• Journal of Aerospace System Engineering
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• v.18 no.3
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• pp.70-75
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• 2024

The insulation performance of a common bulkhead propellant tank for small launch vehicles with variations in insulation thickness was analyzed. The common bulkhead propellant tank composed of a single part allows for lightweight design, as it eliminates the need for tank connections. However, problems such as propellant loss and ignition delay due to heat transfer caused by temperature differences between oxidizer and fuel may arise. Therefore, it is essential to verify the insulation performance of the common bulkhead structure that separates the oxidizer tank and fuel tank. In this study, transient heat transfer analysis was conducted for propellant tanks with insulation thicknesses of (50, 55, 60, 65, and 70) mm to analyze the insulation performance using boil-off mass. Subsequently, the boil-off mass of the oxidizer generated during the first-stage flight time of the propellant tank was determined. The results confirmed that increasing the insulation thickness reduces the boil-off mass, thereby improving the insulation performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.4
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• pp.50-58
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• 2016

Korean Space Launch Vehicle (KSLV-II) Propellant Supply System charges liquid oxygen and kerosene to each propellant tank for the stages. To charge the launch vehicle propellant tank safety, the propellant charge flow rates and scenarios should be defined. First, the Propellant Supply System was modeled with 1D flow analysis program. The control valve capacity and orifice size were calculated by performing the 1D steady state simulation. Second, the 1D transient simulation was performed by using the steady state simulation results. As propellants were being charged at the each tank, the increased tank liquid level decreases the charge flow rate. Consequently, the proposed supply system satisfies the required design charging conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.370-373
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• 2007

In this study, a hybrid rocket motor with separable and detachable oxidizer tank from combustion chamber is developed. Initially, the measured thrust of the motor showed about 30% of the design thrust since the oxidizer supply was not enough. In order to solve this problem, application is made to expand the orifice diameter of oxidizer injector empirically, so that the mass flow rate of oxidizer was improved. The improved performance was about 60% of design thrust, 18kgf, and thrust-to-weight ratio was reasonable, compared with other sounding rockets.

• Bulletin of the Korean Space Science Society
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• 2003.10a
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• pp.90-90
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• 2003

한국항공우주연구원은 액체추진기관 시스템을 이용한 3단형과학로켓(이하 KSR-III)을 국내 최초로 개발하여 비행시험을 수행하였다. 액체추진기관 로켓의 비행시험을 위해서는 이전의 고체 추진기관을 이용한 과학로켓 1, 2와는 달리 비행시험 조건에 부합하게 액체추진제 및 가압제 등을 공급하는 지상설비가 필요하다. 이에 한국항공우주연구원은 독자적으로 비행시험에 필요한 제반 설비를 갖춘 발사장을 구축하였다. KSR-III는 압축 헬륨가스(GHe)를 이용하여 연료(Jet A-1)와 산화제(LOx)를 가압하여 추력을 얻는 액체추진기관 시스템이다. 따라서 발사장에서의 지상공급설비는 유공압 설비와 발사시나리오에 따라 해당 부품을 제어하고 자료를 저장하는 제어/계측 설비 및 기타설비들로 구성되어 있다. 지상공급설비 중 유공압 설비는 LOx의 저장 및 기체 내 산화제 탱크의 충전을 위한 산화제 공급설비, Jet A-1의 저장 및 기체 내 연료 탱크의 충전을 위한 연료 공급 설비, 지상설비용 밸브구동 및 기체 내부 퍼지 등에 필요한 질소($N_2$)를 저장/공급하는 설비, 기체내부 밸브 구동 및 가압제로 사용되는 기체헬륨(He)을 저장/공급하는 설비들로 구성되어 있다. 이러한 구축된 공급설비는 기능시험, 연계시험 등의 각종 입증시험을 통해 그 성능을 검증한 후 단인증모델(SQTM)을 이용하여 발사 시나리오에 따른 추진제 공급능력을 입증한 후 KSR-III의 비행시험을 성공적으로 수행하였다. 수행된 연구결과는 향후 건설되어질 우주센터내의 발사장 기반설비 설계의 기초 자료로 활용할 수 있을 것이다.

• Aerospace Engineering and Technology
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• v.1 no.1
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• pp.55-64
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• 2002

A conceptual design of propulsion system for a geosynchronous communication satellite with 12 years design life is presented in this paper. Propellant mass budget for the design life is calculated using total velocity increment (ΔV) flowed-down from mission requirement analysis. Sizes of the fuel and oxidizer tank are derived based on the calculated propellant mass budget, and mass of the pressurant as well as the size and pressure of pressurant tank are calculated too. Thruster positioning, number of rocket engines, and position of tank are determined through Trade-Off Study with Structure & Mechanical Subsystem. Propulsion system configuration and its schematics are presented finally.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.50-54
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• 2011

This paper presents a two-phase model for hybrid rocket internal ballistics design using $N_2O$ as oxidizer The two-phase model results are compared with data obtained from static firing test. Two-phase model is suitable for blow-down type with saturated compressible fluid as $N_2O$, presented the result by Part 1. HDPE as Fuel, and $N_2O$ as oxidizer were used during the static firing test. The combustor were designed for an average thrust of 30 kgf where oxidizer tank pressure in set to 50 bar. The numerical results of internal ballistic showed good agreements with static firing test results where thrust, oxidizer tank pressure and chamber pressure are compared.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.4
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• pp.33-41
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• 2008

The prelaunch thermal transients in the cryogenic oxidizer tank of liquid propulsion rocket subjected to uniform heat flux from outside are numerically analyzed through thermodynamic equations and heat and mass transfer relations. The prelaunch stage is assumed to be composed of five idealized sub-stages including pressurization process by helium gas injection. The Peng-Robinson equation of state is utilized in the lumped analysis of ullage gas. The liquid region is divided into a number of horizontal layers of uniform properties to account for the thermal stratification. The computational result for the typical case shows that the temperature rise of liquid oxidizer is less than 1K and the adsorbed helium into the liquid is approximately 10g.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.363-367
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• 2009

For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.63-68
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• 2011

For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.