• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.261-264
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• 2008

This study experimentally examines combustion characteristics of a hybrid rocket in which solid paraffin is used as a fuel, while oxidizer is pure oxygen. Especially, the experiment investigates the effects of chamber pressure and configuration of fuel grain. The pressure inside the combustion chamber is varied by changing a flow rate of oxidizer. The regression rate is observed to increase as the chamber pressure does. There also exists the effects of shape of fuel grain on thrust. Characteristic of paraffin hybrid rocket changes with shape of fuel grain. When there is a room near the injector, thrust increases. On the other hand, the room near the nozzle does not contribute to thrust increasement.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.228-231
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• 2011

Burning rate of a propellent is an essential factor when designing a propulsion system. In order to come up with burning rate, first we need to design and build propellent grain to get neutral pressure curve. Then check the pressure with ground test and calculate the burning rate using burning rate equation. This burning rate is then compared to the burning rate of a propellent which was resulted from making a standardized specimen and combusting it using a strand burner. An accurate burning rate is calculated after comparing those two burning rates. For this study, compact propulsion system was designed, produced, tested and analyzed in order to get burning rates, an essential factor in propulsion system design, in an effective way.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.365-368
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• 2004

The multi-phase flow analysis in a solid rocket motor is very important when performing the performance of a motor, and prediction of nozzle ablation. However, only in consideration of regular power, it has analyzed as power which a metal particle receives from a flow until now. We conduct analysis and an experiment about the virtual mass clause which will influence at the place where acceleration is big. We aim at the improvement in accuracy of multi-phase flow analysis from the result.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.69-72
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• 2011

The ground tests of the dual pulse solid rocket motors(SRM) were performed to prove the performance of the developed dual pulse solid rocket motors. To achieve dual pulse, pulse seperation device(PSD) and independent ignition systems were installed between the 1st pulse motor and the 2nd pulse motor. The three major development goals are as follow. First, the total impulse of the dual pulse SRM must be over 90% of the single SRM which has same envelopment and weight. Second, the ignition of the 2nd pulse motor can be started at any time between 0 and 60 seconds after termination of the 1st pulse motor. Last, the rupture disk of PSD must not be broken at the 1st pulse pressure, but be broken at below 30% of the 2nd pulse operating pressure. The results of the ground tests satisfied the development goals.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.18-25
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• 2006

In order to perform system requirements quickly and accurately, an automatic design program of solid rocket motors(SRM) structure designated as the 'ProDes software' has been developed and verified. from given system design criteria and constraints, it has the capabilities to design, analysis, simulation and drawing process to greatly reduce the over 'design cycle time' and manpower of a project. The conception of the program is modular, and calculations are performed step by step allowing parametric design studies and providing final selected design goal. Each configurations of SRM components and joint types composed of various master models is obtained from the data base module of the library. Between the design results of the ProDes software and those of the previous detail design of the established motor showed good agreements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.543-546
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• 2017

Test equipment was designed and manufactured to evaluate thermal reaction characteristic of internal insulators of solid rocket motor. Test is allowed up to chamber pressure 2,500 psi, burn-time 40 s. It is possible to observe and to compare thermal reaction characteristic for a few materials simultaneously, under the condition that the ablation occurs. In efficient average chamber pressure 878 psi, efficient burn-time 10.7 s and gas velocity 100 m/s, test was executed for confirming safety of equipment, being 4 test materials inserted simultaneously. Basic data for understanding thermal characteristic of internal insulator, that is, pressure-time curve, temperature-time curve in the test sample, and thermal destruction depth of test samples was gained successfully.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.9
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• pp.629-641
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• 2019

In order to limit the temperature rise of the structure to a certain level or less while maintaining the aerodynamic shape of solid rocket nozzle by effectively blocking a large amount of heat introduced by the combustion gas of high temperature and high pressure, the heat-resistant materials such as C/C composite having excellent ablation resistance are applied to a position in contact with the combustion gas, and the heat-insulating materials having a low thermal diffusivity are applied to the backside thereof. SiC/SiC composite, which has excellent oxidation resistance, is applied to gas turbine engines and contributes to increase engine performance due to light weight and heat-resistant improvement. Scramjet, flying at hypersonic speed, has been studying the development of C/SiC structures using the endothermic fuel as a coolant because the intake air temperature is very high. In this paper, characteristics, application examples, and development trends of various heat-resistant composites used in solid rocket nozzles, gas turbine engines, and ramjet/scramjet propulsions were discussed.

• International Journal of Aeronautical and Space Sciences
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• v.17 no.4
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• pp.565-578
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• 2016

A probabilistic based systems approach is addressed in this study for the reliability prediction of solid rocket motors (SRM). To achieve this goal, quantitative Failure Modes, Effects and Criticality Analysis (FMECA) approach is employed to determine the reliability of components, which are integrated into the Fault Tree Analysis (FTA) to obtain the system reliability. The quantitative FMECA is implemented by burden and capability approach when they are available. Otherwise, the semi-quantitative FMECA is taken using the failure rate handbook. Among the many failure modes in the SRM, four most important problems are chosen to illustrate the burden and capability approach, which are the rupture, fracture of the case, and leak due to the jointed bolt and O-ring seal failure. Four algorithms are employed to determine the failure probability of these problems, and compared with those by the Monte Carlo Simulation as well as the commercial code NESSUS for verification. Overall, the study offers a comprehensive treatment of the reliability practice for the SRM development, and may be useful across the wide range of propulsion systems in the aerospace community.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.3
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• pp.71-77
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• 2016

Characteristic of thermal reaction of Kevlar/EPDM internal insulator used in a solid rocket motor chamber was investigated at 1,030 psi, 35.86 s and 1,406 psi, 36.63 s. Surface status after test was similar each other and thermal destruction depth was 4.10 mm and 4.18 mm, respectively. Kinetic constant ${\xi}$, ${\zeta}$ and thermal destruction velocity $V_{TD}$ were also similar. It was concluded that characteristics of thermal reaction of Kevlar/EPDM internal insulator were not affected by change of chamber pressure.