• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.175-178
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• 2008

A Numerical Simulation which uses preconditioning algorithm to examine unsteady combustion processes for the AP/HTPB propellant with a converging-diverging nozzle has been compared with experimental data for solid rocket motor. To analysis reacting flow of solid rocket motor, unsteady pressure, temperature contour was simulated by grid moving of propellant.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.197-200
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• 2007

This paper presents solid propulsion manufacturing processes. Solid rocket motors propellants are made of several batches of propellant being mixed in mixers and made of continuous mixers. This paper discusses the technology of continuous mix process of Aerojet, US and SNPE Materiaux Energetiques, France.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.2
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• pp.1-8
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• 2005

The effect of coolants has been studied on the burning properties of low burning rate HTPB/AP composite propellant containing Oxamide or Melamine as coolant for the gas generator. With increasing the content of coolant, the burning rate and the flame temperature could be lowered and the effect on flame temperature was about the same for two coolants. However due to the different thermal decomposition properties of coolant, the burning rate of Melamine propellant was found to abnormally decrease if $200{\mu}m$ AP was partially replaced with $6{\mu}m$ AP.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1994.11a
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• pp.27-37
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• 1994

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.67-72
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• 2006

Several propellants were investigated experimentally for ignition characteristics in subatmospheric pressure. The threshold ignition pressure was 4 psia for HTPB/AP composite propellant. The partial replacement of AP in HTPB/AP propellant by $5{\sim}l5%$ of HMX, HNIW showed that the improvements in ignition delay was over 50% and the threshold pressure was below 0.4 psia. This appears to be due to the characteristics of HMX and HNIW exothermic dissociated at the temperature(${\sim}220^{\circ}C$) love. than that of AP. The ignition substance $B/KNO_3$ was coated thinly on the propellant surface for better ignition performance. As a result, ignition delay time of 15% was improved. NC is applied to $B/KNO_3$ ignition substance as a secondary binder and $NC-B/KNO_3$ suspension solution is coated to the propellant surface.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.6
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• pp.60-65
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• 2021

In this paper, we examined the ignition possibility of the propellant depending on its non-uniform composition of aluminum. Impact and friction sensitivity was investigated by arbitrarily changing the aluminum content in the range of 14~20% to simulate the non-uniform distribution of aluminum in the propellant. As a result of measuring the impact sensitivity, the 50% ignition energy and minimum ignition energy have values around 50 J regardless of the aluminum content. This means that the propellant does not become sensitive to impact even if the aluminum content is increased. On the other hand, the friction sensitivity result shows that as the aluminum content increases, the 50% ignition force and minimum ignition forces were decreased, and thus the propellant becomes sensitive. "Hot Spot" model of propellant ignition is applied, the space inside the propellant is momentarily compressed and ignited by friction stimuli rather than by impact stimuli.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.6
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• pp.42-53
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• 2015

In this study, the influence of Al particle size, as an additive for solid propellant, on the mechanical erosion of the carbon-composite nozzle was evaluated. A new model which can predict the size and distribution of the agglomerated reaction product($Al(l)/Al_2O_3(l)$) was established, and the size of agglomerate were calculated according to the various initial size of Al in the solid propellant. With predicted results of the model, subsequently, the characteristics of mechanical erosion on the carbon-composite nozzle was estimated using a commercial CFD software, STAR CCM+. The result shows that the smaller the initial Al particles are, in the solid propellant, the lower is the mechanical erosion rate of the composite nozzle wall, especially for the nano-size Al particle.

• Aerospace Engineering and Technology
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• v.4 no.2
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• pp.209-215
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• 2005

A qualification of propellant for a AKM(Apogee Kick Motor) has been conducted by firing standard motors. The Standard motor used in developing process of KSLV-I's AKM is ST-6 whose diameter is about six inches. HTPB composite propellant are used in these motors. Several standard motors are fired and compared with simulation results. A mean value of characteristic velocity is 1518 m/s. The characteristic value for prediction is 1516 m/s, and the thrust loss is about 5.9%. Differences between the measurements and prediction are less than 3%.

• Progress in Superconductivity and Cryogenics
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• v.22 no.4
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• pp.51-56
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• 2020

In order to increase thrust of the space launch vehicle, liquid oxygen as an oxidizer and kerosene or liquid hydrogen as a fuel are generally used. The oxidizer tank and fuel tanks are manufactured by composite materials such as CFRP (Carbon Fiber Reinforced Plastic) to increase pay load. The thermal stress of the cryogenic propellant tank should be considered because it has large temperature gradient. In this study, to confirm the design integrity of the oxidizer tank of liquid oxygen, a numerical analysis was conducted on the thermal stress and temperature distribution of the tank for various charging speed of the cryogenic fluid from 100 ~ 900 LPM taking into account the evaporation rate of the liquid nitrogen by convective heat transfer outside the tank and boiling heat transfer inside the tank. The thermal stress was also calculated coupled with the temperature distribution of the CFRP tank. Based on the analysis results, the charging speed of the LN2 can majorly affects the charging time and the resultant thermal stress.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.555-558
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• 2010

This paper presents a design of solid rocket motor with high length to diameter applied composite propellant. Solid rocket motor with high L/D ratio can be generated erosive burning and combustion instability on longitudinal mode. Especially, Erosive burning can effectively prolong the initial pressure spike in some star grain motors. That is, the study shows design of grain, internal ballistics and structural analysis in order to perform system requirements.