• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.24-27
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• 2012

In this study, the full scale flight type Dual-Pulse Rocket Motor(DPRM) with the bulkhead type Pulse Separation Device(PSD) was designed, manufactured, and fire-tested. The bursting time and pressure of PSD were analyzed by the pressure, thrust and vibration results of static fire tests and ablation of PSD was measured with 3-D coordinate measuring machine. As a result, PSD requirements, bursting conditions and thermal safety, were satisfied.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.5
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• pp.12-18
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• 2014

A study of gas generator Fuel-Rich propellant for air-breathing propulsion system was performed in this paper. General solid propellant comprises a mean of 60% or more oxidizing agents. but, to develop the fuel-rich solid propellant increased the content of the metal fuel and reduced the content of the oxidizing agents by approximately 30%. Very high amount of heat per volume of fuel into the metal having the Boron was used. Amorphous Boron Powder was applied to propellant as beads type and it allowed to design more amount of metal fuel in the fuel-rich propellant. And the Combustion characteristics and properties of fuel-rich solid propellant according to the Boron-bead sizes were confirmed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.45-49
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• 2006

The thermal degradation of carbon fiber reinforced phenolic composites have been studied at high temperature by using thermogravimetry analysis (TGA). The aim is that ultimately it can be used to predict the service temperature during solid rocket firing for any level and type of mechanical loading and to recommend protection systems required. To simulate the high heating rate in firing condition, the modified thermal decomposition constant (1000 K/min) was used for FEM analysis. The temperature distribution and the thickness of thermal decomposition were estimated well and we could predict the thickness of thermal decomposition within ${\pm}1mm$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.203-206
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• 2008

In this study, we focused on development of the adhesive insulator using the case for solid rocket motors. Material of insulator used unvulcanized rubber based on EPDM/kevlar. In case of front insulator, preforms was made by using hot press molding, and then modified nylon film was inserted between two preforms for boots manufacturing. Rear insulator included cylinder part was embodied by only one mold with special designed and manufactured shape in the process. Boots part of rear insulator was obtained by cutting machine with hard-metal cemented carbide.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.271-274
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• 2007

In the fuel of the solid fuel ramjet there are metal particles in order to improve the Isp like as solid rocket propellants. Because of the short combustion residence time these metallized fuels have low combustion efficiencies. Therefore it is necessary to increase the combustion efficiency and the inlet air temperature does an important role to this. The main factors to affect the inlet air temperature is the free stream temperature and the flight Mach number. Also the flow velocity in the combustor does an important role, therefore entire range of the air flow; from the stagnation to the sonic velocity in the ramjet combustor is considered.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.3
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• pp.18-30
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• 2020

This study deals an estimation method of thrust measurement uncertainty in solid rocket motors. Guidelines of the force measurement uncertainty estimation have been provided by ISO, domestic and international organizations. However, all of them are described by focusing on the force calibration machines and force transducers with a conceptually-driven way. Thus the guidelines cannot be directly applicable to uncertainty estimation of calibration equation and its linear approximation, which are critical error sources in the thrust measurement. In this paper, the equations taking into account effects of both error sources are derived based on fundamental concepts of measurement uncertainty. These are applied to the real thrust measurement system where a relatively simple estimation method for the thrust measurement uncertainty is proposed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.349-351
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• 2011

Reliability estimation of solid rocket motor is more important to meet the advanced technology and to save the development fund. Domestic reliability requirement of rocket motor has been fulfilled by design capibility. But, reliability estimation process of our own shoud be estabilished by surveying the advanced nation's estimation technology.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.12-12
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• 1998

무기체계의 추진체로서 고체 로켓트 추진기관이 제작의 용이성, 구조의 간단성, 이에 따른 저렴한 제작비, 그리고 고 신뢰도 확보가능 등의 여러 장점으로 대부분의 현존 전술 유도무기에 채택되어 사용되어 오고 있으나 대응, 방어 무기체계의 빠른 발전으로 이에 따른 새로운 추진기관의 유도무기체제가 요구되고 있다. 램제트 기관은 공기흡입추진기관으로 상대적으로 높은 비추력(1000-2000s)과 추력 중량비(∼20)을 가지며, 이로 인해 기존의 로켓 엔진에 비해 4-5배의 성능을 낼 수 있으며, 초음속 장거리 비행에 적합하다며, 또한 높은 속도영역까지 운용가능하고 구조가 비교적 간단하다.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.60-66
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• 2006

Notable are the burning rate characteristics of solid propellant burning at extremely high pressure(10000-20000 psia). The burning rate test using closed bomb shows the discontinuous increment around 4000 psia so that the exponent of burning rate(n) is almost double, from 0.4 to 0.8. The pressure-increasing rate of the test motor is about 300 times as high as that of the motor operating at the conventional pressure, less than 2000 psia, is, therefor the burning rate is augmented about 5-50 times. The performance prediction reflecting the pressure-change-rate effect are fairly comparable with the test data at various test conditions.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.2
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• pp.28-34
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• 2020

In order to relieve the burden of the rear ablative material, the combustion test of the solid rocket motor with the forward deployed multi-pin grain design was successfully performed twice. However, after disassembling the solid rocket motor, a non-uniform ablation pattern was found in the rear ablative material. Periodic repetition of local and regional ablation was measured precisely. Two-dimensional flow and eddy flow, created by the uneven main-pin flow hitting the rear ablative material, were identified as the cause of non-uniform ablation. In addition. when the rear pins were removed, the possibility of securing the soundness of the rear ablative material was confirmed as the average flow velocity and the standard deviation were lowered.