• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.4
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• pp.40-49
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• 2001

The basic research on AKM(Apogee Kick Motor) for space launch vehicle was carried out. AKM which will be used as 3rd stage solid rocket motor in 3-stage Korean Sounding Rocket(III) has been developing. KM is a solid rocket motor using composite propellant based on HTPB and is composed of composite motor case and submerged nozzle. To develop KM rocket motor satisfing a given set of requirement, firstly the full-scale KM with diameter 520mm was designed, then sub-scale motors reduced about 60% were manufactured and tested. Full-scale ground firing test is accomplished two times.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.5
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• pp.62-69
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• 2014

Infrared signature of rocket plume plays an important role for detection, recognition, tracking and minimzing for low observability. Infrared signatures of rocket plume with reduced smoke propellant and smokeless propellant are measured. In order to estimate the infrared signature of rocket plume, CFD analysis for flow structure of plume is performed, and layered integration method for estimating of infrared signature is used. Numerical and experimental results were in good agreement. Both propellants had similar infrared signature. Strong peak at $4.3{\mu}m$ region in the experimental results is appeared due to experimental error arising from the calibration procedure.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.7-10
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• 2009

The development and evaluation of solid propellant were performed for the turbopump pyro starter, which start up the liquid propellant rocket engine for the Space Launch Vehicle (SLV). Requirements for the turbopump pyro starter propellant include the production of low flame temperature, low burning rate and nontoxic gas to protect the mechanical corrosion or air pollution. This study describes the development of the solid propellant composition which is based on PCP binder. DHG (Dihydroxy glyoxime), which has advantages of oxygen balance and ignition, was used as coolant. The mechanical properties and burning rate of the propellants were measured. Finally, static fired test was performed to prove the possibility of development.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.383-387
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• 2011

The supercavitating rocket torpedo SHKVAL was analyzed in view of its system operation procedure and the structure and performance. 3 different propulsion systems installed in SHKVAL were 1st solid rocket booster for launch and acceleration, 2nd solid rocket booster for further acceleration, and Mg-rich Hydroreactive fuel rocket propulsion system for cruising. The gas generator used to help generate the supercavitation bubble was composed of a solid propellant gas generator and a hydroreactive fuel one. The structures and their performance were described based on as much knowledge as we have obtained from cumulative information and up-to-date analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.475-478
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• 2005

Ducted rocket produces thrust by 2 steps, primary incomplete combustion in the gas generator, and secondary complete combustion reaction in combustion chamber mixed by air taken through duct. the range of a rocket is determined by the weight of propellant, especially the weight of fuel. So ducted rocket has more efficiency and high terminal speed compared to traditional solid rocket motor. This propulsion system expected to be applied to various kinds of missile for anti-aircraft, anti-ship

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.1
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• pp.15-23
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• 2019

In this study, an extinguishment model of a three-dimensional solid propellant rocket was developed by combustion and fast depressurization to control the thrust of a solid rocket. Computational fluid dynamics simulation was carried out to ascertain the change in flow patterns in the combustion chamber and the extinguishment process by using a pintle. An ammonium perchloride was used as the target propellant and the dynamic behavior of its major parameters such as temperature, pressure, and burning rate was predicted using the combustion model. The dynamic behavior of the combustion chamber was confirmed by fast depressurization from an initial pressure of 7 MPa to a final pressure of 2.5 MPa at a depressurization rate of approximately -912 MPa/s.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.225-228
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• 2007

The objective of this study is to characterize design parameters of rocket igniters for composite, double base and nitramine propellant. Arc image furnace and fiber optics surface reflectometer were used to measure ignition delay time and reflected optical energy of several compositions of composite, double base and nitramine base rocket propellant at different pressure levels each other. The order of ignitability was double base > composite > Nitramine propellants at initial pressure of over 75 psia. The highest ignition energy was needed to ignite nitramine propellant, however, as the pressure increased up to the range of $75{\sim}400$ psia as the ignition delay time decreased abruptly. The absorbtion of radiation energy could be increased by the addition of small amount of opacifiers as carbon black, ZrC, WC and burning catalyst.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.4
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• pp.44-53
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• 2022

The depressurization rate in a thrust variable solid rocket motor is the major factor that has the greatest influence on the thrust termination performance. In this study, the depressurization rates range of model solid rocket motor was identified and major factors affecting the depressurization rate were found. It is important for actual system design to understand the depressurization rate of the system that can satisfy the target performance as well as the extinguishing characteristics of the solid propellant. The methodology for obtaining the depressurization rate model in this study is considered to be applicable to the optimal design of the thrust terminable propulsion system.

• Transactions of the Korean Society of Mechanical Engineers
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• v.17 no.9
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• pp.2280-2285
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• 1993

Many items, as internal pressure, thrust, temperature, strain, etc. are measured in Ground Firing Test (GFT) of rocket motors. But these items are influenced by various phenomena occurred during propellant combustion. In this study, natural frequencies of motor itself and system(motor+loadcell) on Stand were measured. Also motor responses were measured during burning and analyzed so that the vibration characteristics occurred during GFT and the causes and characteristics of vibration signal appearing on thrust curve were identified.

• 제어로봇시스템학회:학술대회논문집
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• 1991.10a
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• pp.413-418
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• 1991

Two methods are applied to the problem of trajectory optimization for launch vehicles which burn solid propellant. One is 'Optimal Control' theory, the other is 'NonLinear Programming' method. Trajectory optimization for solid rocket motors has a special problem. The special problem is that the payload of launch vehicle is not the function of control variable. This paper deals with this special problem.