• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.4
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• pp.92-97
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• 2019

A study was conducted to improve the adhesion of propellant, liner,and insulation. Insulation was shown to be more advantageous in improving the adhesion when a barrier coat was applied compared to a bare insulation layer. It was confirmed that the adhesion strength between the insulation and the propellant improves as the thickness of the liner coating increases. The liner was cured for 24 h. If the liner is cured for a long time, it will adversely affect adhesion. Adhesion is also improved when a bonding agent is applied. As the bonding agent content increases, the adhesion improves. There is a change in the adhesive strength depending on the type of bonding agent used. HX-868 shows slightly more improved adhesion than HX-752.

• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.551-554
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• 2008

In this study, numerical methods for the code development of the interior ballistics have been conducted. Mathematical models and numerical methods for the analysis technique of the granular solid propellants have been investigated. As the results of applying the methods of errors have been generated by calculation for the specific surface area of the granular solid propellants. To remove these error, the developed Eulerian-Larangian method for multiphase flows has been suggested.

• International Journal of Aerospace System Engineering
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• v.3 no.2
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• pp.9-13
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• 2016

This study is about technical analysis in launching SLBM of North Korea. We expect that North Korea develop ICBM and SLBM by improving the technique called R-27. Also it is expected that they attempt to achievement in covertness and ambush by completing technique of cold launching. Recently, SLBM of North Korea rised 40 ~ 50 m on surface after launching in an underwater when they showed the scene of firing SLBM. We expect that they actively use not general technique but super cavitation technique. Also, they might improve the launching technique by doing SLBM launching test. This type is about that whole rocket is separated two parts and ignited with high velocity and we might think that 1st rocket is used in solid propellant to maneuver in high velocity in an underwater. After then, they might use liquid propellant for the long-range ballistic missile.

• Journal of the Korean Society of Combustion
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• v.5 no.1
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• pp.43-53
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• 2000

Experimental studies on determination of the supply leading time of propellants to combustion chamber have been made to stably and efficiently guarantee the ignition process with liquid rocket engine. The propellant used is a Jet A-1 as fuel and a liquid oxygen as oxidizer. Unlike impinging FOOF type of injectors are arranged radially and the designed O/F ratio is 2.34. The present experiment program also includes the stability on the quadlet type of ignitor using the triethylalumimum as an ignition source and injector life tests. Experimental results clarifies that the propellant supply through LOx leading to combustion chamber is proper for stable ignition and combustion processes based on the fuel and oxidizer manifold pressures, combustion chamber pressure, and the variation of flame length from the nozzle exit with lapse time, and shows that the leading supply time of propellants affects the engine performance little. The effect of positioning cooling holes is remarkable to protect the injector face.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.368-373
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• 2017

Control method for DACS with variable burning area is designed and the performance of the control method is analyzed by doing simulation at various conditions. DACS, which got solid propellant on board, is designed as end-burning type typically. End-burning type DACS has the merit of controlling pressure and thrust, but it discharges the combustion gas which does not using for getting thrust. Therefore, optimal design of propellant grain and burning area changes over time as a result. Variable burning area can be assumed as a disturbance and adaptive control method is useful for pressure control of DACS effected by disturbance.

• Membrane Journal
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• v.22 no.4
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• pp.235-242
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• 2012

Ammonium perchlorate (AP) is primarily derived from the process of liquid incineration treatment when dismantling a solid rocket propellant. A series of batch dead-end nanofiltration (NF) and reverse osmosis (RO) membrane experiments were conducted to explore the retention mechanisms of AP under various hydrodynamic and solution conditions. Low levels of silicate type of siloxane had been detected through the GC/MS and FTIR analysis of liquid solutions extracted from solid ammonium perchlorate composite propellant (APCP). It is indicated that NF/RO membranes fouling in the presence of APCP was mainly attributed to the AP interactions because the concentration of silicate type of siloxane was negligible compared to that of AP. The osmotic pressure of AP was presumably resulted in the flux declines ranging from 13 to 17% in the case of the application of low-pressure (551 and 896 kPa for NF and RO) compared to those in application of high-pressure. The retention of AP by NF/RO membranes significantly varied from approximately 10 to 70% for NF and 26 to 87% for RO, depending on the operating and solution water chemistry conditions. The results suggested that retention efficiency of AP was fairly increased by reducing concentration polarization (i.e. application of low-pressure and stirring speed of 600 rpm) and increasing the pH of a solution. The result of this study was also consistent with the previous modeling of 'solute mass transfer of NF/RO membranes' and demonstrated that hydrodynamic and solution water chemistry conditions are to be a key factor in the retention of AP by NF/RO membranes.

• 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 Mechanical Science and Technology
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• v.14 no.4
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• pp.466-476
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• 2000

This paper presents atomization characteristics of a double impinging F -0-0- F type injector with four streams. A phase Doppler particle analyzer was employed to measure the droplet-size and water was used as the inert simulant liquid instead of reactive propellant liquids. The droplet mean diameter (SMD) and size distribution were measured to investigate the effects of the momentum ratio and pressure drop variations. This experimental results can be used during the preliminary design stage of a impinging stream type injector for liquid rockets.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.622-627
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• 2017

In this study, the numerical analysis of propellant supply manifold of a sub-scale model combustor with the same combustion and acoustic resonance conditions as the real combustor was carried out. The analysis of the results is based on the shape of the manifold and the number and type of inflow channels. The manifold form of the main propellant was rounded so that the recirculation area was small and easy to manufacture. The liquid oxygen mainfold included a distributor to uniformly supply the flow, and since the kerosene manifold was directly in contact with the flame side, the inflow channel was formed at a radial $360^{\circ}$ so as to minimize the recirculation region. The liquid nitrogen manifold was configured as a radial inflow channel to reduce the velocity difference near the injector.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.375-383
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• 2005

The present paper describes the first development of a LOX/kerosene type turbopump in Korea. The liquid rocket engine, that the turbopump can be applied to, has a 30-ton(metric) level of vacuum thrust and employs a gas generator cycle. The turbopump consists of two single-stage centrifugal pumps, that is, LOX and kerosene pumps, and one single-stage impulse turbine. Inter-propellant seal(IPS) is located between the LOX pump and the kerosene pump to avoid any interaction between the propellants. A series of component and TPU(Turbopump Unit) test has been completed in the level of simulant propellants and ready for hot firing tests.