• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.511-514
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• 2009

In this study, the combustion characteristics of hybrid rocket fuel with tapered grain port were studied. The regression rate was increased about 17.5% by using the convergence port shape fuel. On the other hand, in case of divergence port shape fuel, any notable difference of regression rate was not observed when compared with regression rate of the cylindrical port shape fuel. Also, in case of convergence port shape fuel, characteristic velocity efficiency was increased. From these results, one can notice that convergence port shape of hybrid rocket fuel can be effective configuration in terms of improvement of combustion efficiency and performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.2
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• pp.75-84
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• 2002

A numerical analysis of unsteady motion in solid rocket motors with a nozzle has been conducted. The numerical formulation including modified $\kappa$-$\varepsilon$ turbulence model treats the complete conservation equation for the gas phase and the one-dimensional equations in the radial direction for the condensed phase. A fully coupled implicit scheme based on a dual time-stepping integration algorithm has been adopted to solve the governing equations. After obtaining a steady state solution, pulse and periodic oscillations of pressure are imposed at the head-end to simulate acoustic oscillations of a travelling-wave motion in the combustion chamber. Various steady and unsteady state features in the combustion chamber of a rocket motor has been analyzed as results of numerical calculations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.181-184
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• 2009

A development of a composite solid propellant is carried out for an application to gas generators as an energy source of rocket system. With HTPB as a propellant binder which has 80% of particle loading ratio, a favorable rheology, and moderate curing properties at the range of $-50^{\circ}C{\sim}70^{\circ}C$, AN is selected as the first kind of oxidizer having the characteristics of a low flame temperature, minimal particle residual as well as nontoxic products. AP is the second oxidant for ballistic property control. A series of experiments for the improvement of physical properties were conducted and resulted in the propellant formulation having 30% of strain rate at 8 bar of max. stress.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1994.04a
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• pp.23-26
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• 1994

130mm D.B. 추진기관의 고온 시험에서 나타난 극심한 이상 연소 현상을 해결하기 위해 미세한 고체 입자들을 연소 가스에 분산시켜 불안정 연소를 억제하는 particulate damping 효과를 연구하였다. 고체 입자로서 효과적인 것으로 알려진 $K_2$$SO_4$. ZrC, Graphite를 CTPB, HTPB 고분자 물질에 충진시켜 epoxide, isocyanate 반응기와 가교 반응을 일으킴으로써 고무상의 탄성체 성질을 갖게 하는 $K_2$$SO_4$/CTPB, ZrC/Graphite/HTPB, ZrC/Graphite/AP/HTPB, ZrC/AP/HTPB 조성의 연소 안정제를 개발하였다. 이 연소 안정제는 외경 17mm, 길이 1000mm의 안정봉 형태로 제작하여 모타의 중심 cavity에 조립한 후 지상 연소 시험을 통하여 성능을 확인하였다. 시험 결과, 조성에 AP를 포함시켜 연소 안정제에 일정한 연소 속도를 부여하여 추진제 grain 연소 동안 고체 입자를 연소 가스에 분산되게 설계한 ZrC/Graphite/AP/HTPB, ZrC/AP/HTPB 조성의 연소 안정제가 불안정 연소 억제에 효과적인 것으로 나타났다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.433-438
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• 2009

The main objective of this research effort is to develop the performance of C/SiC composites manufactured by LSI (Liquid Silicon Infiltration) method for solid and liquid rocket propulsion system and ensure the performance analysis technique. The high performance and reliability of C/SiC composite are proved for solid and liquid rocket propulsion system. And the performance analysis technique related to mathematical ablation model is originated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.189-192
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• 2010

In this study, insensitive munitions(IM) policies and technologies of advanced countries for solid rocket motor were investigated. Development trends and caseworks of each part such as propellant and motor case of rocket motor for IM were also studied. Based on these investigation and analysis for IM rocket motor, directions of the development for IM rocket motor in our country were suggested.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.7
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• pp.593-600
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• 2015

In this study, a combined hybrid rocket system is newly introduced which has characteristics of both gas generators and afterburner type hybrid rockets. In particular, a combined gas generator utilizing solid fuel and liquid/gas oxidizer was designed as a primary combustor of the system. Combustion tests were carried out with various equivalence ratio affected by parameters such as fuel length, oxidizer flow rate, fuel port diameter and fuel type. In general, fuel-rich gas generator produces low combustion gas temperature to meet the temperature requirement and the target temperature was transiently set less than 1600 K. Since it was found that controlling parameters showed limited effects on the change of equivalence ratio, mixture of $O_2$ and $N_2$ as an oxidizer was additionally introduced. As a result, a combined gas generator successfully produced combustion gas temperature of less than 1600 K Future studies will carry out more combustion tests to attain fuel-rich combustion gas temperature less than 1200 K, which was a temperature requirement of a gas generator system in the previous studies.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.233-236
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• 2009

In this study, the characteristic of the hybrid rocket motor with $LN_2O$(Liquid Nitrous oxide) was investigated experimentally. HDPE(High Density PolyEthlene) was used as fuel with different sized single port. When used $LN_2O$, combustion efficiency is lower than using $GN_2O$(Gas Nitrous oxide), because of completeness of vaporization of droplet and mixing. And regression rate was changed by different oxidizer phase. This behavior was considered that flame temperature and combustion of solid fuel front/end surface.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.466-475
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• 2011

This paper presents development status of sounding rockets containing scientific payload and telemetry at home and abroad. The case of outside, United States is launching sounding rockets in 20-30 flights per year by the NASA program which offers to carry payload weights of 38-680 kg and altitude of 88-1500 km. Europe is launching in 4-5 flights per year by the ESA program. The case of Korean sounding rockets was successful with the launch of three times(KSR-I,II,III), but Korea lags far behind the advanced countries in the field of development technologies for space launch vehicle. We expect that our scientific and industrial technologies will be improved through the research and development of sounding rockets. In this study we proposed necessity and future direction of development in domestic sounding rockets.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.3
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• pp.70-79
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• 2011

This paper presents development status of sounding rockets containing scientific payload and telemetry at home and abroad. The case of outside, United States is launching sounding rockets in 20-30 flights per year by the NASA program which offers to carry payload weights of 38-680 kg and altitude of 88-1500 km. Europe is launching in 4-5 flights per year by the ESA program. The case of Korean sounding rockets was successful with the launch of three times(KSR-I,II,III), but Korea lags far behind the advanced countries in the field of development technologies for space launch vehicle. We expect that our scientific and industrial technologies will be improved through the research and development of sounding rockets. In this study we proposed necessity and future direction of development in domestic sounding rockets.