• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.2
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• pp.1-8
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• 2015

Dual pulse rocket motor is a variant of solid rocket motor with two propellant grain separated by a pulse separation device. The major performance of such a rocket motor is influenced by the change in the hole area of pulse separation device to nozzle throat area ratio. In this study, we performed flow analysis to investigate the internal flow characteristics according to the pulse separation device hole area to nozzle throat area ratio change. Gases used flow analysis were used combustion gas of HTPB/AP composite propellant and nitrogen gas. Flow analysis results of the dual pulse rocket motor were validated by comparison with experimental results of pneumatics. Commercial CFD code ANSYS FLUENT 14.5 is used in this study to simulate flow analysis.

• Journal of Aerospace System Engineering
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• v.15 no.3
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• pp.79-86
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• 2021

For a preliminary study on a thrust-throttleable Variable Flow Ducted Rocket, a gas generator and flow control valve were developed, and ground combustion tests were performed. The gas generator and flow control valve operated at the required performance level for parameters, such as heat-resistance, combustion-time, pressure, and temperature. The combustion characteristics of a fuel-rich solid propellant mixed with Boron/MgAl/AP, etc., were also analyzed. A Plunger-type flow control valve was designed to control the discharge flow area, and it was confirmed that the flow control valve was able to control the combustion gas flow rate and pressure. However, due to the reduction of the discharge flow area caused by adhesion of combustion products, the combustion pressure continuously increased. The analysis of the pressure increase is covered in Part 2 of this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.6
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• pp.31-37
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• 2010

This study was carried out to investigate the thermal decomposition and execute EIDS slow cook-off test for the propellant ingredients and 2 kinds of HTPE propellants. The thermal analysis of the propellant ingredients used in this study showed that the thermal stability of these materials decreases in the following order : AP > HTPE > AN > BuNENA. In addition, propellant HTPE 002 containing AN showed that an endothermic process at around $125^{\circ}C$ corresponding to the solid phase change(II$\rightarrow$I) of AN was followed by the exothermic process of BuNENA/AN mixture up to $200^{\circ}C$. In EIDS slow cook-off tests, HTPE 001 and HTPE 002 reacted at around $250^{\circ}C$ and $152^{\circ}C$ respectively, and both of them showed sudden temperature increase curves at $115^{\circ}C$. The critical temperatures, $T_c$, of thermal explosion for the propellants HTPE 001 and HTPE 002, were obtained from both the non-isothermal curves at various heating rates and Semenov's thermal explosion theory. Kissinger's method that was used to calculate $T_c$ was also employed to obtain the activation energies for thermal decompositions.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.95-103
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• 1999

Low pressure combustion characteristics of the composite solid propellants were studied in terms of the propellant burning rate, ignition processes, and the structure of the extinguished surfaces. Optical Vacuum Strand Burner(OVSB) system was designed and configured for this purpose. Burning rates of the propellants were measured at subatmospheric pressure by developed test method in OVSB. The ignition and combustion phenomena of the studied propellants in the combustion chamber of OVSB were recorded and analyzed with the camera and VCR(30 frames/s). Burning surfaces of the propellants were extinguished by rapid depressurization method and analyzed with Scanning Electron Microscope(SEM).