutoff valve is closed when the pilot pressure is removed from the diaphragm type actuator. During designing process of life cycle is when should be analyzed according to the characteristics of forces with respect to the opening and closing of diaphragm actuator. A valve has been designed to adjust the control pressure which is required to open a poppet and to determine the working fluid pressure at which a valve starts to close. During flow capacity test under room temperature as a part of life cycle tests, the leakage in diaphragm was occurred due to the leakage of sheet welding sections. The operating cycle of the diaphragm type actuator is about 61 times with 22 MPa of pilot pressure.