• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.723-725
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• 2010

Calculation program to predict required pressurant mass for propellant tank was verified through flight test data. This program was already developed and verified through ground test data, but to increase reliability of program, it was compared with flight test data of KSR-III launched in 2002. Because pressurant temperature incoming to propellant tank was not measured in flight test, that was assumed in calculation program. Required pressurant mass and inside temperature of oxygen tank dome was compared. Validation of calculation program was verified by showing required pressurant mass accuracy of 6%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.85-88
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• 2011

In the trend of world wide environment preservation, researchers tried to find new environment friendly propellant instead of highly toxic propellant, Hydrazine. Among the candidates, ionic liquid propellants have lower toxicity, higher density, and higher specific impulse than Hydrazine. These ionic liquid propellants have high combustion chamber temperature, so catalyst supports such as gamma alumina cannot withstand in that temperature. Therefore, a catalyst that showed stable characteristic in high temperature is needed. Barium dopped alumina can be changed to Hexaaluminate in high temperature, and its characteristic in high temperature is superior than gamma alumina. Barium dopped Alumina is wet impregnated with Platinum and heated up to $1300^{\circ}C$ and $1400^{\circ}C$ for 2 hours. Those catalysts were examined by XRD, SEM, EDS, BET, and Drop test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.413-416
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• 2005

The study has been performed on the improvement of bonding process and bond strength of HTPB propellant and liner using a polymeric curative. In case of liner using polymeric curative prepared from reaction of HTPB and TDI, migration of curative was decreased at bond interface. So EPDM insulation sanding and Desmodur RE coating process could be omitted in motor case preparation and bond strengths between the HTPB propellant and liner were increased. Also deterioration phenomena of bond strength could not be observed in accelerated aging test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.215-218
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• 2007

Internal Ballistic modeling and performance prediction for solid propellant micro thruster was performed with heat loss to the chamber wall as an important factor of miniaturization. Simple l-D end-burner type thruster and general HTPB-AP type composite propellant were selected for computation model. The results showed that the performance loss with the heat loss to the surroundings becomes larger as the surface-to-volume ratio is increased. In this case, the total impulse was reduced about 3% of the case in adiabatic condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.2
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• pp.110-114
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• 2006

The study has been performed on the improvement of bonding process and bond strength of HTPB propellant and liner using a polymeric curative. In case of liner using polymeric curative prepared from reaction of HTPB and TDI, migration of curative was decreased at bond interface. So EPDM insulation sanding and Desmodur RE coating process could be omitted in motor case preparation and bond strengths between the HTPB propellant and liner were increased. Also deterioration phenomena of bond strength could not be observed in accelerated aging test.

• Journal of the Korea Institute of Military Science and Technology
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• v.25 no.3
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• pp.321-328
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• 2022

The current development tend of the gun propellants that they should have low sensitivity and high energy. We studied a nitrocellulose based propellant composition that replaced sensitive NG with RDX and DEGDN which high energy and low sensitivity. The important factors in the design of the gun propellant were impetus and flame temperature. NC-based propellant containing RDX showed similar impetus but low flame temperature compared to KM30A1, a triple-based propellant. The developed propellant composition didn't show any abnormal combustion reaction and the characteristics of ballistic resistance were also confirmed.

• Journal of the Korean Society for Precision Engineering
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• v.15 no.3
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• pp.31-37
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• 1998

The structure of satellite consists of six parts which are control system, power system, thermal control system, remote measurement command system, propellant system and thrust system. In these parts, propellant system consists of propellant tank and thrust device. What we want to perform is optimum design to minimize the weight of propellant tank. In order to design optimal propellant tank, several parameters should be adopted from the tank geometry like the relative location of the lug and variation of the wall thickness. The analysis was executed by finite element analysis for finding optimal design parameters. The structure was divided into three parts consisting of the initial thickness zone, the transitional Bone, and the weak zone, whose effects on the pressure vessel strength was investigated. Finally the optimal lug location and the three zone thickness were obtained and the weight was compared with the uniform thickness vessel.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.39-45
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• 2013

Basically, in order to apply solid propellant as ignition source to high energy metal particle combustion system, we analyzed combustion characteristics of the HTPB/AP/Al, HTPE/AP/Al propellants by using a strand burner. The propellants were tested in a high-pressure windowed strand burner, which was pressurized up to 300 psia with pure argon gas. Strand burner was visualized with two quartz windows and ignition was accomplished by a 10 W $CO_2$ laser. The burning rate of propellant was measured with high-speed camera method for frame analysis and photodiode method for combustion time analysis. Emission spectrum was measured with spectrometer at 300 nm ~ 800 nm and 1500 nm ~ 5000 nm and then we analyzed species during propellant combustion.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.181-184
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• 2003

We manufactured standard motor to measure burning rate of propellant, used to estimate burning rate of main motor by static fired testing. We installed static fired testing facility and developed standard motor more lightly to accomplish the test. As a result of the tests, we could get the burning rate closer than acquired by existing method to the main motor's on. And we confirmed repeatability by many times of tests. We will use this method positively for R&D and quality assurance activity of mass production.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.41-47
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• 1999

The burning rate of solid propellant is one of the key parameter associated with the dynamic characteristics of combustion and the combustion performances. In the AP propellants, the evaporation on the reacting surface as well as the decomposition of the propellant is of great importance in determining the overall burning rate. In this study, a theoretical analysis was conducted to obtain the expression for burning rate in the steady state combustion with the energy and species equations in the condensed phase when the radiative heat flux partially contributes to the total heat transfer to the propellant surface.