• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.97-100
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• 2008

A hybrid sounding rocket carrying about 10kg payload reaching up to 15km altitude has been designed. The commercial seamless aluminium tube and liquid ${N_2}O$ without pressurization devices were chosen as rocket motor case and oxidizer supply system respectively. A hybrid rocket engine performing required propulsion impulse is designed with time dependent internal ballistic scheme. Engine performance, aerodynamic characteristics, and trajectory were predicted by a integral technique of internal ballistics and external ballistics. The design results were evaluated by comparison with previous experimental data, technical reports, and literatures.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.144-147
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• 2017

A hybrid rocket engine capable of thrust throttling and underwater-working was developed for the underwater high-speed vehicle propulsion system. The hybrid rocket engine was designed and made by two types of ground test motor and underwater working motors. An engine performance was verified by the ground tests with the ground test motor and in the case of underwater motors the ground tests and underwater tests were performed. For the underwater operation a two-stage ignition system was adopted and a rupture disc was installed at the end of nozzle for a water-tight just before an ignition. Successful ignition and propulsion were confirmed in the underwater test with the final selected double rupture disc.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.251-254
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• 2008

Ground firing test facility and test field for firing test of hybrid rocket engine were constructed. Ground firing test facility were composed of hybrid rocket engine, thrust stand, oxidizer storage/supply system, control system and data acquisition system. Firing tests of thrust 50 kgf class were conducted. Stable performance data was obtained and operational reliability of ground firing test facility were found.

• International Journal of Aerospace System Engineering
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• v.2 no.2
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• pp.29-33
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• 2015

The multidisciplinary design optimization (MDO) of a launch vehicle (LV) with a hybrid rocket engine (HRE) was carried out to investigate the ability of an HRE for a single-stage LV. The non-dominated sorting genetic algorithm-II (NSGA-II) was employed to solve two design problems. The design problems were formulated as two-objective cases involving maximization of the downrange distance over the target flight altitude and minimization of the gross weight, for two target altitudes: 50.0 km and 100.0 km. Each objective function was empirically estimated. Several non-dominated solutions were obtained using the NSGA-II for each design problem, and in each case, a trade-off was observed between the two objective functions. The results for the two design problem indicate that economical performance of the LV is limited with the HRE in terms of the maximum downrange distances achievable. The LV geometries determined from the non-dominated solutions were examined.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.3
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• pp.112-118
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• 2019

The development of an underwater rocket propulsion system was described in this paper. Throttle able liquid propellant and hybrid rocket propulsion systems were selected for underwater propulsion. A subscale liquid rocket combustion chamber and it's portable stand were developed and their applicability was examined. 1.5-ton.f and 1.8-ton.f hybrid rockets were developed for underwater applications. The test results indicated that the 18-ton.f hybrid rocket fully complies to the performance and underwater cruise stability requirements; thus, its development was concluded to be successfully complete.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.238-238
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• 2012

Report is devoted to place and role of numerical simulation in design of rocket propulsion systems. In introduction advanced solutions in liquid propellant rocket engines design are presented. Further essence of design process described briefly. The central place of method of solution of direct problem in design process was shown. Numerical simulation for solving direct problem of fluid dynamic was used as the alternative to theoretical and experimental approaches. Main features of numerical models of processes in propulsion systems were observed. Some results of simulation and (or) design of different types of chemical propulsion system were presented also. The combined rocket engine, rocket engine with injection of after-turbine gas into supersonic part of the nozzle, solid propellant engine and hybrid propulsion engine are under consideration.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.131-134
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• 2003

There has been a renewed interest in the use of hydrogen peroxide as an oxidizer in bipropellant liquid rocket engines as well as in hybrid rocket engines. This is because hydrogen peroxide is a propellant of low toxicity and enhanced versatility. The present paper details the features of the designed engine of l00N thrust and its facility. Also explained is the arrangement of the distillation unit to be used to prepare rocket-grade hydrogen-peroxide propellant. Results of the simulated "cold" tests are presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.415-418
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• 2006

하이브리드 추진제는 연소 중 연료의 후퇴율 및 연소 면적의 변화로 인해 O/F비가 변화하므로 이로 인해 설계 및 해석에 어려움을 가지게 된다. 본 연구에서는 요구임무를 만족하는 하이브리드 로켓모터에 대한 설계를 수행하였다. 하이브리드 모터의 기초 설계를 실시하고 그 결과를 이용하여, 일정한 산화제 유량이 공급된다는 가정 하에, 연소 시 연소면적변화에 따른 O/F비 변화를 고려한 하이브리드 모터의 내탄도 해석을 실시하였으며, 그 결과를 이용하여 로켓의 외탄도 해석을 실시하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.248-251
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• 2007

The blowdown oxidizer feeding system is effective in the respect of higher reliability by the small number of parts and the absence of additional pressurization tanks, but it also has the unfavorable disadvantage such as thrust variation during the operation. Thus, in order to understand the these performance characteristics inherent in the Hybrid Rocket Motor (HRM) with blowdown oxidizer feeding system, this study proposed the integrated mathematical model to describe physical phenomena in the following parts: the oxidizer tank, combustion chamber, fuel grain, nozzle and injector.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.180-185
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• 2004

Our previous study showed that although the hybrid rocket engine with swirling gaseous oxygen had high performance, a direct injection of LOX with swirl into the combustion chamber of the hybrid rocket engine lowered the performance of the engine, compared to that with gaseous oxygen. In order to clarify this reason, combustion tests of a small PMMA combustor with an inner port diameter of 2 mm were conducted in liquid oxygen flow by comparison with gaseous oxygen flow. Although the oxygen mass fluxes of LOX were about two orders of magnitude larger than those of gaseous oxygen, the fuel regression rate of LOX were remarkably smaller than those of gaseous oxygen. For both liquid and gaseous oxygen, diffusion flames in the port of the grain controlled the combustion process of PMMA in oxygen flow. These results may be explained by the fact that only small amount of LOX vaporized and consumed in the combustion with PMMA while flowing through the port due to relatively larger latent heat of injected liquid oxygen compared to the heat of release by combustion which depended on the burning surface area of PMMA.