• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.11a
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• pp.21-21
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• 1997

인공위성용 추진제 탱크를 개발하기 위해 여러 설계인자를 설정하여 각 인자가 탱크벽면에 미치는 응력분포 영향을 구하고, 또한 최적의 인자 값을 구하기 위해 각 인자의 변화에 따라서 구조해석을 수행하였다. 탱크 지지부 위치와 탱크 벽면 두께 변화에 따른 탱크 벽면에 미치는 응력분포 영향을 고찰하기 위해 1/4 모델을 설정하였고, 연료배출구의 위치변화(경사각돈)에 따른 응력분포는 1/2 모델을 설정하여 해석을 하였다. 탱크에 작용하는 하중은 연료압력에 의해 발생하는 정하중(350 psi)을 가하며 또한, 발사 시 발사체로부터 전달되는 최대동하중(llg)을 고려하였다. 그리고, 탱크가 인공위성에 장착될 때에 발생하는 다양한 장착조건에 대해서 구조해석을 수행하였고, 추진제 배출구 각도가 $0^{\circ}$ 에서 $25^{\circ}C$까지 변화할 때 탱크 벽면에 미치는 응력분포 영향을 구했다. 그래서 각 조건에서 구한 상당응력분포와 인자의 최적 값은 추진제 탱크를 설계하기 위한 기초적인 자료로 활용하고자 한다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1171-1175
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• 2017

Cryogenic helium gas is used as the pressurant for the oxidizer pressurization of DR(Damper Receiver) sphere in the PSD(Pogo Suppression Device) system and liquid oxygen is used as the oxidizer for the propellant in Korea Space Launch Vehicle-II. The helium gas is stored in pressurant cylinders inside the cryogenic liquid oxygen tank and liquid oxygen is stored in the oxidizer tank. In this study, the performance test of PSD liquid oxygen drain valve for KSLV-II was considered.

• Journal of Drive and Control
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• v.18 no.2
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• pp.32-37
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• 2021

The piston reservoir is mainly used in hydraulic blow-down system for aerospace engineering. The reservoir is heavy due to both hydraulic cylinder and piston in pressurization. The positive expulsion tank with rubber diaphragm has been mostly applied propellant and fuel tank at low pressure to satellites. To reduce weight, the reservoir that can be used at high pressure with rubber diaphragm was developed. In this research, the prediction of life-time for the rubber diaphragm was implemented through an accelerated life test, as a part of development of new reservoir. Also, the diaphragm was stored in an temperature chamber at the same condition as and operation with hydraulic oil. As a result, the life-time for a rubber diaphragm was successfully evaluated via Arrhenius law and Time-Temperature Superposition based on failure times over temperatures in the accelerated test.

• Composites Research
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• v.36 no.2
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• pp.132-139
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• 2023

A linerless composite propellant tank was designed and manufactured by using the carbon fiber-reinforced composite materials which have superior strength-to-weight ratio in order to reduce weight of the tank. In this research, we designed a sub-scale composite propellant tank with a diameter of 800 mm to withstand an MEOP of 1.7 MPa. We manufactured the boss of the tank by using the same composite materials to reduce the thermal expansion difference between the boss and the secondary-bonded composite layers of the barrel in the cryogenic environment. We used the collapsible mandrel to manufacture the tank without any liner. The mandrel was made from epoxy-based composite tooling prepregs to reduce weight of the mandrel. We manufactured the test tanks by laying up the carbon fiber fabric prepregs manually on the mandrel and then applying the autoclave cure process. We performed a proof test, a helium tightness test, a repeated pressurization test, and a burst test in room temperature. The test results demonstrate that the proposed design and manufacture process satisfies all strength requirements as well as an anti-leakage requirement.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.7
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• pp.592-600
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• 2007

Nowadays, researches for replacing material systems for cryotanks by composites have been being performed for the purpose of lightweight launch vehicle. In this paper, a type 3 propellant tank, which is composed of the composite developed for cryogenic use and an aluminum liner, was fabricated and tested considering actual operating environment, that is, cryogenic temperature and pressure. For this aim, liquid nitrogen (LN2) was injected into the fabricated tank and in turn, gaseous nitrogen (GN2) was used for pressurization. During this test procedure, strains and temperatures on the tank surface were measured. The delamination between hoop layer and helical one, was detected during the experiment. Several attempts were followed to investigate the cause analytically and experimentally. Thermo-elastic analysis in consideration of the progressive failure was done to evaluate the failure index. Experimental approach through a LN2 immersion test of composite/aluminum ring specimens suitable for simulating the Type 3 tank structure.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.25-37
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• 2005

For pump-fed rocket propulsion system, the temperature of LOX to be supplied to turbopump inlet should be satisfied with pump inlet temperature requirement during all operating stages, as excessive temperatures can result in cavitation due to reduction in NPSH, thus either damaging the pump or adversely affecting pump performance rise. So exact estimation of LOX temperature rise is absolutely needed for developing reliable propulsion system. This paper presents systematic analysis scheme for estimating inner process of cryogenic propellant tank which is needed for LOX temperature rise. And this paper presents LOX temperature rise and thermal stratification for all rocket operating stages including cooling, filling, waiting, pre-pressurization and firing, with the application of buoyancy driven boundary layer theory.

• Aerospace Engineering and Technology
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• v.2 no.1
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• pp.117-123
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• 2003

This paper introduces ground feeding facility for flight test of sounding rocket(KSR-III) which use liquid propellants and addresses facility configuration, development process and results. Supply of propellants and pressurization gases to vehicle according to predefined launch scenario is the primary goal of ground feeding facility. It was constructed at KSR-III launch site, verified by several tests and used for the flight test successfully.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.11
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• pp.1028-1034
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• 2015

Solenoid valves for space launch vehicles require the strict limitations on the size, weight and current consumption comparing to industrial solenoid valves. The preliminary design of a cryogenic and high pressure solenoid valve for propellant tank pressurization which can ensure the operation of solenoid valve under such strict limitation conditions was preformed. The Copper and Constantan materials in coil design have used to prevent the excessive rise of the current at cryogenic state. The measured current of solenoid valve at cryogenic temperature satisfies a design requirement.

• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.26-32
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• 2020

In this study, a facility was constructed to supply liquid nitrogen to simulate combustion instability in a liquid rocket combustor. The pressurization and supply performances were predicted and verified through different experiments. The liquid nitrogen supply system was composed of a pressurized supply system, and a dome regulator was used to adjust the pressure of the pressurant. A cavitation venturi was used to control the mass flow rate of liquid nitrogen. The condition of liquid nitrogen supply was a mass flow rate of 2.55 kg/s and the venturi inlet pressure was above 100 bar. Based on the initial experiment, it was observed that the predicted amount of the pressurant was not sufficiently supplied and the target pressure was not supplied due to a drop in tank pressure. Through the modification of the established facilities, the target mass flow rate was successfully supplied and the cryogenic liquid nitrogen supply facility was verified.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.259-265
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• 2005

To measure burning rate of solid propellants using ultrasonic technique, a special closed bomb and an ultrasonic and pressure measurement system are fabricated. During pressurization tests and burning tests on propellants, ultrasonic and pressure signal are acquired in realtime fashion by this system. Based on acquired signals, analysis programs using two different algorithm which can measure burning rates corresponding to pressures are compared. One algorithm is to correct sound velocity variation of propellants and solid couplant, another one is only to correct sound velocity variation of propellants. And accuracies of homing rates measured through these algorithms are calculated through comparison with homing rates measured using strand burner method.